AVIATION ELECTRICIAN'S MATE (AE)

NOTE Standard for Metric Practice, ASTM E 380-84 of 30 May 1984 was adopted for use by Department of Defense personnel. Refer to it whenever conflicts exist. MEASUREMENTS An AE must understand measurement because much of his/her professional responsibility concerns some type of measurement. The instruments an AE works with include relatively simple devices such as meters to measure voltage, current, and resistance; as well as relatively sophisticated aircraft-installed measuring devices, such as altimeters, airspeed indicators, and position indicators. In any study of physics, specific words and terms have specific meanings that must be mastered. If you do not know the exact meaning of a term, you won’t understand the principles involved in the use of that term. Once you understand the term, however, a discussion of principles illustrates and/or emphasizes a particular point. The first part of this chapter defines some physical terms and briefly discusses a few principles. To evaluate results, you must often ask the questions "how much, how far, how many, how often, or in what direction." As scientific investigations become more complex, measurements must become more accurate and new methods must be developed to measure new things. Measurements are classified into three broad categories; magnitude, direction, and time. These categories are broken down into several types, each with its own standard units. Measurements of direction and time are standardized and have comparatively few subdivisions. Magnitude, on the other hand, is an extremely complex category with many classes and subdivisions. The unit of measurement is just as important as the number that precedes it, and both are necessary to give an accurate description. There are two widely used sets of fundamental units of measurement; the metric and English units. The metric unit is normally used to express scientific observations. When using metrics, the basic unit for measuring distance is the meter; for measuring mass, the kilogram is used; and for measuring time, the second. This is called the meter-kilogram-second (mks) system. The English system uses the foot for distance, the pound for mass, and the second for time; thus, it is called the foot-pound-second (fps). See Table 1-1 for other frequently used units.

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Table 1-1 — Frequently Used Units of Measurement ENGLISH SYSTEM METRIC SYSTEM GENERAL Acre bar (metrology) ELECTRICAL Btu (British thermal unit) calorie ampere bushel gram coulomb dram hectare decibel foot hertz farad gallon hour henry hertz joule ohm horsepower liter volt hour meter watt inch metric ton (1,000 kilogram) knot micrometer LIGHT mil minute candela mile newton lumen minute quintal ounce second peck stere pint pound quart second slug ton (short 2,000 pounds, long 2,240 pounds)

yard Distance The AE should be familiar with both the metric and English systems of measurement, since personnel in the AE rating will use both systems in measuring distance and length. Metric Units of Length Metric units of length are based on the standard meter, first measured as one ten­ millionth part of the distance between the earth's equator and one of the poles. In 1960, 1-3

the International Bureau of Weights and Measures defined the meter as 1,650,763.73 wavelengths of the spectral emission of the krypton isotope, Kr86. Kilometers are used to measure large distances; 1 kilometer (km) = 1,000 meters (m). For smaller measurements, the meter and its subdivisions are used. One meter equals 100 centimeters (1m = 100 cm) and 1 centimeter equals 10 millimeters (1cm = 10 mm); therefore, a meter equals 1,000 millimeters (1 m = 1,000 mm). The micrometer is even smaller. It is the unit often used to state the wavelength of light, or to refer to the size of a particle of foreign matter that may pass through a certain screen or filter in the liquid cooling system of electronic equipment. The micrometer is one-thousandth of a millimeter or one-millionth of a meter. The nanometer is one thousandth of a micrometer and one millionth of a millimeter or one billionth of a meter. English Units of Length The common units of the English system of distance measurement are inches, feet, yards, and miles; where 1 foot equals 12 inches, 1 yard equals 3 feet, and 1 mile equals 1,760 yards. Whereas a standard mile is 5,270 feet, a nautical mile is equal to 6,076.115 feet. A mil is equal to 1/1000 inch. In 1866, the United States, by an act of Congress, defined the yard to be 3600/3937 part of a standard meter, or in decimal form, 0.9144 meter. Thus, other conversions between the systems may be found by proper multiplication or division. Some approximate conversions are listed in Table 1-2.

Table 1-2 — Conversion Factors for Units of Length

km

m

cm

mm

in.

ft

yd

mile 1 km = 1 1,000 100,000 1 X 10+6 39,370 3,280.83 1,093.61 0.621369 1m= 0.001 1 100 1,000 39.37 3.28083 1.09361 6.214 x 10-4 1 c m= 1 X 10-5 0.01 1 10 0.3937 0.032808 1.094 x 10-2 6.214 x 10-6 1 mm = 1 x 10-6 1 x 10-3 0.1 1 0.03937 3.28 x 10-3 1.094 x 10-3 6.214 x 10-7 1 in.= 2.54 X lO-5 2.54 x 10-2 2.54 25.4 1 0.08333 0.02777 1.58 X 10-5 1 ft = 3.048 x 10-4 0.3048 30.48 304.8 12 1 0.33333 1.89 x w-4 1 yd = 9.144 x 10-4 0.9144 91.44 914.4 36 3 1 5.68 x w-4 1 mile = 1.60934 1,609.34 160,934 1,609,344 63,360 5,280 1,760 1 NOTE: When a number is multiplied by a power of ten, the decimal point is moved the number of places represented by t he power. A negative power moves the decimal point to the left. A positive power moves it to the right. Thus, 84 x 10-2 = .84, and 84 x 10 2 = 8,400. Simply stated, a power of ten merely moves the decimal point left or right. 1-4

Figure 1-1 — Volume measurement. Figure 1-2 — Measuring the volume of an irregular object. Volume Volume is the amount of space enclosed within the bounding surfaces of a body. To determine the volume of a regularly shaped body, you will need three measurements; length, width, and height (depth). Volume = Length x Width x Depth (Height) V = LWH Volume measurement (Figure 1- 1) is expressed as dimensions of length cubed because it is the product of three length measurements. The unit of volume is a cube having edges of unit length. A great deal of ingenuity is often needed to measure the volume of irregularly shaped bodies. Sometimes it is practical to divide a body into a series of regularly shaped parts and then apply the rule that the total volume is equal to the sum of the volumes of all individual parts. Figure 1-2 shows another way to measure the volume of small irregular bodies. The volume of water displaced by a body submerged in water is equal to the volume of the body. Measuring the volume of floating bodies can be done in a similar way. A floating body displaces its own weight of liquid. This may be proved by filling a container to the brim with liquid, then lowering the body to the surface of the liquid. Next, the liquid that flows over the brim is collected. By weighing the displaced liquid and the original body, the statement, “A floating body displaces its own weight of liquid,” can be proven.

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NOTE Even though the pound is not mass unit (it is a force unit), it is often used as Mass, Weight, and Force The measure of the quantity of matter that a body contains is called mass. The mass of a body does not change. It may be compressed or expanded, but the quantity or mass of matter remains the same. For practical purposes, the metric unit of mass is based on the gram (g), which is equal to the mass of 1 cubic centimeter of pure water at a temperature of 4° Celsius. The standard pound (lb.) is the mass equal to 453.6 grams.

The mass of a body is constant no matter where the body is located. However, its weight (the force with which it is attracted toward the earth) is not constant. Its weight is slightly higher at the poles than at the equator, and it becomes less as the body moves away from the earth's surface. In addition to using grams (and pounds) as units of mass, these units are used to describe the weight of a body by comparing the body's weight to the weight of a standard mass unit. Unless otherwise specified, when an object is described as having a weight of 1 pound, it means the object has the same pull of gravity that a mass of 1 pound would have when located near sea level. To avoid confusion, the slug is used as the unit of mass. It weighs 32 pounds at sea level in the English system. In the metric system, the term newton, the force that causes 1 kilogram mass to be accelerated , is used. Conversion between the weight units of the metric system is simple. It is only a matter of moving the decimal point. For example:  1,000 milligrams (mg) = 1 g  1,000 g = 1 kg  1,000 kg = 1 metric ton Conversion between units of the English system requires more effort, since the pound is divided into 16 ounces and the ounce into 16 drams. The short ton is 2,000 pounds, while the long ton is 2,240 pounds. The metric ton is fairly close to the long ton, c onverting to 2,205 pounds. DERIVED UNITS Units based on combinations of two or three fundamental units are expressed as some combination of these units. The watt (unit of power) is written as a joule (unit of work) per second. The joule, in turn, could be expressed as newtons (force) times meters (distance), and the watt then becomes newton-meters per second. Likewise, the unit of horsepower is expressed in foot-pounds per second. Although there are conversion factors between derived units of the English system and the metric system, fundamental units of the two systems are not combined. For instance, if force is given in pounds and distance in meters, one or the other must be changed before combining them to get work units. 1-6

Speed and Velocity One example of a derived unit is the knot, a unit of speed. This unit combines the nautical mile as the unit of distance and the hour as the unit of time. The knot is derived by dividing the distance traveled by the time required. Thus, if a ship traveled at a constant rate for 15 minutes (0.25 hr) and moved a distance of 6 nautical miles, its speed would be 6 divided by 0.25 or 24 knots. The rate of travel (speed) is also used to solve for distance traveled when time is known. If the above ship were to travel for 3 hours at 24 knots, it would move 72 nautical miles. Likewise, the time required to move a certain distance may be determined when the speed is known. A movement of 36 nautical miles by a ship traveling at 24 knots would require 36/24 = 1.5 hours, or 1 hour 30 minutes. Often, speed is expressed with two fundamental units, such as miles per hour, kilometers per hour, or feet, inches, meters, or centimeters per minute or per second. Conversion is a matter of replacement of one unit by its equivalent in another unit. For example, a speed of 60 miles per hour (60 mph) is converted to feet per second by replacing the mile with 5,280 feet and the hour with 3,600 seconds. Thus, a speed of 60 mph = 60 (5,280 ft/3,600 s) = 88 feet per second. Table 1-3 gives the conversion factors between meters per second, feet per second, kilometers per hour, miles per hour, and knots.

Table 1-3 — Conversion Factors for Speed and Velocity Speed m/s ft/s km/hr mi/hr knots 1 m/s = 1 3.281 3.6 2.24 1.94 1 ft/s = 0.3048 1 1.0973 0.6818 0.5921 1km/hr = 0.27778 0.9113 1 0.6214 0.5396 1 mi/hr = 0.44704 1.4667 1.6093 1 0.8684 1 knot = 0.5148 1.689 1.853 1.152 1

The terms speed and velocity are sometimes used as if they had the same meaning. However, velocity is a vector quantity; that is, it is speed in a given direction. Thus, a car may move around a circular path with a constant speed while its velocity is continuously changing. When a body moves with constant speed along a straight line whose direction is specified, it is customary to speak of its velocity (which is numerically equal to its speed). When a body moves along a curved path or along a straight path with no reference being made to direction, it is proper to speak of its speed. Work and Energy Units of work and energy are derived units. They are the product of the units of force and distance. The joule is the unit of work in the mks system, where 1 newton acts through a distance of 1 meter. 1-7

In the English system, the unit foot-pound is defined as the work done in lifting 1 pound a distance of 1 foot against the force of gravity. Thus, the work done in lifting a mass of 5 pounds vertically 4 feet (5 lb x 4 ft = 20 foot-pounds) is 20 foot-pounds. Do not confuse the term foot-pound with the term used to measure torque. Since 1 pound force equals 4.448 newtons, and 1 foot equals 0.3048 meter, then 1 foot-pound is approximately 1.356 joules. The British thermal unit (Btu) is the heat energy required to raise the temperature of 1 pound of water 1 degree Fahrenheit. It is equivalent to 252 calories and, incidentally, to 777.8 foot-pounds of mechanical energy. Power All units of power include measurements of force, distance, and time because power equals work, which is force times distance divided by time. The watt is the unit of power frequently used with electrical units, and it is also the rate of doing 1 joule of work in 1 second. Thus, if a force of 5 newtons acts through a distance of 12 meters in 3 seconds, the power required is P = 5 x 12 = 20 watts 3 If the same work is to be done in 2 seconds, 30 watts will be required. Horsepower is a larger unit of power. It is equal to 550 foot-pounds per second, or 746 watts; therefore, 1 foot-pound per second is 746/550 watts, or about 1.356 watts. TEMPERATURE If an object is hot to the touch, it is said to have a high temperature; if it is cold to the touch, it is said to have a low temperature. In other words, temperature is used as a measure of the hotness or coldness of an object being described. Hotness and coldness are only relative. For example, on a cold day, metals seem colder to the touch than nonmetals because they conduct heat away from the body more rapidly. Also, upon leaving a warm room, the outside air seems cooler than the actual air temperature. Coming from the cold outside into a warm room, the room seems warmer than is the actual room temperature. In other words, the temperature a person feel s depends upon the state of their body. Temperature Conversion There are many systems of temperature measurement, and you will often need to convert from one to the other. The four most common scales (Figure 1-3) in use today are the Fahrenheit (F), Celsius (C), Kelvin (K), and Rankine (R) scales. FAHRENHEIT SCALE – The most familiar scale to most Americans is the Fahrenheit scale, which was established so that its zero point approximates the temperature produced by mixing equal quantities by weight of snow and common salt. Under standard atmospheric pressure, the boiling point of water is 212 degrees above zero, and the freezing point is 32 degrees above zero. Each degree represents an equal division, and there are 180 such divisions between freezing and boiling. CELSIUS SCALE – This scale uses the freezing point and boiling point of water under standard atmospheric pressure as fixed points of 0 and 100, respectively, with 100 equal divisions between them. 1-8

Figure 1-3 — Comparison of the four common temperature scales. The 100 divisions represent the same difference in temperature as the 180 divisions of the Fahrenheit scale. This creates a ratio of 100/180, which reduces to 5/9. This means that a change of 1°F is equal to a change of 5/9°C. A change of 5 degrees on the Celsius scale, therefore, is equal to a change of 9 degrees on the Fahrenheit scale. Because 0 degrees on the Celsius scale corresponds to 32 degrees on the Fahrenheit scale, a difference in reference points exists between the two scales. (See Figure 1-3.) To convert from the Fahrenheit scale to the Celsius scale, subtract the 32 degree difference and multiply the result by 5/9. Try converting 68°F to Celsius.

(68 - 32) =

x 36 = 20 °C To convert Celsius to Fahrenheit, reverse the procedure. First multiply the reading on the Celsius thermometer by 9/5 and then add 32 to the result.

(20) + 32 = 36 + 32 = 68 °F One way to remember when to use 9/5 and when to use 5/9 is to keep in mind that the Fahrenheit scale has smaller divisions than the Celsius scale. In going from Celsius to Fahrenheit, multiply by the ratio that is larger; in going from Fahrenheit to Celsius, use the smaller ratio. Another method of temperature conversion that uses these same ratios is based on the fact that the Fahrenheit and Celsius scales both register the same temperature at -40 degrees; that is, -40°F equals -40° C. This method of conversion, sometimes called the 40 rule, is calculated by adding 40 to the temperature to be converted, whether it is Fahrenheit or Celsius. Then, multiply the number by 9/5 when changing from Celsius to Fahrenheit or by 5/9 when changing from Fahrenheit to Celsius. Finally, subtract 40 from the previous number. This is the answer. For example, to convert 100 degrees to the equivalent Fahrenheit temperature using the 40 rule, complete the following steps: 1. 100 + 40 = 140 2. 140 x 9/5 = 252 1-9

3. 252- 40 = 212 Therefore, 100°C = 212 °F Remember that the multiplying ratio for converting Fahrenheit to Celsius is 5/9, and the multiplying ratio for converting Celsius to Fahrenheit is 9/5. Also, remember to ADD 40 first, multiply, and then SUBTRACT 40, regardless of the direction of the conversion. It is important for you to be able to read thermometers and convert from one scale to the other. In some types of electronic equipment, thermometers are provided as a check on operating temperatures. Thermometers are also used to check the temperature of a charging battery. KELVIN SCALE – Also known as the absolute scale, the Kelvin scale has as its zero point the temperature at which all molecular motion ceases and no additional heat can be extracted from the substance. Theoretically, this is referred to as absolute zero temperature. This point is -273.16°C, but -273°C is used for most calculations (Figure 1- 3). The spacing between points on the Kelvin scale is the same as the spacing between degrees on the Celsius scale; conversion from Celsius to Kelvin is made by adding 273 to the Celsius temperature. RANKINE SCALE – This scale has the same spacing between degrees as the Fahrenheit scale, but has its zero corresponding to 0.01 Kelvin (absolute zero). This is calculated to be -459.67°F; however, -460°F is usually used. To convert from the Fahrenheit scale to the Rankine scale, add 460 degrees to the Fahrenheit temperature. Since the Rankine and Kelvin scales both have the same zero point, conversion between the two scales requires no addition or subtraction. Rankine temperature is equal to 9/5 the Kelvin temperature, and Kelvin temperature is equal to 5/9 the Rankine temperature. Thermometers The measurement of temperature is known as thermometry. Many modern thermometers use liquids in sealed containers. Water was the first liquid used, but because it freezes at 0°C, it could not measure temperatures below that point. After much experimentation, scientists decided that the best liquids to use in the construction of thermometers are alcohol and mercury because of the low freezing points of these liquids. LIQUID THERMOMETERS – The construction of the common laboratory thermometer gives some of the meaning to a change of 1 degree in temperature. A bulb is blown at one end of a piece of glass tubing that has a small bore. The tube and bulb are filled with a liquid. The temperature of both the liquid and the tube during this process are kept at a point higher than the thermometer will reach in normal usage. The glass tube is then sealed and the thermometer cooled. During the cooling process, the liquid falls away from the top of the tube and creates a vacuum within the thermometer. Next, the thermometer is marked. The thermometer is placed in melting ice, and the height of the liquid column is marked as the 0° C point. Next, the thermometer is placed in steam at a pressure of 101,325 Pascals (Pa), and a mark is made at the point to which the liquid inside rises. That point is the boiling point or the 100°C mark. The space between these two marks is then divided into 100 equal parts. These spacing are known as degrees. This is the Celsius thermometer, which is used in laboratory work and in testing electrical equipment. SOLID THERMOMETERS – Because the range of liquid thermometers is limited, other methods of thermometry are necessary. Most liquids freeze at temperatures between 1-10

Figure 1-4 — Thermocouple. 0°C and -200°C. At the upper end of the temperature range where high heat levels are encountered, the use of liquid thermometers is limited by the high vapor pressures of those liquids. Among the most widely used types of thermometers, other than the sta ndard liquid thermometers, are the resistance thermometer and the thermocouple. The resistance thermometer makes use of the fact that the electrical resistance of m etals change as the temperature changes. Resistance thermometers are usually made of platinum wire wound on a mica form and enclosed in a thin­ walled, silver tube. It is extremely accurate from the lowest temperature to the melting point of the unit. The thermocouple (Figure 1-4) is an electric circuit. Its operation is based on the principle that when two unlike metals are joined and the junction is at a different temperature from the remainder of the circuit, an electromotive force is produced. This electromotive force can be measured with great accuracy by a galvanometer. Thermocouples can be located wherever measuring temperature is important and wires run to a galvanometer located at any convenient point. By using a rotary selector switch, one galvanometer can be used to read the temperatures of thermocouples at any number of widely separated points.

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Figure 1-5 — Compound bar. The principle used in compound bar thermometers (Figure 1-5) is the Coefficient of linear expansion which deals with Thermal Expansion in different types of metals (Temperature Differential). The bar may be in the shape of a spiral or a helix. Through the use of these shapes in a given enclosure, a greater length of the compound bar may be used; the movement of the free end per degree of temperature change is increased. Also, the indicating pointer may be joined to the moving end of the compound bar by a distance multiplying linkage to make the thermometer easier to read. Often this linkage is arranged to give the pointer a circular movement.

SOUND The range of sound that the human ear can detect varies with the individual. The normal range extends from about 20 to 20,000 vibrations per second. In the faintest audible speech sounds, the intensity at the ear is about watts/ at the threshold of feeling, the maximum intensity that the ear perceives as sound is about watts/ . If the ear is tested with tones of any one frequency, the threshold of audibility is, reached when intensity is reduced to such a low level that auditory sensation ceases. On the other hand, the threshold of feeling is reached when intensity is increased to such a high level that the sound produces the sensation of feeling and becomes painful. If this procedure is performed over a wide frequency range, the data can be used to plot two curves, one for the lower limit of audibility and the other for the upper limit (Figure 1- 6). Below the lower curve, the sound is too faint for you to hear. Above the upper curve, the sensation is one of feeling rather than of hearing; that is, the sensation of sound is masked by pain. The area between the two curves shows the pressure ranges for auditory response at various frequencies. 1-12

Figure 1-6 — Field of audibility. Sound Units The loudness of sound is not measured by the same type of scale used to measure length. Units of sound measurement vary logarithmically with the amplitude of the sound variations. These units are the bel and decibel, which refer to the difference between sounds of unequal intensity or sound levels. The decibel (one-tenth of a bel) is the minimum change of sound level that human ear perceives. Hence, the decibel merely describes the ratio of two sound levels. A sound for which the power is 10 times as great as that of another sound level differs in power level by 1 bel, or 10 decibels. For example, 5 decibels may represent almost any volume of sound, depending on the intensity of the reference level on which the ratio is based. In sound system engineering, decibels are used to express the ratio between electrical powers or between acoustical powers. Intensity Level An arbitrary zero reference level is used to describe the loudness of various sounds. This zero reference level is the sound produced by watts per square centimeter of surface area facing the source. This level approximates the least sound perceptible to the ear, and it is usually called the threshold of audibility. Therefore, the sensation experienced by the ear when subjected to a noise of 40 decibels above reference level would be 10,000 times as great as when subjected to a sound that is barely perceptible. STRUCTURE OF MATTER All matter is composed of atoms, which are composed of smaller subatomic particles. The subatomic particles of major interest in basic physics are the electron, the proton, and the neutron. They are electrical in nature, with the proton representing a positive charge, the electron representing a negative charge, and the neutron being neutral (neither positive nor negative). Although the composition of matter follows a consistent pattern for all atoms, the detailed arrangement of subatomic particles is different for each distinct substance. It is the combination and arrangement of the subatomic particles that gives a substance its distinguishing chemical and physical characteristics. 1-13

Figure 1-7 — An atom. NOTE Because of their smaller mass, electrons are more easily repositioned than protons. The protons and the neutrons of an atom are closely packed together in a nucleus (core), with the electrons revolving around the nucleus (Figure 1-7.) Normally, atoms are electrically neutral; that is, they contain an equal number of electrons and protons. However, atoms are not electrically neutral under all conditions. Atoms that contain an equal number of electrons and protons are known as balanced atoms; those with an excess (too many electrons) or a deficiency (too few electrons) of electrons are known as ions. The proton and the neutron have approximately the same mass (1,836 times that of an electron). In any atom, nearly all the mass is contained in the nucleus. Under normal conditions, if there were a change in the composition of the atom, there would be a change in the number or arrangement of the electrons.

ELEMENTS The word element identifies one of about 100 substances that comprise the basic substance of all matter. Two or more elements combine chemically to form a compound; and any combination that does not result in a chemical reaction between the different elements is known as a mixture. The atom is the smallest unit that exhibits the distinguishing characteristics of an element. An atom of one element differs from an atom of any other element in the number of protons in the nucleus. All atoms of a given element contain the same number of protons. Therefore, the number of protons in the nucleus determines the type of matter. Elements are tabulated according to the number of protons they contain. The number of protons in the nucleus of the atom is referred to as the atomic number of the element. Nucleus The study of the nucleus of the atom is known as nucleonic or nuclear physics. Experiments dealing with nuclei (plural of nucleus) usually involve the bombardment of the nucleus of an atom by various types of nuclear particles. By doing this, the composition of the nuclei is changed, and usually results in the release of energy. The change to the nuclei may occur as an increase or a decrease in the number of protons and/ or neutrons. 1-14

If the number of protons is changed, the atom has become an atom of a different element. Changing one element into another is known as transmutation (attempted by alchemists during the Middle Ages through chemical means, giving an impetus to the development of chemistry). If the number of neutrons in the nucleus of an atom is changed, the atom remains an atom of the same element. All atoms of a particular element have the same number of protons (atomic number). Atoms of certain elements may contain various numbers of neutrons; for example, hydrogen, which is the sole exception to the rule that all atoms are composed of three kinds of subatomic particles. Normally, a hydrogen atom contains a single proton, a single electron, and no neutrons. If the hydrogen atom contains a neutron, it is known as deuterium. Such an atom (although still a hydrogen atom) is also known as heavy hydrogen. (This atom is termed heavy because the addition of the neutron has nearly doubled the weight of the atom.) The atomic weight of an atom is an indication of the total number of protons and neutrons in the nucleus. Atoms of the same element having different atomic weights are known as isotopes. Nearly all elements have several isotopes; some are common, and some are rare. A few isotopes occur naturally; but most are produced by nuclear bombardment, and are radioactive or have unstable nuclei. These unstable isotopes undergo a spontaneous nuclear bombardment that eventually results in either a new element or a different isotope of the same element. The rate of spontaneous radioactive decay is measured by half-life, which is the time required for one-half the atoms of a sample of radioactive material to change (by spontaneous radioactive decay) into a different substance. For example, uranium, after a few billion years and several substance changes, becomes lead. Electron Shells The physical and chemical characteristics of an element are determined by the number and distribution of electrons in the atoms of that element. The electrons are arranged in successive groups of electron shells that rotate around the nucleus. Each shell can contain no more than a specific number of electrons. An inert element (a gas element that does not combine chemically with any other element) is a substance in which the outer electron shell of each atom is completely filled. In all other elements, there are one or more electrons missing from the outer shell. An atom with only one or two electrons in its outer shell can give up those electrons. Conversely, an atom whose outer shell needs only one or two electrons to be completely filled can accept electrons from another element that has one or two extras. The concept of needed or extra electrons arises from the basic fact that all atoms have a tendency toward completing (filling) the outer shell. For example, if an atom’s outer shell has only two electrons, it needs to collect six additional electrons (no easy task, from an energy standpoint) to have the eight required for that shell to be full. A much easier way to achieve the same objective is to give up the two electrons in the outer shell and let the full shell next to it serve as the new outer shell. In chemical terminology, this concept is termed a valence, which is the determining factor in predicting chemical combinations. COMPOUNDS AND MIXTURES Under certain conditions, two or more elements can be brought together so that they unite chemically to form a compound. The resulting substance may differ widely from any of its component elements. For example, ordinary drinking water is formed by the 1-15

chemical union of two gases; hydrogen and oxygen. When a compound is produced, two or more atoms of the combining elements join chemically to form the molecule that is typical of the new compound. The molecule is the smallest unit that exhibits the distinguishing characteristics of a compound. The combination of sodium and chlorine to form the chemical compound sodium chloride (common table salt) is a typical example of the formation of molecules. Sodium is a highly caustic, poisonous metal whose atom contains 11 electrons. Its outer shell has a single electron, which may be considered extra (a valence of +1). Chlorine, a highly poisonous gas whose atom contains 17 electrons, needs a single electron to fill its outer shell (has a valence of -1). When the atom of sodium gives up its extra electron, it becomes a positively charged ion. (It has lost a unit of negative charge.) The chlorine, having taken on this extra unit of negative charge (electron) to fill its outer shell, becomes a negative ion. Since opposite electric charges attract, the ions stick together to form a molecule of the compound sodium chloride. The attracting force that holds the ions together in the molecular form is known as the valence bond. You will see this term when learning about transistors. Remember, in the chemical combination, there is no change in the nucleus of either atom; the only change has occurred in the distribution of electrons between the outer shells of the atoms. Also, the total number of electrons has not changed, although they have been redistributed. Therefore, the molecule is electrically neutral, and has no resultant electrical charge. Not all chemical combinations of atoms are on a one-for-one basis. In the case of drinking water, two atoms of hydrogen (valence of +1) are required to combine with a single atom of oxygen (valence of -2) to form a single molecule of water. Some of the more complex chemical compounds consist of many elements having various numbers of each atom. All molecules, like all atoms, are normally considered electrically neutral. One exception to this rule is the chemical activity in batteries. Elements or compounds may be physically combined without undergoing any chemical change. Grains of finely powdered iron and sulfur stirred and shaken together retain their own identity as iron or sulfur. Salt dissolved in water is not a compound; it is merely salt dissolved in water. Each chemical substance retains its chemical identity, even though it may undergo a physical change. This is the typical characteristic of a mixture. STATES OF MATTER In their natural condition, forms of matter are classified and grouped in many different ways. One way to classify matter is according to its natural state; solid, liquid, or gas. This classification is important because of the common characteristics possessed by substances in one group distinguish them from substances in the other groups. However, you should remember that most substances can be made to assume any of the three forms. In all matter, molecules are in constant motion, and it is the extent of this motion that determines the state of that matter. The moving molecular particles in all matter possess kinetic energy of motion. The total of this kinetic energy is the equivalent of the quantity of heat in a sample of the substance. When heat is added, the energy level is increased, and molecular agitation (motion) is increased. When heat is removed, the energy level decreases, and molecular motion diminishes. In solids, the motion of the molecules is greatly restricted by the rigidity of the crystalline structure of the material. In liquids, the molecular motion is somewhat less restricted, 1-16

and the substance is permitted to flow. In gases, molecular motion is almost entirely random; the molecules are free to move in any direction and are almost constantly in collision both among themselves and with the surfaces of a container. Solids The outstanding characteristic of a solid is its tendency to retain its size and shape. Any change in these values requires the exchange of energy. The common properties of a solid are cohesion, adhesion, tensile strength, ductility, malleability, hardness, brittleness, and elasticity. Ductility is a measure of the ease with which the material can be drawn into a wire. Malleability is the ability of some materials to assume a new shape when pounded. Hardness and brittleness are self-explanatory terms. The remaining properties are discussed in the following paragraphs. COHESION AND ADHESION – Cohesion is the molecular attraction between like particles throughout a body, or the force that holds any substance or body together. Adhesion is the molecular attraction existing between surfaces of bodies in contact, or the force that causes unlike materials to stick together. Cohesion and adhesion are possessed by different materials to widely varying degrees. Generally, solid bodies are highly cohesive but only slightly adhesive. Conversely, fluids (liquids and gases) are usually highly adhesive but only slightly cohesive. Generally, a material having one of these properties to a high degree possesses the other property to a relatively low degree. TENSILE STRENGTH – The cohesion between the molecules of a solid explains the property known as tensile strength. Tensile strength is a measure of the resistance of a solid to being pulled apart. Steel possesses this property to a high degree; therefore, it is very useful in structural work. When a break does occur, the pieces of the solid cannot be stuck back together because merely pressing them together does not bring the molecules into close enough contact to restore the molecular force of cohesion. However, melting the edges of the break (welding) allow the molecules on both sides of the break to flow together, bringing them into the close contact required for cohesion. ELASTICITY – If a substance will spring back to its original form after being deformed, it has the property of elasticity. This property is useful in materials used as springs. Steel and bronze are examples of materials that exhibit this property. Elasticity of compression is exhibited to some degree by all solids, liquids, and gases. The closeness of the molecules in solids and liquids makes them hard to compress, but gases are easily compressed because the molecules are farther apart. Liquids Liquids tend to retain their own volume while assuming the shape of their container. Therefore, liquids are considered almost completely flexible and highly fluid. Liquids are practically incompressible; applied pressure is transmitted through them instantaneously, equally, and undiminished to all points on the enclosing surfaces. Hydraulic apparatus can increase or decrease input forces, and provides an action similar to mechanical advantage in mechanical systems. Because of the properties of liquids, hydraulic servomechanisms have advantages and limitations when compared with other systems. ADVANTAGES – The fluidity of hydraulic liquids permits the component parts of a system to be placed at widely separated points when necessary. Hydraulic power units transmit energy around corners and bends without using complicated gears and levers. 1-17

They operate with a minimum of slack and friction, which are often excessive in mechanical linkages. Uniform action is obtained without vibration, and the operation of the system remains largely unaffected by variations in the load. The accumulator (a component that provides the necessary pressurization of the system to furnish practically instantaneous response) can be pressurized during periods of non-action, eliminating the buildup time characteristic of electric servos. DISADVANTAGES – Hydraulic hoses used to transmit fluid from unit to unit are bulky and heavy when compared to electric wiring. Many of the hydraulic fluids are messy and are safety hazards, such as contributing to the danger of slipping. Also, hydraulic fluids cause deterioration of electric wiring insulation and they conduct electricity, increasing the hazards of short circuits. Some hydraulic fluids are flammable. Gases The most notable characteristics of a gas are its tendency to assume the shape and volume of its containers and the definite relationship that exists between the volume, pressure, and temperature of a confined gas. The ability of a gas to assume the shape and volume of its container is the result of its extremely active molecular particles, which are free to move in any direction. There is little cohesion between the molecules so they tend to separate and distribute themselves uniformly throughout the volume of the container. In an unpressurized container of liquid, pressure is exerted on the bottom and the sides of the container up to the level of the liquid. In a like container of gas, however, the pressure is also exerted against the top surface, and the pressure is equal at all points on the enclosing surfaces. The relationship of volume, pressure, and temperature of confined gas are explained by Boyle's law, Charles' law, and the general gas law. Many laboratory experiments based on these laws make use of the ideas of standard pressure and standard temperature. These are not natural standards, but are standard values selected for convenience in laboratory usage. Standard values are used at the beginning of an experiment, or when a temperature or a pressure is to be held constant. Standard temperature is 0°C, the temperature at which pure ice melts. Standard pressure is a pressure of 101,325 Pa. In many practical uses, these standards must be changed to other systems of measurement. All calculations based on the laws of gases make use of absolute temperature and pressure. These topics require a somewhat more detailed explanation. GAS PRESSURE – Gas pressure is indicated in either of two ways absolute pressure or gauge pressure. Since the pressure of an absolute vacuum is zero, any pressure measured with respect to this reference is referred to as absolute pressure. In this section, absolute pressure represents the actual pressure exerted by the confined gas. At sea level, the average atmospheric pressure is about 14.7 pounds per square inch (PSI). In a mercury barometer, this pressure would support a column of mercury 760 millimeters in height and be equal to 101,325 Pa. However, the actual pressure at sea level varies considerably and the pressure at any given altitude may differ from that at sea level. Therefore, the actual atmospheric pressure must be considered when converting absolute pressure to gauge pressure (or vice versa). When a pressure is expressed as the difference between its absolute value and that of the local atmospheric pressure, the measurement is termed gauge pressure and is usually expressed in pounds-per-square-inch gauge (PSIG). You can convert gauge pressure to absolute pressure by adding the local atmospheric pressure to the gauge pressure. 1-18

NOTE The capital P and T in the figure below represent absolute pressure and temperature. ABSOLUTE ZERO – Absolute zero, one of the fundamental constants of physics, is usually expressed in terms of the Celsius scale. It is used in the study of the kinetic theory of gases. According to the kinetic theory of gases, if the heat energy of a given gas sample could be progressively reduced, some temperature should be reached at which the motion of the molecules would cease entirely. If accurately determined, this temperature could then be taken as a natural reference, or a true absolute zero value. Experiments with hydrogen indicate that if a gas were cooled to -273.15°C (-273°C is used for most calculations), all molecular motion would cease and no additional heat could be extracted from the substance. In theory, at this point both the volume and the pressure of gas would shrink to zero. When temperatures are measured with respect to the absolute zero reference, they are expressed in the absolute or Kelvin scale; absolute zero may be expressed either as 0 K or as -273°C. BOYLE’S LAW – The English scientist, Robert Boyle, was among the first to study what he called the springiness of air. By direct measurement, he discovered that when the temperature of an enclosed sample of gas was kept constant and the pressure doubled, the volume was reduced to half the former value; as the applied pressure was decreased, the resulting volume increased. From these observations, he concluded that for a constant temperature, the product of the volume and pressure of an enclosed gas remains constant. CHARLES’ LAW – The French scientist, Jacques Charles, provided the foundation for the modern kinetic theory of gases. He found that all gases expand and contract in direct proportion to the change in the absolute temperature, provided the pressure is held constant. Since any change in the temperature of a gas causes a corresponding change in volume, it is reasonable to expect that if a given sample of a gas were heated while confined within a given volume, the pressure should increase. By actual experiment, it was found that the increase in pressure was approximately 1/273 of 0°C pressure for each 1°C increase. Because of this fact, it is normal practice to state this relationship in terms of absolute temperature. GENERAL GAS LAW – Figure 1-8 below shows Boyle's law in view A, while the effects of temperature changes on pressure and volume (Charles' law) are shown in views B and C.

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Figure 1-8 — Boyle’s, Charles’, and general gas laws.

By combining Boyle's and Charles’ laws, a single expression can be derived that states all the information contained in both. This expression is called the general gas equation (Figure 1-8, view D). Y ou can see in Figure 1-8, views A, B, and C, the three equations are applications of the general equation. Thus, if the temperature remains constant, equals , and both can be eliminated from the general formula, which then reduces to the form shown in view A. When the volume remains constant, equals thereby reducing the general equation to the form shown in view B. Similarly, is equal to for constant pressure, and the equation is that shown in view C. The general gas law (Figure 1-8, view D) applies only when one of the three measurements remains constant. When a gas is compressed, the work of compression is done to the gas. Work energy is converted to heat energy in the gas, so dynamic heating takes place. For example, when air at 0°C is compressed in a non-conducting cylinder to half its original volume, its temperature will rise to 90°C. When compressed to one tenth its original volume, its temperature rises to 429°C. The general gas law (Figure 1-8, view D) applies with exactness only to ideal gases in which the molecules are assumed to be perfectly elastic. However, it describes the behavior of actual gases with enough accuracy for most practical purposes. MATTER AND ENERGY Matter is defined as anything that occupies space and has weight or mass. It exists naturally in three states; solid, liquid, or gas. Matter may be changed or combined by physical, chemical, or nuclear means. Matter has many properties; properties possessed by all forms of matter are called general properties, while those properties possessed only by certain classes of matter are referred to as special properties. Energy is defined as the capacity for doing work, and it is classified in many ways. In this discussion, however, energy is classified as mechanical, chemical, radiant, heat, light, sound, electrical, or magnetic. Energy is constantly being exchanged from one object to another and from one form to another. 1-20

Law of Conservation Matter may be converted from one form to another with no change in the total amount of matter. Energy may also be changed in form with no resultant change in the total quantity of energy. In addition, a third statement has been added within the past half century, "Although the total amount of matter and energy remains constant, matter can be converted into energy or energy into matter.'' This statement is known as the law of conservation for energy and matter. The basic mathematical equation that shows the relationship between matter and energy is: E = Where;  E represents the amount of energy  m represents the amount of matter (mass)  c represents the velocity of light

This equation mathematically states that ''the destruction of matter creates energy, and that the creation of matter requires an expenditure of energy.'' From this observation, it may be inferred that a given quantity of matter is the equivalent of some amount of energy. In common usage, it is usually stated that ''matter possesses energy.'' General Properties of Matter All forms of matter possess certain properties. In the basic definition of matter, it is stated that "matter occupies space and has mass." Those two ideas contain most, if not all, of the general properties of matter. Space – The amount of space occupied by, or enclosed within, the bounding surfaces of a body is termed volume. In the study of physics, this concept is modified to be completely accurate. As stated previously, matter may appear as a solid, as a liquid, or as a gas, each having special properties. For a specific substance, the volume may vary with changes in circumstances, and liquids and solids tend to retain their volume when physically moved from one container to another, and gases tend to assume the volume of the container. To clarify the concept of occupying space, minute particles of matter must be dealt with. These are composed of still smaller particles separated from each other by empty space (which contains no matter). This idea is used to explain two general properties of matter, impenetrability and porosity. Impenetrability of Matter – Two objects cannot occupy the same space at the same time. This statement defines the concept known as the impenetrability of matter. The actual space occupied by the individual subatomic particles cannot be occupied by any other matter. The impenetrability of matter may, at first glance, seem invalid when a cup of salt is poured into a cup of water. The result is considerably less than two cups of salt water. However, matter has an additional general property called porosity. Porosity – Porosity explains this apparent loss of volume; the water simply occupies space between particles of salt. Porosity is present in all material, but some material is more porous than other material. Generally, gases are extremely porous, liquids only slightly porous, and the porosity of solids varies from that of the sponge to the steel ball. Inertia – Every object tends to maintain a uniform state of motion. A body at rest never starts to move by itself; a body in motion will maintain its speed and direction unless it is 1-21

caused to change. To cause a body to deviate from its condition of uniform motion, a push or pull must be exerted on it. This requirement is due to the general property of all matter known as inertia. The greater the tendency of a body to maintain uniform motion, the greater its inertia. The quantitative measure of inertia is the mass of the body. Acceleration – Any change in the state of motion of a body is known as acceleration. In other words, acceleration is the rate of change in the motion of a body and acceleration may represent either an increase or a decrease in speed and/or change in direction of motion. The amount of acceleration is stated as, ''the change of velocity divided by the time required to make the change." For example, if a car traveling 15 mph increased its speed to 45 mph in 4 seconds, the 30 mph increase divided by 4 seconds gives 7.5 miles per hour per second as its acceleration. By converting the 30 mph speed to 44 feet per second, the acceleration could be expressed as 11 feet per second per second, or as 11 . Force – Force is the action or effect on a body that tends to change the state of motion of the body acted upon. A force tends to move a body at rest. It tends to increase or decrease the speed of a moving body, or it tends to change the body's direction of motion. The application of a force to a body does not necessarily result in a change in the state of motion; it may only tend to cause such a change. A force is any push or pull, which acts on a body. Water in a can exerts a force on the sides and bottom of the can. A tug exerts a push or a pull (force) on a barge. A person leaning against a bulkhead exerts a force on the bulkhead. In the above examples, a physical object is exerting the force and is in direct contact with the body upon which the force is being exerted. Forces of this type are called contact forces. There are other forces, which act through empty space without contact; in some cases without even seeming to have any mass associated with them. The force of gravity exerted on a body by the earth (known as the weight of the body) is an example of a force that acts on a body through empty space and without contact. Such a force is known as an action-at-a-distance force. Electric and magnetic forces are other examples of these action-at-a-distance forces. The space through which these action- at-a-distance forces are effective is called a force field. Force is a vector quantity; that is, it has both direction and magnitude. A force is completely described when its magnitude, direction, and point of application are given. In a force vector diagram, the starting point of the line represents the point of application of the force. Any given body, at any given time, is subjected to many forces. In many cases, all these forces may be combined into a single resultant force, which is then used to determine the total effect on the body. Each body of matter in the universe attracts every other body with a force that is directly proportional to the mass of the bodies and inversely proportional to the square of the distance between them. This force is called the universal force of gravitational attraction. When considering the forces acting on a single body and since everybody exerts this force on every other body, it is an almost universal practice to resolve all gravitational forces into a single resultant. At or near the surface of the earth, this becomes a fairly simple process. Because of its extremely large mass, the earth exerts so large a gravitational attraction that it is entirely practical to ignore all other such attractions and merely use the earth's gravitational attraction as the resultant. 1-22

NOTE In the metric system, the specific gravity of a substance has the same numerical value as its density. Although gravitational attraction is exerted by each body on the other, it’s often more convenient to consider the force as being exerted by the larger mass on the smaller mass. Usually, this happens when there is a great difference in the mass of two bodies. Therefore, it is commonly stated that the earth exerts a gravitational force of attraction on a body. The gravitational attraction exerted by the earth on a body is known as gravity. The gravitational force exerted by the earth on a body is called the weight of that body, and is expressed in force units. In the English system, force is expressed in pounds. If a body is attracted by a gravitational force of 160 pounds, the body is said to weigh 160 pounds. The gravitational force between two bodies decreases as the distance between them increases; therefore, a body weighs less a mile above the surface of the ocean than it weighs at sea level; it weighs more a mile below sea level. Density and Specific Gravity The density of a substance is its weight per unit volume. A cubic foot of water weighs 62.4 pounds; the density of water is 62.4 pounds per cubic foot. (In the metric system the density of water is 1 gram per cubic centimeter.) The specific gravity (sp gr) of a substance is the ratio of the density of the substance to the density of water, or: sp gr = Weight of the substance Weight of equal volume of water

Specific gravity is not expressed in units but as a pure number. For example, if a substance has a specific gravity of 4, 1 cubic foot of the substance weighs 4 times as much as a cubic foot of water: 62.4 x 4 = 249.6 pounds. In metric units, 1 cubic centimeter of a substance with a specific gravity of 4 weighs 1 times 4 or 4 grams.

Specific gravity and density are independent of the size of the sample. They depend only upon the substance of which the sample is made. Look at Table 1-4 below. It contains some typical values of the specific gravity for various substances.

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Table 1-4 — Typical Values of Specific Gravity Substance Specific Gravity Aluminum 2.7 Brass 8.6 Copper 8.9 Gold 19.3 Ice 0.92 Iron 7.8 Lead 11.3 Platinum 21.3 Silver 10.5 Steel 7.8 Mercury 13.6 Ethyl alcohol 0.81 Water 1.00

Pressure and Total Force Pressure and force, while closely related, are not identical. A weight of 10 pounds resting on a table exerts a force of 10 pounds. However, the shape of the weight determines the effect of the weight. If the weight consists of a thin sheet of steel resting on a flat surface, the effect is quite different than if the same sheet of steel were resting on a sharp corner. Pressure is concerned with the distribution of a force with respect to the area over which that force is distributed. Pressure is defined as the force per unit of area, or P= F/A. A flat pan of water with a bottom area of 24 square inches and a total weight of 72 pounds exerts a total force of 72 pounds, or a pressure of 72/24 or 3 pounds per square inch on the flat table. If the pan is balanced on a block with a surface area of 1 square inch, the pressure is 72/1 or 72 pounds per square inch. An aluminum pan with a thin bottom is suitable for use on a flat surface, but may be damaged if placed on the small block. The concept of pressure and force explains why a sharp knife cuts more easily than a dull one. The smaller area concentrates the applied force (increases the pressure) and penetrates more easily. For hydraulic applications, the relationship between pressure and force is a basic operating principle. In enclosed liquids under pressure, the pressure is equal at every point on the surfaces of the enclosing container; therefore, the force on a given surface is dependent on the area.

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Figure 1-9 — Center of gravity in various bodies. Kinetic Energy Moving bodies possess energy because they are capable of doing work. The energy of mass in motion is called kinetic energy, which is expressed by the equation: Kinetic Energy =

Where m represents the mass of the body, and v is the velocity of its motion. When the moving body is stopped, it loses its kinetic energy. The energy is not destroyed, but is merely converted into other forms of energy such as heat and potential energy. Remember, bodies at rest also possess energy by virtue of their position. MECHANICS Mechanics is the branch of physics that deals with force, mass, and motion. Normally considered the fundamental branch of physics, it deals with matter. Many of its principles and ideas may be seen, measured, and tested. Since all other branches of physics are also concerned (to some extent at least) with force, mass, and motion, understanding this section will help you understand other branches of physics. For fu rther study on mechanics, you should refer to Basic Machines, NAVEDTRA 10624- A1. Force, Mass, and Motion Each particle in a body is acted upon by gravitational force. In every body, there is one point at which a single force, equal to the gravitational force and directed upward, would sustain the body in a condition of rest. This point is known as the center of gravity (cg), and it represents the point at which the entire mass of the body appears to be concentrated (Figure 1-9). The gravitational effect is measured from the center of gravity. In symmetrical objects of uniform mass, this is the geometrical center. In the case of the earth, the center of gravity is near the center of the earth. When discussing the motion of a body, it usually describes the path followed by the center of gravity. The natural tendency of a moving body is to move in a manner so that the center of gravity travels in a straight line. Movement of this type is called linear motion. Some moving bodies, however, do not move in a straight line, but describe an arc or a circular path. Circular motion falls into two general classes; rotation and revolution. Objects come in many different shapes, and to discuss rotary and revolutionary motion, it is necessary to consider the location of the center o f gravity with respect to the body. While reading this discussion, refer to Figu re 1- 9. 1-25

Figure 1-9 — Center of gravity in various bodies. Figure 1-10 — Center of gravity and center of rotation. In view A, the center of gravity of a ball coincides with the physical center of the ball. However, in the flat washer (view B), the center of gravity does not coincide with any part of the object, but is located at the center of the hollow space inside the ring. In irregularly shaped bodies (view C); the center of gravity may be difficult to locate exactly. If the body is completely free to rotate, the center of rotation coincides with the center of gravity. But, a body may be restricted so its rotation is about some point other than the center of gravity. Here, the center of gravity revolves around the center of rotation. These conditions are illustrated in Figure 1-10. Generally, the gyro rotor (view A) is said to rotate about its axis, and the ball (view B) is said to revolve about a point at the center of its path. Masses in Motion Motion is defined as the act or process of changing place or position. The state of motion refers to the amount and the type of motion possessed by a body at some definite instant (or during some interval) of time. A body at rest is not changing in place or position; it has zero motion, or is motionless. The natural tendency of any body at rest is to remain at rest; a moving body tends to continue moving in a straight line with no change in speed or direction. A body that obeys this natural tendency is in uniform motion. Any change in the speed or direction of motion of a body is known as acceleration and requires the application of some force. The acceleration of a body is directly proportional to the force causing that acceleration; acceleration also depends upon the mass of the body. For example, the greater mass of a lead ball makes it harder to move than a wood ball of the same diameter. The wood ball moves farther with the same push. These observations point to a connection between force, mass, and acceleration. They indicate that the acceleration of a body is directly proportional to the force exerted on that body and inversely proportional to the mass of that body. In mathematical form, this relationship may be expressed as: a =

Or, as it is more commonly stated, "Force is equal to mass times acceleration (F = ma).'' 1-26

Acceleration Due to Gravity The small letter g used in formulas for solving weight when mass is known (W = mg) represents the acceleration of a body in free fall, neglecting any friction. This can happen only in a vacuum. At sea level near the equator, g has the approximate values of 32 ft/s2 in the fps system, and 9.8 m/s2 in the mks system. Transposing the formula W = mg to solve for m, the absolute units of mass of a body may be determined when its weight is known. If the accelerating force is applied to the center of gravity in such a manner as to accelerate the body with no rotation, it is called a translational force. A force applied in such a manner as to cause the body to rotate about a point is called a torque force. Laws of Motion Among the most important discoveries in theoretical physics are Newton’s three fundamental laws of motion. These laws have been used in explanations of various topics earlier in this chapter. At this point, they clarify and summarize much of the discussion about mechanical physics.  Every body tends to maintain a state of uniform motion unless a force is applied to change the speed or direction of motion.  The acceleration of a body is directly proportional to the magnitude of the applied force and inversely proportional to the mass of the body; acceleration is in the direction of the applied force.  For every force applied to a body, the body exerts an equal force in the opposite direction. Momentum Every moving body tends to maintain uniform motion. Quantitative measurement of this tendency is proportional to the mass and velocity of the body (momentum = mass x velocity). The concept of momentum explains why heavy objects in motion at a given speed are harder to stop than lighter objects, and also why it’s easier to stop a given body moving at low speed than it is to stop the same body moving at high speed. Work, Power, and Energy You learned earlier that energy is the capacity for doing work. In mechanical physics, work involves the idea of a mass in motion; it is usually thought of as the product of the applied force and the distance through which the mass is moved (work = force x distance). For example, if a man raises a weight of 100 pounds to a height of 10 feet, he accomplishes 1,000 foot-pounds of work. The amount of work accomplished is the same regardless of the time involved. However, the rate of doing the work varies greatly. The rate of doing work, known as power , is defined as the work accomplished per unit of time (power = work/time). Refer to the previous example about the man raising a weight of 100 pounds. If the work is accomplished in 10 seconds, power is being expended at the rate of 100 foot-pounds per second; if it takes 5 minutes (300 seconds), power is being expended at the rate of approximately 3.3 foot-pounds per second. Therefore, measurements of power include force, distance, and time. In the English system of measurements, the unit of mechanical power is known as horsepower, and it is the equivalent of 33,000 foot-pounds per minute or 550 foot- 1-27

pounds per second. Energy is readily convertible from one form to another; therefore, the work and power measurements based on the conversion of energy must be readily convertible. For example, the electrical unit of power is the watt. Electrical energy can be converted into mechanical energy; therefore, electrical power must be converted into mechanical power. One horsepower is the mechanical equivalent of 746 watts of electrical power, and 1 horsepower is capable of doing the same amount of work as 746 watts in the same time. The accomplishment of work always involves a change in the type of energy, but does not change the total quantity of energy. Thus, energy applied to an object may produce work, changing the composition of the energy possessed by the object. Potential Energy A body has potential energy if, by virtue of its position or its state, it can do work. A wound clock spring and a cylinder of compressed gas both possess potential energy since they can do work in returning to their uncompressed condition. Also, a weight raised above the earth has potential energy since it can do work in returning to the ground. Thus, potential energy results when work has been done against a restoring force. The water in a reservoir above a hydroelectric plant has potential energy regardless of whether the water was placed there by work applied via a pump or by the work done by the sun to lift it from the sea and place it in the reservoir in the form of rain. Kinetic Energy The ability of a body to do work by virtue of its motion is called its kinetic energy. A rotating wheel on a machine has kinetic energy of rotation. A car moving along the highway has kinetic energy of translation. For a given mass (m) moving in a straight line with a velocity (v), the kinetic energy is determined by: kinetic energy =

in foot-pounds m is the slugs and v is in feet per second.

For example, the kinetic energy of a 3,200 pound car that is traveling at 30 miles per hour can be found by expressing the 3,200 pounds as 100 slugs and the 30 mph as 44 feet per second. Inserting these values into the formula makes; kinetic energy = x 100 x 44 x 44 = 96,800 foot-pounds of energy. This amount of kinetic energy is the result of 96,800 foot-pounds of work (plus that to overcome friction) having been applied to the car to get it traveling at the rate of 44 feet per second. The same amount of energy could do the work of lifting the 3,200 pounds vertically to a distance of 30.25 feet. Also, it would have the potential energy to move the car if it were at rest on an incline and then allowed to coast to a point that is vertically 30.25 feet below its starting point (again neglecting friction). Efficiency Provided there is no change in the quantity of matter, energy is convertible with no gain or loss. However, the energy that results from a given action may not be in the desired form; it may not even be usable in its resultant form. In all branches of physics, this concept is known as efficiency. 1-28

The energy expended is always greater than the energy recovered. An automobile in motion possesses a quantity of kinetic energy that depends on its mass and velocity. To stop the car, energy must be converted into potential energy. When the car comes to rest, its potential energy is considerably less than the kinetic energy it possessed while in motion. The difference, or the energy lost, is converted into heat by the brakes. The heat serves no useful purpose, so the recovered energy is less than the expended energy. The system is less than 100 percent efficient in converting kinetic to potential energy. The term efficiency, normally used in connection with work and power, means the ratio of the input to the output work, power, or energy. It is always expressed as a decimal or as a percentage less than unity. Friction In mechanical physics, the most common cause for the loss of efficiency is friction. Whenever one object slides or rolls over another, irregularities in the contacting surfaces interlock and cause an opposition to the force being exerted. Even rubbing two smooth pieces of ice together produces friction. Friction also exists in the contact of air with all exposed parts of an aircraft in flight. When a nail is struck with a hammer, the energy of the hammer is transferred to the nail, and the nail is driven into a board. The depth of penetration depends on the momentum of the hammer, the size and shape of the nail, and the hardness of the wood. The larger or duller the nail and the harder the wood, the greater the friction; therefore, there is lower efficiency and less depth of penetration, but greater heating of the nail. Friction is always present in moving machinery, which means that the useful work accomplished by the machine is never as great as the energy applied. Work accomplished to overcome friction is usually not recoverable. Friction can be minimized by decreasing the number of contacting points, by making the contacting areas as small and as smooth as possible, by the use of bearings, or by the use of lubricants. There are two kinds of friction; sliding and rolling. Rolling friction is usually of lower magnitude, so most machines are built so rolling friction is present rather than sliding friction. The ball bearing and the roller bearing are used to convert sliding friction to rolling friction. A third type of bearing, the common (or friction) bearing requires application of a lubricant to surfaces that have been made as smooth as possible. Many new types of machines use self-lubricating bearings to minimize friction and maximize efficiency. Mechanical Advantage The concept of mechanical advantage has proved to be one of the great discoveries of science. Mechanical advantage is the basic principle involved in levers, block and tackle systems, screws, hydraulic mechanisms, and other work saving devices. However, in the true sense, these devices do not save work; they merely enable humans to accomplish tasks that they could not otherwise do. For example, a normal human could not lift the rear end of a truck to change a tire; but with a jack, a block and tackle, or a lever, a human can do the job. 1-29

Figure 1-11 — Mechanical advantage. Mechanical advantage is usually considered with respect to work. Work represents the application of a force through a distance to move an object through a distance. Thus, it may be seen that there are two forces involved, each with an appropriate distance. These forces are shown by the simple lever in Figure 1-11. Refer to this figure as you read this section. Assuming perfect efficiency, the work input ( ) is equal to the work output ( ). Assuming equal distances and , a force of 10 pounds must be applied at the source to counteract a weight of 10 pounds at the load. If the fulcrum is moved nearer the load, less force is required to balance the same load. This is an example of the mechanical advantage of force. If force is applied to raise the load 1 foot, the source must be moved through a distance greater than 1 foot. Thus, mechanical advantage of force represents a mechanical disadvantage of distance. By moving the fulcrum nearer the source, these conditions are reversed. Ideally, the input work equals the output work (assuming no losses), and the mechanical advantage may be stated as a ratio of the force or of the distances. In actual situations, friction results in energy loss and decreased efficiency, thereby requiring an even greater input to accomplish the same work. Revolving Bodies Revolving bodies represent masses in motion; therefore, they possess all the characteristics and obey all the laws associated with moving bodies. In addition, since they possess a specific type of motion, they have special properties. Revolving bodies travel in a constantly changing direction, and they are constantly subjected to an accelerating force. Momentum tends to produce linear motion, but this is prevented by application of a force that restrains the object. This restraining force that prevents the object from continuing in a straight line is known as centripetal force. Remember Newton's third law of motion that states “for every force applied to a body, the body exerts an equal force in the opposite direction”; therefore, the centripetal force exerted on a revolving body must be opposed by an equal force that tends to produce linear motion. This second force is known as centrifugal force. The two forces, their relationships, and their effects are shown in Figure 1-12.

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Figure 1-12 — Forces on revolving bodies. The various forces involved in revolving bodies may be shown by using a ball and string. A slip knot is tied in the center of a 10 foot length of string to shorten the line to 5 feet; a rubber ball is attached to one end of the string. Holding the other end of the line, whirl the ball slowly in a circle. Note that the ball exerts a force against the hand (through the string); and to restrain the ball in its circular path, the hand must exert a force (through the string) on the ball. Now, make the ball revolve at a higher speed. Note that as the forces increase, the ball continues in a circular path. At some rotational speed, the forces involved become great enough to overcome inertial friction and the knot slips. At this time, allow the velocity of the rotation to stabilize (keep whirling the ball at the same speed) so you can analyze the existing conditions. When the knot slips, the ball is temporarily unrestrained and is free to assume linear motion in the direction of travel at that instant (tangent to the circle at the instantaneous position). The ball travels in a straight line until the string reaches its full length; during this time, no force is exerted on or by the hand. As soon as all the slack is taken up, there is a sharp jerk. An accelerating force is exerted to change the direction of motion from its linear path into a circular rotation. The ball again assumes rotational motion, but with an increase in radius. The ball does not make as many revolutions in the same time (rotational velocity is decreased), but it does maintain its former linear velocity. (The kinetic energy and the momentum of the ball have not changed.) Since the change in direction is less abrupt with a large radius than with a small one, less accelerating force is required, and the hand will feel less force. If you accelerate the ball to the same rotational velocity as you were doing just before the knot slipped, the linear velocity of the ball becomes much greater than before, as do the centripetal and centrifugal forces. In this example, the hand is fixed at a point that represents the center of rotation. This assumption is not actually accurate but it doesn't affect the general conclusions. For practical purposes, the two forces are equal at all points along the string at any given time, and the magnitude of each force is equal at all points along the string. In revolving or rotating bodies, all particles of the matter that are not on the axis of rotation are subjected to the forces just described. The statement is true whether the motion is through a complete circle or merely around a curve: An aircraft tends to skid when changing course; an automobile tends to take curves on two wheels. The sharper the curve (smaller radius), or the higher the velocity, the greater the tendency to skid. 1-31

HEAT Heat represents a form of energy. Therefore, it must be readily exchangeable with, or convertible into, other forms of energy. When a piece of lead is struck a sharp blow with a hammer, part of the kinetic energy of the hammer is converted into heat. In the core of a transformer, electrical and magnetic energy are exchanged; but due to hysteresis and eddy currents, some of the energy is lost as heat. These are some examples of the unwanted conversions, but many times the production of heat is desirable. Many devices are used to produce heat. Regardless of how or why heat is produced, it possesses certain characteristics that make it important to the technician. Knowledge of the nature and behavior of heat may help you understand the operation of some types of electronic equipment or to determine why equipment doesn't operate or operates incorrectly. Nature of Heat There are several theories regarding the nature of heat, none of which explain all the characteristics and properties exhibited by heat. The two theories most commonly included in discussions regarding the nature of heat are the kinetic theory and the radiant energy theory. In the kinetic theory, it is assumed that the quantity of heat contained by a body is represented by the total kinetic energy possessed by the molecules of the body. The radiation theory treats radio waves, heat, and light as the same general form of energy, differing primarily in frequency. Heat is considered as a form of electromagnetic energy involving a specific band of frequencies falling between the radio spectrum and light. A common method used to produce heat energy is the burning process. Burning is a chemical process in which the fuel unites with oxygen and a flame is usually produced. The amount of heat liberated per unit mass or per unit volume during complete burning is known as the heat of combustion of a substance. Through experimentation, scientists have found that each fuel produces a given amount of heat per unit quantity burned. Transfer of Heat There are three methods of heat transfer; conduction, convection, and radiation. In addition to these, a phenomenon called absorption is related to the radiation method of heat transfer. Conduction The metal handle of a hot pot may burn the hand; a plastic or wooden handle, however, remains relatively cool even though it is in direct contact with the pot. This phenomenon is due to a property of matter known as thermal conductivity. All materials conduct heat, some very readily, some to an almost negligible extent. When heat is applied to a body, the molecules at the point of application become violently agitated, strike the molecules next to them, and cause increased agitation. The process continues until the heat energy is distributed evenly throughout the material. Aluminum and copper are used for cooking pots because they conduct heat very readily to the food being cooked. Wood and plastic are used as handles because they are very poor conductors of heat. As a general rule metals are the best conductors of heat, although some metals are considerably better than others. 1-32

Figure 1-13 — Water is a poor conductor of heat. Among solids, there is an extremely wide range of thermal conductivity. In the original example, the metal handle transmits heat from the pot to the hand, with the possibility of burns. The wooden or plastic handle does not conduct heat very well, so the hand is given some protection. Materials that are extremely poor conductors are called insulators and are used to reduce heat transfer. Some examples are the wood handle of soldering irons, the finely spun glass or rock wool insulation in houses, or the wrapping used on steam pipes. Liquids are generally poorer conductors than metals. Look at Figure 1-13. The ice in the bottom of the test tube has not yet melted, although the water at the top is boiling. Water is such a poor conductor that the rate of heating of the water at the top of the tube is not sufficient to cause rapid melting of the ice at the bottom. Since thermal conduction is a process by which molecular energy is passed on by actual contact, gases are generally even poorer conductors than liquids because the molecules are farther apart and molecular contact is not so pronounced. A double pane window with airspace between the panes is a fair insulator. Convection Convection is the process by which heat is transferred by movement of a hot fluid. For example, an electron tube gets hotter and hotter until the air surrounding it begins to move. The motion of the air is upward because heated air expands in volume and is forced upward by the denser cool air surrounding it. The upward motion of the heated air carries the heat away from the hot tube by convection. Transfer of heat by convection is sped by using a ventilating fan to move the air surrounding a hot object. The rate of cooling of a hot vacuum tube can also be increased by providing copper fins to conduct heat away from the hot tube. The fins provide large surfaces against which cool air can be blown. A convection process may take place in a liquid as well as in a gas. One example is a transformer in an oil bath. The hot oil is less dense (has less weight per unit volume) and rises; while the cool oil falls, is heated, and rises in turn. When the circulation of gas or liquid is not rapid enough to remove sufficient heat, fans or pumps may be used to accelerate the motion of the cooling material. In some installations, pumps are used to circulate water or oil to help cool large equipment. In airborne installations, electric fans and blowers are used to aid convection. 1-33

Radiation Conduction and convection cannot wholly account for some of the phenomena associated with heat transfer. For example, heating through convection cannot occur in front of an open fire because the air currents are moving toward the fire. It cannot occur through conduction because the conductivity of the air is very low, and the cooler currents of air moving toward the fire would more than overcome the transfer of heat outward. Therefore, heat must travel across space by some means other than conduction and convection. The existence of another process of heat transfer is still more evident when the heat from the sun is considered. Since conduction and convection take place only through molecular contact within some medium, heat from the sun must reach the earth by some other method. (Outer space is an almost perfect vacuum.) Radiation is the name given to this third method by which heat travels from one place to another. The term radiation refers to the continual emission of energy from the surface of all bodies. This energy is known as radiant energy. It is in the form of electromagnetic waves and is identical in nature with light waves, radio waves, and x­rays, except for a difference in wavelength. Sunlight is a form of radiant heat energy that has traveled a great distance to reach the earth. These electromagnetic heat waves are absorbed when they come in contact with nontransparent bodies. The net result is that the motion of the molecules in the body is increased, as indicated by an increase in the temperature of the body. The differences in conduction, convection, and radiation are described as follows:  Although conduction and convection are extremely slow, radiation takes place with the speed of light. For example, an eclipse of the sun shuts off the heat from the sun at the same time as the light is shut off.  Radiant heat may pass through a medium without heating it. For example, the air inside a greenhouse may be much warmer than the glass through which the sun’s rays pass.  Although conducted or convected heat may travel in roundabout routes, radiant heat always travels in a straight line. For example, radiation can be cut off with a screen placed between the source of heat and the body to be protected. Absorption The sun, a fire, and an electric light bulb all radiate energy, but a body need not glow to give off heat. A kettle of hot water or a hot soldering iron radiates heat. If the surface is polished or light in color, less heat is radiated. Bodies that do not reflect are good radiators and good absorbers, and bodies that reflect are poor radiators and poor absorbers. For this reason, white clothing is worn in the summer. A practical example of the control of heat is the thermos bottle. The flask itself is made of two walls of silvered glass with a vacuum between them. The vacuum prevents the loss of heat by conduction and convection, and the silver coating reduces the loss of heat by radiation. The most effective color for heat transfer is dull black; dull black is the ideal absorber and also the best radiator. Thermal Expansion Nearly all substances expand or increase in size when their temperature increases. Railroad tracks are laid with small gaps between the sections to prevent buckling when the temperature increases in summer. Concrete pavement has strips of soft material 1-34

inserted at intervals to prevent buckling when the sun heats the roadway. A steel building or bridge is put together with red hot rivets so that when the rivets cool they shrink, and the separate pieces are pulled tightly together. As a substance is expanded by heat, the weight per unit volume decreases because the weight of the substance remains the same, while the volume is increased by the application of heat. Density decreases with an increase in temperature. Experiments show that for a given change in temperature, the change in length or volume is different for each substance. For example, a given change in temperature causes a piece of copper to expand nearly twice as much as a piece of glass of the same size and shape. For this reason, the connecting wires into an electronic tube are not made of copper but of a metal that expands at the same rate as glass. If the metal does not expand at the same rate as the glass, the vacuum in the tube is broken by air leaking past the wires in the glass stem. The metal usually used for this purpose is an alloy called Kovar . The amount that a unit length of any substance expands for a 1 degree rise in temperature is known as the coefficient of linear expansion for that substance. To estimate the expansion of any object, such as a steel rail, you need to know three things about the object; its length, the rise in temperature to which it is subjected, and its rate of coefficient of linear expansion. You can use the following equation to find the amount of expansion. Expansion = coefficient x length x rise in temperature or, e = k l ( - ) where k is the coefficient of expansion for the substance (In some instances, the Greek letter alpha (α) is used to indicate the coefficient of linear expansion.). l represents the length, and minus is the difference of the two temperatures. Use the equation shown above to solve the following problem: If a steel rod measures exactly 9 feet at 21°C, what is its length at 55°C? (The coefficient of linear expansion for steel is 11 x .) If the equation e = kl ( - ) is used, then; e = (11 x ) x 9 x (55 – 21) e = 0.000011 x 9 x 34 e = 0.003366 This amount, when added to the original length of the rod, makes the rod 9.003366 feet long. (Since the temperature has increased, the rod is longer by the amount of e. If the temperature had been lowered, the rod would have become shorter by a corresponding amount.) The increase in the length of the rod is relatively small; but, if the rod were placed where it could not expand freely, there would be a tremendous force exerted due to thermal expansion. You can see that thermal expansion is a consideration when designing ships, buildings, and all forms of machinery. Table 1-5 is a list of the coefficients of approximate linear expansion of some substances per degree Celsius.

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Figure 1-14 — Thermostat. Table 1-5 — Linear Expansion Coefficients SUBSTANCE COEFFICIENTS OF APPROXIMATE LINEAR EXPANSION Aluminum 24 x Brass 19 x Copper 17 x Glass 4 to 9 x Kovar 4 to 9 x Lead 28 x Iron, Steel 11 x Quartz 0.4 x Zinc 26 x

As an AE, you can see the practical application for the differences in the coefficients of linear expansion in the thermostat (Figure 1-14). This instrument is made from two strips of dissimilar metal (compound bar) fastened together. When the temperature changes, the bar bends because of the unequal expansion of the metals. Thermostats are used in overload relays for motors, in temperature sensitive switches, and in electric ovens. The coefficient of surface or area expansion is approximately twice the coefficient of linear expansion. The coefficient of volume expansion is approximately three times the coefficient of linear expansion. It is interesting that if a plate contains a hole, the area of the hole expands at the same rate as the surrounding material. When a volume is enclosed by a thin solid wall, the volume expands at the same rate as the solid body containing the volume of material. 1-36

Tabulation of Heat A unit of heat must be defined as the heat necessary to produce some agreed upon standard of change. The internationally accepted unit in common use is the Btu. One Btu is the quantity of heat necessary to raise the temperature of 1 pound of water 1°F. The terms quantity of heat and temperature are commonly misused. The distinction between them should be understood clearly. For example, place two identical pans containing different amounts of water of the same temperature, over identical gas burner flames for the same length of time. At the end of that time, the smaller amount of water will have reached a higher temperature. Equal amounts of heat have been supplied, but the increases in temperatures are not equal. In another example, if the water in both pans is the same temperature, say 80°F, and both are to be heated to the boiling point, more heat must be supplied to the larger amount of water. The temperature rises are the same for both pans, but the quantities of heat necessary are different. Mechanical Equivalent Mechanical energy is usually expressed in joules, or foot-pounds. Energy in the form of heat is expressed in calories or in Btu. In a precise experiment in which electric energy is converted into heat in a resistance wire immersed in water, the results show that 4.184 joules equals 1 thermochemical calorie, or that 778 foot-pounds equals 1 Btu. The following equation is used when converting from the English system to the metric system: 1 Btu = 252 calories. Specific Heat One way that one substance differs from another is that they require different quantities of heat to produce the same temperature change in a given mass of substance. The thermal capacity of a substance is the quantity, measured in calories, of heat needed per gram mass to increase the temperature 1°C. The specific heat of a substance is the ratio of its thermal capacity to the thermal capacity of water at 15°C. Specific heat is expressed as a number, which, because it is a ratio, has no units and applies to both the English and the metric systems. Water has a high thermal heat capacity. Large bodies of water on the earth keep the air and the surface of the earth at a fairly constant temperature. A great quantity of heat is required to change the temperature of a large lake or river. Therefore, when the temperature of the air falls below that of such bodies of water, they give off large quantities of heat to the air. This process keeps the atmospheric temperature at the surface of the earth from changing very rapidly. To find the heat required to raise the temperature of a substance, multiply its mass by the rise in temperature times its specific heat. For example, it takes 1,000 Btu to raise the temperature of 100 pounds of water 10°F, but only 31 Btu to raise 100 pounds of lead 10°F. Table 1-6 gives the specific heats of several common substances listed in descending order.

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Table 1-6 — Specific Heats of Some Common Substances Hydrogen (at constant pressure) 3.409 Water at 4°C 1.0049 Water at 15°C 1.0000 Water at 30°C 0.9971 Ice at 0°C 0.502 Steam at 100°C 0.421 Air (at constant pressure) 0.237 Aluminum 0.217 Glass 0.160 Iron 0.114 Copper 0.093 Brass, Zinc 0.092 Silver 0.057 Tin 0.056 Mercury 0.033 Gold, Lead 0.031 Change of State A thermometer placed in melting snow behaves in a strange manner. The temperature of the snow rises slowly until it reaches 0°C. Provided that the mixture is stirred constantly, it remains at the point until all the snow has changed to water. When all the snow has melted, the temperature again begins to rise. A definite amount of heat is required to change the snow to water at the same temperature. This heat is required to change the water from crystal form to liquid form. Heat of Fusion Heat of Fusion is the amount of heat required to convert a unit mass of a solid at its melting point into a liquid without an increase in temperature. Eighty gram calories of heat are required to change 1 gram of ice at 0°C to water at 0°C. In English units, the heat required to change 1 pound of ice at 32°F to water at 32°F is 144 Btu. These values (80 gram-calories and 144 Btu) are called the heat of fusion of water. The heat used while the ice is melting represents the work done to produce the change of state. Since 80 calories are required to change a gram of ice to water at 0°C, when a gram of water is frozen, it gives up 80 calories. Many substances behave very much like water. At a given pressure, they have a definite heat of fusion and an exact melting point. Many materials, however, do not change from a liquid to a solid state at one temperature. Molasses, for example, gets thicker and thicker as the temperature decreases; but there is no exact temperature at which the change of state occurs. Wax, celluloid, and glass are other substances that 1-38

do not change from a liquid to a solid state at any particular temperature. In fact, measurements of the glass thickness at the bottom of windows in ancient cathedrals tend to indicate that the glass is still flowing at an extremely slow rate. Most types of solder used in electronics maintenance also tend to become malleable before melting. Heat of Vaporization Damp clothing dries more rapidly under a hot, flat iron than under a cold one. A pool of water evaporates more rapidly in the sun than in the shade. It may be concluded that heat has something to do with evaporation. The process of changing a liquid to a vapor is like the process that occurs when a solid melts. Heat of Vaporization is the amount of heat required to convert a unit mass of a liquid at its boiling point into vapor without an increase in temperature. If a given quantity of water is heated until it evaporates, a much greater amount of heat is used than that necessary to raise the same amount of water to the boiling point. For example, 540 calories are required to change 1 gram of water to vapor at a temperature of 100°C. It takes 972 Btu to change 1 pound of water at 212°F to water vapor (steam) at 212°F. The amount of heat necessary for this change is called the heat of vaporization of water. Over five times as much heat is required to change a given amount of water to vapor than to raise the same amount of water from the freezing to the boiling point. The change from water to vapor occurs as follows. As the water molecules take up more and more energy from the heating source, their kinetic energy increases. The motion resulting from the high kinetic energy of the water molecules causes a pressure known as the vapor pressure. As the velocity of the molecules increases, the vapor pressure increases. The boiling point of a liquid is that temperature at which the vapor

pressure equals the external or atmospheric pressure. At normal atmospheric pressure at sea level, the boiling point of water is 100°C or 212°F. While the water is below the boiling point, a number of molecules acquire enough kinetic energy to break away from the liquid state into a vapor. For this reason some evaporation takes place below the boiling point. At or above the boiling point, large numbers of molecules have enough energy to change from liquid to vapor, and evaporation rapidly occurs. If the molecules of water are changing to water vapor in an open space, the air currents carry them away quickly. In a closed container, they rapidly become crowded and some of them bounce back into the liquid as a result of collisions. When as many molecules are returning to the liquid state as are leaving it, the vapor is said to be saturated. Experiments have shown that saturated vapor in a closed container exerts a pressure and has a given density at every temperature. LIGHT The exact nature of light is not fully understood, although men have been studying the subject for many centuries. Some experiments seem to show that light is composed of tiny particles, and some indicate that it is made up of waves. First, one theory, and then the other, attracted the approval and acceptance of the physicists. Today, there are scientific phenomena that are explained only by the wave theory and another large group of occurrences that are explained only by the particle or corpuscular theory. Physicists, constantly searching for some new discovery that would bring these contradictory theories into agreement, gradually have come to accept a theory concerning light that is a combination of these two views. 1-39

According to the view now generally accepted, light is a form of electromagnetic radiation; that is, light and similar forms of radiation are made up of moving electric and magnetic forces. Characteristics Light waves travel in straight lines. When they encounter any substance, they are transmitted, reflected, or absorbed. Those substances that permit clear vision through them, and that transmit almost all the light falling upon them, are said to be transparent. Those substances that allow the passage of part of the light, but appear clouded and impair vision substantially, are called translucent. Those substances that transmit no light are called opaque. Objects that are not light sources are visible only because they reflect part of the light reaching them from some luminous source. If light is neither transmitted nor reflected, it is absorbed or taken up by the medium. When light strikes a substance, some absorption and reflection always take place. No substance completely transmits, reflects, or absorbs all the light that reaches its surface. Luminous Intensity and Intensity of Illumination Though these two terms may sound like the description of the same property, they really aren't. Luminous intensity refers to the total light produced by a source, while intensity of illumination describes the amount of light received per unit area at a distance from the source. Some of the terms used to describe luminous intensity and intensity of illumination are defined in the following paragraphs. Candela – A unit of luminous intensity, in a given direction, of a source that emits mono­ chromatic radiation. Footcandle – The intensity of illumination of a surface (il luminance) is directly proportional to the luminous intensity of the light source and is inversely proportional to the square of the distance between the light source and the surface. Look at Figure 1-15, which shows the inverse square law of light. 1-40

Figure 1-15 — Inverse square law of light. If a card is placed 1 foot from a light source, the light striking the card is of certain intensity. If the card is moved 2 feet away, the intensity of light decreases with the square of the distance (2 x 2 or 4 times) and is 1/4 as bright. If the card is moved to 3 feet away from the light source, the intensity decreases the square of the distance (3 x 3 or 9 times), and the light is 1/9 the intensity it was at 1 foot. If the card is moved to 4 feet away, the light is 1/16 as intense as it was at 1 foot. The footcandle is one unit of measuring the intensity of incident light, and it can be computed by using the following formula: Illumination in footcandles =

A surface 1 foot from a 1 candlepower source would have an illumination of 1 footcandle, but if it was moved to a distance of 4 feet, a 16-candlepower source would be required for the same illumination. The inverse square law of light holds true for undirected light only; that is, light emissions not controlled by a reflector or lens. For light that is directed, the rate at which its intensity diminishes is dependent upon the rate or divergence of the beam.

Lumen – is the amount of light flowing through a solid angle of 1 steradian (sr) from a standard candle. A light source of 1 candlepower, placed in the center of a sphere that 1-41

has a radius of 1 foot, will illuminate every point on the surface of the sphere at an intensity of 1 footcandle. Then every square foot of the surface receives 1 lumen of light. The output of light bulbs may be given either in candlepower or in lumens, but since light bulbs do not distribute light equally in all directions, the lumen is most frequently used. It is customary for light bulb manufacturers to measure the light output in all directions and specify its total output in lumens. The common, gas filled, tungsten-filament light bulbs are usually more efficient in the larger sizes. For example, a 25-watt light bulb produces about 260 lumens (10.4 lumens per watt), while a 200-watt bulb produces 3,640 lumens (18.2 lumens per watt). Lux – is the illumination given to a surface 1 meter away from a 1 candlepower source and is sometimes called a meter-candle. Luminance – whether a body is self-luminous or just a reflector of the light that falls upon it luminance (or brightness) refers to the light a surface gives off in the direction of the observer. Reflection Light waves obey the law of reflection. Optical devices incorporated specifically for the purpose of reflecting light are generally classed as mirrors. They may be of a polished opaque surface, or they may be a specially coated glass. In the case of the glass mirror, there is some refraction as well as reflection; however, if the glass is of good quality and not excessively thick, the refraction will cause no trouble. The following discussion is based on a polished surface mirror. Several classes of mirrors are shown in Figure 1-16, views A, B, and C. All the devices work on the basis of the law of reflection. The applications of the law are briefly summarized here. Basically, the reflector is used to change the direction of a light beam (view A), to focus a beam of light (view B), or to intensify the illumination of an area (view C). In Figure 1-16, view A, the angle of the reflected light may be changed to some degree by changing the angle at which the incident light impinges upon the mirror. In Figure 1-16, view B, the focusing action of a concave mirror is indicated. The point of focus may be made any convenient distance from the reflector by proper selection of the arc of curvature of the mirror. The sharper the curvature, the shorter the focal length. In Figure 1-16, view C, the principle of intensification of illumination for a specific area is shown. The flashlight is an example of this application. In the system shown, the light source (bulb) is located approximately at the principal focus point, and that all rays reflected from the surface are parallel. Also, the reflector alone does not concentrate all the rays; some are transmitted without being reflected and are not included in the principal beam. 1-42

Figure 1-17 — The law of refraction.

Refraction As light passes through a transparent substance, it travels in a straight line. However, as it passes into or out of that substance, it is refracted in the same manner as other waves. Refraction of light waves results from the fact that light travels at slightly different velocity in different transparent media (Figure 1-17). To simplify the problem of understanding the action of light refraction and to make it possible to predict the outcome of specific applications, many transparent substances have been tested for refractive effectiveness. The ratio of the speed of light in air to its speed in each transparent substance is called the index of refraction for that substance. For example, light travels about 1.5 times as fast in air Figure 1-16 — Reflectors of light. 1-43

Figure 1-18 — Passage of light through a prism. as it does in glass, so the index of refraction of glass is about 1.5. When using the law of refraction in connection with light, a "denser" medium refers to a medium with a higher index of refraction. Refraction through a piece of plate glass is shown in Figure 1-17. The ray of light strikes the glass plate at an oblique angle along path AB. If it were to continue in a straight line, it would emerge from the plate at point N; but obeying the law of refraction, it is bent toward the normal (RS) and emerges from the glass at point C. Upon entering the air, the ray does not continue on its path but is bent away from the normal (XY) and along the path CD in the air. If the two surfaces of the glass are parallel, the ray leaving the glass is parallel to the ray entering the glass. The displacement depends upon the thickness of the glass plate, the angle of entry into it, and the index of refraction for the glass. All rays striking the glass at any angle other than perpendicular are refracted in the same manner. In the case of a perpendicular ray, no refraction takes place, and the ray continues through the glass and into the air in a straight line. PRISMS – When a ray of light passes through a flat sheet of glass, it emerges parallel to the incident ray. This holds true only when the two surfaces of the glass are parallel. When the two surfaces are not parallel, as in a prism, the ray is refracted differently at each surface of the glass and does not emerge parallel to the incident ray. Figure 1-18, view A, shows that both refractions are in the same direction, and that the ray coming out of the prism is not parallel to the ray going into it. The law of refraction explains what has happened. When the ray entered the prism, it was bent toward the normal; and when it emerged, it was bent away from the normal. Notice that the deviation is the result of the two normals not being parallel. If two triangular prisms are placed base to base (Figure 1-18, view B), parallel incident rays passing through them are refracted and caused to intersect. The rays passing different parts of the prisms, however, do not intersect at the same point. In the 1-44

Figure 1-19 — Positive lenses. ca se of two prisms, there are only four refracting surfaces. The light rays from different points on the same plane are not refracted to a point on the same plane behind the prism. They emerge from the prisms and intersect at different points along an extended common baseline, as shown in Figure 1-18, view B, points A, B, and C. Parallel incident light rays falling upon two prisms that have been joined apex to apex (Figure 1-18, view C) are spread apart. The upper prism refracts light rays toward its b ase; and the lower prism refracts light rays toward its base, causing the two sets of rays to diverge. POSITIVE LENSES – A positive (convergent) lens acts like two base-to- base prisms with their surfaces rounded off into a curve. Rays that strike the upper half of the lens bend downward, and rays that strike the lower half bend upward. A good lens causes all wavelengths within each ray to cross at the same point behind the lens (Figure 1-19, view A). When the incident ray of light enters the denser medium (the lens), it bends toward the normal. When it passes through into the less dense medium (the air), it bends away from the normal. Look at Figure 1-19, view B. It shows the refraction of only one ray of light, but all rays passing through a positive lens are affected in the same way. All incident light rays, either parallel or slightly diverging, will converge to a point after passing through a positive lens. The only ray of light that can pass through a lens without bending is the ray that strikes the first surface of the lens at a right angle, perpendicular or normal to the surface. It passes through that surface without bending and strikes the second surface at the same angle. It leaves the lens without bending. This ray is shown in Figure 1-19, view B. Positive lens and convergent lens are synonymous terms, since either of them may be used to describe the action of a lens that focuses (brings to a point of convergence) all light rays passing through it. All simple positive lenses are easy to identify since they are thicker in the center than at the edges. The three most common types of simple positive lenses are shown in Figure 1-19, view C. 1-45

Figure 1-20 — Negative lenses. NEG ATIVE LENSES – Figure 1- 18, view C, shows the refraction of light rays by two prisms apex to apex. If the prism surfaces are rounded, the result is a negative (divergent) lens. A negative lens is called a divergent lens since it does not focus the rays of light passing through it. Light rays passing through a negative lens diverge or spread apart, as shown in Figure 1-20, view A. Figure 1-20, view B, applies the law of refraction to one ray of light passing through a negative lens. Just as in a positive lens, a ray of light passing through the center of a negative lens is not affected by refraction and passes through without bending. Three simple negative lenses are shown in Figure 1-20, view C. They are often referred to as concave lenses and are readily identified by their concave surfaces. The simple negative lenses are thicker at the edges than at the center. They are generally used in conjunction with simple positive lenses where their primary use is to help form a sharper image by eliminating or subduing various defects present in an uncorrected simple positive lens.

1-46

Figure 1-21 — Electromagnetic wavelengths and the refraction of light. FREQUENCIES AND COLOR The electromagnetic waves that produce the sensation of light are very high in frequency, which means that they have very short wavelengths. These wavelengths are measured in nanometers (millionths of millimeters). Figure 1-21 indicates that light with a wavelength of 700 nanometers is red, and that a light with a wavelength of 500 nanometers is blue-green. This drawing is subject to erroneous interpretation. In actual fact, the color of light is dependent on its frequency, not its wavelength. When the wavelength of 700 nanometers is measured in air, it produces a color known as red, but the same wave measured in another medium has a wavelength other than 700 nanometers. When red light that has been traveling through air enters glass, it loses speed and its wavelength becomes shorter or compressed, but it continues to be red. This illustrates that the color of light is dependent upon frequency and not on wavelength. The color scale in Figure 1-21 is based on the wavelengths in air. All color component wavelengths of the visible spectrum are present in equal amounts in white light. Variations in composition of the component wavelengths result in other characteristic colors. If a beam of white light is passed through a prism, as shown in Figure 1-21, it is refracted and dispersed into its component wavelengths. Each of these wavelengths react differently on the eye, which then sees the colors making up the visible spectrum. 1-47

NOTE These are not the primaries used in mixing pigments. The visible spectrum is a mixture of red, orange, yellow, green, blue, indigo, and violet. When the primaries (red, green, and blue) are mixed together in overlapping beams of light, white light results.

Furthermore, the complementary or secondary colors (magenta, yellow, and cyan) are shown by mixing any two of the primary colors in overlapping beams of light. If red and green light is mixed in equal intensities, it will make yellow light; mixing green and blue light produces blue-green light (cyan); and mixing blue and red light correctly produces magenta (a purplish red color).

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End of Chapter 1 BASIC PHYSICS Review Questions 1-1. What are the three broad categories of measurement?

A. English, Metrics, and Mass B. Magnitude, Direction, and Time C. Meter, Kilogram, and Second D. Foot, Pound, and Btu

1-2. What unit of measurement is normally used to express scientific measurements?

A. English B. Farad C. Metric D. Lumen

1-3. What element is being measured when using the term kilogram?

A. Distance B. Mass C. Time D. Heat

1-4. What is the difference between the mass of a body and the weight of a body?

A. The mass of a body changes in form but the weight of the body remains constant. B. The mass of a body and weight are constant at all times. C. Mass of a body does not change but the weight of a body changes when in a solid form. D. The mass of a body is constant no matter where the body is located; however its weight is not constant.

1-5. What is meant when a person is described as weighing 195 pounds?

A. That person has the pull of gravity of 195 grams. B. The mass of the body is equivalent to the volume of 195. C. That person has the same pull of gravity a mass of 195 would have located near sea level. D. That person has the same pull of gravity when in constant motion.

1-49

1-6. What term is defined as the work done in lifting 1 pound a distance of 1 foot against the force of gravity?

A. Foot-pound B. Inch-pound C. Force D. Energy

1-7. List the measurements included in the units of power?

A. Joule and Horsepower B. Force, Distance, and Time C. Watt, Foot-pound, and Newtons D. Velocity and Mass

1-8. What are the four types of temperature scales?

A. Celsius, Fahrenheit, Kelvin, and Rankine B. Hot, Cold, Fahrenheit, and Celsius C. Atmospheric, Liquid, Solid, and Gas D. Freezing, Boiling, Steam, and Vapor

1-9. What procedure is used to convert from the Fahrenheit scale to the Celsius scale?

A. Subtract the 32 degree difference and multiply the result by 5/9. B. First multiply the reading on the thermometer by 5/9 and then add 32 to the result. C. Add 32 degrees and multiply the result by 5/9. D. Add 40 to the temperature then multiply by 9/5.

1-10. What type of thermometer is usually used in the laboratory?

A. Liqu id B. Solid C. Gas D. Resistant

1-11. What principle is used in compound bar thermometers?

A. Pascals B. Resistance C. Coefficient of linear expansion D. Thermocouple

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1-12. What terms are used to describe sound?

A. Frequency and change B. Loudness and frequency C. Watts and audibility D. Bel and Decibel

1-13. What gives a substance its distinguishing chemical and physical characteristics?

A. The atom’s negative charge B. The arraignment of the atom’s neutral particle and the density of the hydrogen atom C. The combination and arrangement of the atom’s subatomic particles’ D. The physical characteristics of the atom

1-14. What is a balanced atom?

A. An atom that has more protons than electrons B. An atom that contains an equal number of protons and electrons C. An atom that contains more electrons than neutron D. An atom that has neither positive nor negative charge

1-15. How is the atomic number of an element determined?

A. By the number of electrons B. By the chemical reaction with the different elements C. By the amount of mixture D. By the number of protons in its nucleus

1-16. The outer electron shell of an element is completely filled. What type of element is this?

A. Electron B. Proton C. Inert D. Neutral

1-17. What is the smallest unit that forms a compound?

A. Molecule B. Element C. Neutron D. Hydrogen

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1-18. What are the states of matter?

A. Liquid, Mass, and Hardness B. Solid, Liquid, and Gas C. Solid and Liquid D. Gas, Vapor, and Solid

1-19. Which is a common property of solids?

A. Resistant to heat B. Cohesion and adhesion C. Material compound D. Mixture of properties

1-20. What is one of the main uses of absolute zero?

A. A natural source reference B. It’s a fundamental constant of physics C. Change in temperature changes volume D. To study kinetic gases

1-21. What are the absolute zero points on the Kelvin and Celsius scales?

A. 0 Kelvin and -100° Celsius B. 0 Kelvin and -273° Celsius C. 10 Kelvin and -300° Celsius D. 15 Kelvin and -373° Celsius

1-22. What does Charles' law state?

A. Temperature of an enclosed sample of gas was kept constant and the pressure doubled; the volume was reduced to half the former value. B. When temperature changes, the pressure changes. C. Constant temperature and volume of an enclosed gas remains constant. D. All gases expand and contract in direct proportion to the change in the absolute temperature, provided the pressure is held constant.

1-23. What is mathematically stated by the formula E = mc2?

A. The destruction of matter creates energy, and that the creation of matter requires an expenditure of energy. B. Anything that can occupy space and has weight or mass. C. Energy can be changed in form with no resultant change in the total quantity of energy. D. Matter can be converted from one form to another.

1-52

1-24. Why is force considered a vector quantity?

A. Force has a direction. B. Force has both direction and magnitude. C. The direction of force is constant. D. The mass remains constant and the amount of force is variable.

1-25. In the English system of measurement, what force is expressed in pounds?

A. The momentum of a moving object in a given direction B. When the mass of an object remains constant if movement remains constant C. Gravitational force exerted by the earth on the body, known as weight of that body, expressed in force units D. The gravitational force it takes to keep an object in one location

1-26. How is the density of a substance described?

A. Its mass and shape B. By the form of the substance when in its solid state C. Its total weight D. Its weight per unit volume

1-27. What term describes the energy of mass in motion?

A. Density B. Universal force C. Kinetic energy D. Volume

1-28. What branch of physics deals with force, mass, and motion?

A. Hydraulics B. Mechanics C. Kinetics D. Specific gravity

1-29. What point of an object is its center of gravity?

A. The point where the gravitational force is at its greatest B. The point that is farthest away from the largest mass area of an object C. The point where a single force is less than the gravitational direction of an object D. The point where a single force equal to the gravitational force and directed up, and sustains the body at rest

1-53

1-30. Generally, a gyro rotor ________ about its axis.

A. revolves B. is stationary C. rotates D. exceeds gravity

1-31. What type of force is an accelerating force applied to the center of gravity of a body so that the body is accelerated with no rotation?

A. Translational B. Radial C. Momentum D. Torque

1-32. What is the most common reason for efficiency loss in mechanical physics?

A. Motion B. Force C. Time D. Friction

1-33. What is the principle that allows man to accomplish work that he normally could not do?

A. Efficiency B. Mechanical advantage C. Kinetic energy D. Potential energy

1-34. What force prevents an object from continuing along a straight line?

A. Momentum B. Linear C. Centripetal D. Centrifugal

1-35. In the radiation theory, heat is generally treated the same as which form of energy?

A. Radio waves B. Absorption C. Kinetic theory D. Insulators

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1-36. What are the three methods of heat transfer?

A. Absorption, radiation, and insulation B. Contact, thermal, and air space C. Conductivity, absorption, and air D. Conduction, convection, and radiation

1-37. In what state is matter the poorest conductor of heat?

A. Solid B. Liquid C. Gas D. Mass

1-38. What principle is involved in temperature-sensitive switches?

A. Refraction of thermal energy B. Coefficient of linear expansion C. Change of state D. Heat of vaporization

1-39. What effect does the heat of fusion have on solder?

A. It causes it to become malleable before it melts; that is, it flows at a very slow rate. B. It causes the solder to become brittle as the temperature increases. C. It causes the solder to have an exact melting point. D. It causes the solder to be thicker at the bottom.

1-40. What is meant by the term luminous intensity?

A. The constant color level of the light source B. The maximum angle in degree of reflected light without loss of brightness C. The total light produced by a source D. The reflected light without loss of color

1-41. What term is usually used to describe the output of a light bulb?

A. Lumen B. Watt C. Joule D. Vector

1-55

1-42. What are the principle uses of reflectors?

A. To focus in one area B. Focus in two directions simultaneously without loss of directional light C. To intensify the illumination of an area only D. Focus a beam of light, change direction of a light beam, or intensify the illumination of an area

1-43. Which are the primary colors of light frequencies?

A. Magenta, yellow, and cyan B. Green, blue, and yellow C. Red, orange, and yellow D. Red, green, and blue

1-44. What is the result if you mix the primary colors together?

A. Yellow light B. Blue-green light C. White light D. Purple light

1-56

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1-57

CHAPTER 2 ELECTRICAL MAINTENANCE AND TROUBLESHOOTING

Aircraft maintenance falls into two broad categories; scheduled maintenance and unscheduled maintenance. Scheduled maintenance is the action taken to reduce or eliminate failure and prolong the useful life of the equipment. Unscheduled maintenance is the action taken when a system part or component has failed, and the equipment is out of service. In maintenance work of any kind, you use two fundamentals; knowledge and skills. As an Aviation Electrician’s Mate (AE), you must have specific information about the particular equipment you repair or keep in good condition. Also, you must have certain general skills and knowledge that apply to many kinds of equipment and types of work assignments. The specific information required consists of special procedures and processes, and detailed step-by-step directions approved by proper authority. You can find this information in publications or checklists authorized by the Naval Air Systems Command (NAVAIR), type commanders, and other authorities. General maintenance skills and procedures are based on knowledge that is not in equipment manuals. These skills come from schools, On-the-J ob Training (OJT), and from training manuals. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Identify safety precautions regarding aircraft, personnel, material, and tools. 2. Restate troubleshooting techniques, including how to analyze, detect, and correct faults in electrical equipment. 3. Describe the various types of general-purpose test equipment associated with aircraft electrical maintenance, and identify tests the AE will make using these equipment. 4. Recognize aircraft wire and cable characteristics and various means of identifying and splicing wires and cables. 5. List the use and characteristics of aircraft electrical and mechanical hardware. 6. Explain the aircraft electrical motor and generator maintenance procedures. 7. Specify the characteristics of printed circuits, including modules and potted components, and recognize circuit construction features. 8. Describe the functions, capabilities, and operating characteristics of various types of aircraft test equipment. 9. Identify the hazards to Electrostatic Discharge ( ESD) sensitive devices and their preventative measures through proper handling and packaging techniques. 2-1

WARNING

Do not perform CPR unless you have had the proper training and certification. SAFETY In the performance of your duties, you come across many potentially dangerous conditions and situations. You install, maintain, and repair electrical and electronic equipment in confined spaces where high voltages are present. Among the hazards of this work are injury caused by electric shock, electrical fires, and harmful gases. Also, you must include improper use of tools among these hazards. Common sense and carefully following established rules will produce an accident-free naval career. When working, there is one rule to stress strongly, SAFETY FIRST. Whether you are working in the shop, on the line, or during a flight, you should follow prescribed safety procedures. When working on or near aircraft, there is the danger of jet blast, losing your balance, or being struck by propeller or rotor blades. Because of these dangers, you need to develop safe and intelligent work habits. You should become a safety specialist, trained in recognizing and correcting dangerous conditions and unsafe acts. Knowledge of first aid procedures is important. You must know how to treat burns and how to give artificial ventilation (respiration) to persons suffering from electric shock. In some cases, you may have to perform external heart compression in addition to artificial ventilation to restore the heartbeat. Artificial ventilation and external heart compression performed together are known as Cardiopulmonary Resuscitation (CPR). The life of a shipmate could easily depend upon your CPR and first aid skills. You should study the CPR training section of Basic Military Requirements (BMR), NAVEDTRA 14325. CPR training is a MANDATORY requirement for all AE's. Personnel could also benefit from additional first aid training. CPR training and certification is available at most commands, many Navy medical facilities, as well as local fire stations and Red Cross or American Heart Association agencies. It is important for you to be currently certified in the special skills of CPR. You must remain current on your CPR certification.

Cooperation of personnel and being vigilant prevents most accidents that occur in noncombat operations. You can learn more about general aviation safety by studying Aviation Electricity and Electronics— Maintenance Fundamentals Reference, NAVEDTRA 14318. General Precautions Only authorized personnel can repair and maintain electronic and electrical equipment because of the chance of injury, the danger of fire, and possible material damage. When you work on electrical equipment, open and tag the circuit breakers, main supply s witches or cutout switches. The tag should read as follows: “This circuit is open for repairs and shall not be closed except by direct order of ___________ (usually the person directly in charge of the repairs).” Securely cover fuse boxes and junction boxes, except when you are working on them. Don’t alter or disconnect safety devices, such as interlocks, overload relays, and fuses except when replacing them. Never change or modify safety or protective devices in any 2-2

way without proper authorization. Remove and replace fuses only after the circuit is de- energized. If a fuse blows, replace it with a fuse of the same current rating only. When possible, carefully check the circuit before making the replacement since the burned-out fuse often results from a circuit fault. You should move slowly when working around electrical equipment. Maintain good balance and DO NOT lunge after falling tools. DO NOT work on electrical equipment if you are mentally or physically exhausted. DO NOT touch energized electrical equipment when standing on metal, damp, or other well-grounded surfaces. DO NOT handle energized electrical equipment when you are wet or perspiring heavily. DO NOT TAKE UNNECESSARY RISKS. Some general safety precautions that you should follow are shown below:  REPORT ANY UNSAFE CONDITION, or any equipment or material that you consider to be unsafe, to the immediate supervisor.  WARN OTHERS you believe to be endangered because of known hazards or who fail to follow safety precautions.  WEAR or USE APPROVED PROTECTIVE CLOTHING or EQUIPMENT for the safe performance of work or duty.  REPORT to the supervisors ANY INJURY or evidence of impaired health occurring during work or duty.  EXERCISE REASONABLE CAUTION during any unforeseen hazardous occurrence, as is appropriate to the situation. The safety precautions that apply to the work of Aviation Electrician’s Mates include those you should follow when working in and around aircraft and in the electrical or battery shop. In addition to these, you need to know the authorized fire-fighting methods for electrical fires, procedures for treatment of burns, and artificial respiration. Some of the publications you should be familiar with are listed below:  Airman, NAVEDTRA 14014  Navy Safety and Occupational Health (SOH) Program Manual for Forces Afloat, OPNAVINST 5100.19 (series)  Navy Safety and Occupational Health Program Manual, OPNAVINST 5100.23 (series) These publications contain a variety of operations and functions in the Navy; therefore, they are basic and general in nature. Activities use these manuals as a basis for establishing specific safety instructions for their particular equipment, weapons system, or locality. Precautions Regarding Aircraft As an AE, you are exposed to flight line hazards. You will be working around moving machinery, which is dangerous. Therefore, you need to be alert when working around aircraft. Always follow your activity’s instructions on the application of external power. Make sure you get a thorough safety indoctrination from your supervisor. The Maintenance Instructions Manual (MIM) for each type of aircraft has an illustration, such as that shown in Figure 2-1 for the F/A-18 aircraft. Study the illustrations for each aircraft in your operating area. Most safety instructions require the anti-collision light to 2-3

Figure 2-1 — Radiation, intake, exhaust, and turbine blade failure danger areas. be operating whenever the engine or engines are operating. This gives an additional warning so you will be aware of propellers, rotors, or jet intakes and exhausts. Precautions Regarding Personnel and Materials If possible, you shouldn’t make repairs on energized circuits. When repairs on operating equipment are necessary, only experienced personnel under the supervision of a senior AE should do the work. Follow every known safety precaution carefully. Make sure there is enough light for good illumination, and there is insulation from ground, using a suitable non-conducting material. Station helpers near the main switch or the circuit breaker to de-energize the equipment immediately in case of emergency. While making the repair, someone qualified in CPR should be standing by in case of injury due to electrical shock. High-Voltage Precautions NEVER work alone near high- voltage equipment. Don’t use tools and equipment containing metal parts within 4 feet of high- voltage circuits or any wiring having exposed surfaces. The handles of all metal tools, such as pliers and cutters, should have rubber insulating tape covers. Always use authorized tools, utilization of locally manufactured tools and insulation devices are explicitly unauthorized except as prescribed by applicable instructions and manuals. Before touching a capacitor, short-circuit the terminals to discharge the capacitor completely. Permanently attach grounded shorting prods to workbenches where electrical devices receive regular servicing. Don’t work on any type of electrical apparatus with wet hands or while wearing wet clothing. Don’t wear loose or flapping clothing. Don’t wear thin-soled shoes with metal plates or hobnails; wear safety shoes with non-conducting soles when available. Don’t wear flammable articles. 2-4

Before you work on an electrical or electronic apparatus, remove all rings, wristwatches, bracelets, and similar metal items. Make sure that your clothes do not contain exposed zippers, metal buttons, or any type of metal fastener. Make sure warning signs and suitable guards are posted to prevent personnel from coming into accidental contact with high voltages. Low-Voltage Precautions Most people never realize the dangers of low-voltage electric shock. These hazards are present and dangerous. You need to be aware of their existence and treat low voltage equipment with the same precautions as high voltage equipment. You should be aware of any voltage greater than 15 volts. Degree of Shock The current that may pass through the body without causing damage depends on the individual, and the type, path, and length of contact time. The resistance of your body varies. For example, if the skin is dry and unbroken, body resistance will be quite high, on the order of 300,000 to 500,000 ohms. However, if the skin becomes moist or broken, body resistance may drop to as low as 300 ohms; a potential as low as 30 volts could cause a fatal current flow. Therefore, any circuit with a higher value potential is more dangerous. If a 60-hertz alternating current passes through the chest cavity, it has the following effects:  At 1 milliampere (0.001 ampere), you will feel the shock.  At 10 milliamperes (0.01 ampere), the shock paralyzes muscles, and a person may be unable to release the conductor.  At 100 milliamperes (0.1 ampere), the shock is usually fatal if it lasts for 1 second or more. It is important to remember that, fundamentally, current, rather than voltage, is the criterion of shock intensity. First Aid for Electric Shock Electric shock produces a jarring, shaking sensation. The victim usually feels like he/she just received a sudden blow. If the voltage and resulting current is high enough, the victim may become unconscious. Severe burns may appear on the skin at the place of contact. Muscular spasm may occur, causing the victim to clasp the apparatus or wire causing the shock. If this happens the victim is unable to release it. Use the following procedures for rescuing and caring for shock victims: 1. Remove the victim from electrical contact at once. DO NOT ENDANGER YOURSELF. Remove the victim by throwing the switch if it is nearby, or cut the cable or wires to the apparatus using an axe with a wooden handle. Protect your eyes from the flash when the wires are severed. Also, you can use a dry stick, rope, belt, coat, blanket, or any other nonconductor of electricity to drag or push the victim to safety. 2. Determine whether the victim is breathing. Keep the person lying down in a comfortable position and loosen the clothing about the neck, chest, and abdomen for easy breathing. Protect from exposure to cold, and watch closely. 3. Keep the victim from moving. In this condition, the heart is very weak. Any sudden muscular effort or activity of the patient may result in heart failure. 2-5

WARNING

Never use a solid stream of water to extinguish electrical fires in energized equipment. 4. Do not give stimulants or opiates. Send for a medical doctor at once, and do not leave the patient until adequate medical care is available. 5. If the victim is not breathing, apply artificial ventilation without delay, even though the patient may be lifeless. Do not stop artificial respiration until the victim revives, or proper authority pronounces the victim is beyond help. Electrical Fires The three general classes of fires are Alpha, Bravo, and Charlie. Class Alpha fires involve wood, paper, cotton and wool fabrics, rubbish, and the like. Class Bravo fires involve oil, grease, gasoline and aircraft fuels, paints, and oil-soaked materials. Class Charlie fires involve insulation and other combustible materials in electrical and electronic equipment. Electrical or electronic equipment fires are caused by overheating, short circuits, friction (static electricity), or radio-frequency arcs. Also, equipment may ignite from exposure to nearby class Alpha or Bravo fires. Since class Charlie fires involve electrical circuits, electrical shock is an added hazardous condition. Whenever possible, immediately de- energize any electrical equipment exposed to class Alpha or class Bravo fires, or ignited by such a fire. If the equipment cannot be de-energized completely, use protective measures to guard against electrical shock. Extinguishing agents other than gases contaminate delicate instruments, contacts, and similar electrical devices. Carbon dioxide (CO2) is the preferred extinguishing agent for electrical fires. It does not conduct electricity, rapidly evaporates, and leaves little or no residue. It reduces the possibility of electrical shock to personnel and damage to equipment as a result of contamination. A dry chemical extinguishing agent, chiefly composed of potassium bicarbonate (more commonly known as Purple-K Powder or PKP), can be used on electrical fires. It is a nonconductor, which provides protection against electrical shock. However, damage to electrical or electronic parts may result from the use of this agent.

Water usually contains minerals that make it conductive. The conductivity of seawater is many times greater than that of fresh water. Pure distilled water is not a good electrical conductor, and it is suitable for emergency use on small electrical fires. If you must use fresh water or seawater, use a fog head or tip hose nozzle in electronic equipment spaces. The fog is a fine diffusion or mist of water particles, with very little conductivity. However, there is still a danger of electric shock unless the equipment is completely de- energized. Also, condensation from the fog frequently damages the electrical components. Be careful when using the fire hose. Pressure at the fireplug may be as much as 100 pounds per square inch (PSI). An unrestrained fire hose may result in whiplash for the person holding it. Do not use foam on electrical fires. Foam can damage equipment and possibly pose a shock hazard to personnel. If necessary, you can use foam on de-energized circuits. When a blanket of foam goes on a burning substance, the foam smothers the fire, cutting off the air supply to the burning substance. As the supply of oxygen decreases, the fire dies out. 2-6

When fighting electrical fires, you should use the following general procedures: 1. Promptly de-energize the circuit or equipment affected. 2. Sound an alarm according to station regulations or ship’s fire bill. When ashore, notify the fire department; if afloat, notify the officer of the deck. Give the fire location, and state what is burning. If possible, report the extent of the fire. 3. Close compartment air vents or windows. 4. Control or extinguish the fire using a C02 fire extinguisher. 5. Avoid prolonged exposure to high concentrations of carbon dioxide in confined spaces. You can suffocate in confined spaces unless you’re using special breathing apparatus. 6. Administer artificial ventilation and oxygen to a person overcome by carbon dioxide fumes, and keeps the victim warm. Even under normal conditions, fire aboard a Navy vessel at sea can kill and injure more people and cause more ship damage than battle. All personnel need to know the dangers of fire. You need to know the type and location of all fire-fighting equipment and apparatus in your immediate working and berthing spaces and throughout the ship. Volatile Liquids Prior to use of any chemical compound ensure that you have reviewed and understand all hazards, precautions and treatments provided by the Material Safety Data Sheets (MSDS) for that compound. MSDSs will be available in a central location for all compounds used within a work center or command spaces, see your supervisor or Hazardous Materials (HAZMAT) Program Manager for details pertaining to your command. Volatile liquids, such as insulating varnish, lacquer, turpentine, and kerosene, are dangerous when used near operating electrical equipment because sparks from the equipment can ignite the fumes. When these liquids are used in compartments containing non-operating equipment, make sure there is enough ventilation to avoid an accumulation of fumes. Also, make sure the space is clear of all fumes before energizing the equipment. When cleaning of equipment or spaces is required use only authorized cleaning materials and quantities. Today, all aviation fuels are hydrocarbons. Handling hydrocarbon products is hazardous because of their flash point. Products such as gasoline, solvents, and most crude oils begin to vaporize at or below 80°F; their flash point is reached at 80°F. Their flash point makes them the most hazardous petroleum products to handle. Other petroleum products, such as kerosene and lubricating oils, have a flash point above 80°F, making them less hazardous. The vapors of petroleum, gasoline, and other petroleum products cause drowsiness when inhaled. Petroleum vapors in concentrations of 0.1 percent can cause dizziness to the extent of inability to walk straight after 4 minutes of exposure. Longer exposure and/or greater concentrations may cause unconsciousness or death. The first symptoms of exposure to toxic (poisonous) vapors are headaches, nausea, and dizziness. When working in an area where there are possible toxic vapors, stay alert. If you get a headache, become dizzy, or become nauseous, you might be exposed. You should leave the area and report the condition. 2-7

CAUTION Compressed air is a special tool; do not use as a substitute for a brush to clean machines, clothing, or your person. In those cases where an air hose is essential for blowing out fixtures and jigs, wear eye protection and maintain air pressure below a maximum of 30 PSI. It is also desirable to place screens around work to confine the blown particles. You recover from early symptoms quickly when you move to an area having fresh air. If you find people overcome by vapors, get them immediate medical attention. First aid c onsists of the prevention of chilling, and, if breathing has stopped, artificial respiration. If gasoline remains in contact with your skin, it may irritate the skin, particularly under soaked clothing or gloves. Repeated contact with gasoline removes protective oils from the skin, causing drying, roughening, chapping, and cracking, and in some cases, infection. Remove clothing or shoes soaked with gasoline at once. Wash gasoline from the skin with soap and water. While removing your clothes, an arc, caused by static electricity, can cause the fuel to ignite. For this reason, remove fuel soaked clothes in a running shower. If a person swallows gasoline, give first aid immediately. You should give the victim large amounts of water or milk and 4 tablespoons of vegetable oil, if available. Do not induce vomiting. Get medical attention immediately. JP-4 fuel shares some of the characteristics of gasoline. Due to other characteristics, such as lower vapor pressure and high aromatic content (compounds added to increase the performance number), handle this fuel very carefully. JP-5 is a kerosene type of fuel. It has a vapor pressure close to 0 PSI. Since its tendency to vaporize is lower than the more volatile grades, the vapor-air mixture above its liquid surface is too lean to ignite. For ignition to occur, the liquid’s surface must reach 140°F. Handle this fuel with care. Take precautions to prevent personnel from breathing fumes from any fuel. Also prevent fuel from coming in contact with the skin, especially if the skin has abrasions or sores. Compressed Air Compressed air, when misused, is dangerous. Injuries occur through hose or fitting failures, causing the hose to whip dangerously and propelling fitting parts through the air. Blown dust and small particles become an eye hazard. Air under pressure may cause internal injury or even death by introducing an airstream into body tissue, usually through an existing cut or scratch. Air can rupture cell tissues and cause severe wounds through injection of minute foreign bodies into the skin. Impurities always exist in a shop’s air supply. Falls can result from tripping over air lines thoughtlessly left lying on the floor.

The National Safety Council has published the following general safety rules for working with compressed air: 1. Use only sound, strong hose with secure couplings and connections. 2. Make sure there aren’t any sharp points on metal hose parts. 2-8

NOTE Do not etch beryllium-copper tools. 3. Close the control valve in portable pneumatic tools before turning on air. 4. Turn off air at the control valve before changing pneumatic tools. Never kink a hose to stop the airflow. 5. Wear suitable goggles, mask, protective clothing, or safety devices. 6. Never use air to blow dust chips from work clothing or from workbenches. 7. Never point the hose at anyone. Practical jokes with compressed air have caused painful deaths. 8. When using compressed air, see that nearby workers are not in line of airflow. Precautions Regarding Tools The tools you use shall conform to Navy standards as to quality and type. You should use them only for the purposes of their design. Maintain tools in good repair, and turn in all damaged or nonworking tools. As an AE, you will use hand tools. By using them correctly, you will improve the quality of maintenance and reduce the chance of failures. It is recommended that AE's should complete Tools and Their Uses, NAVEDTRA 14256. Carelessness is the biggest menace in any shop. A machine doesn’t inflict injury. Operator inattention is the cause of most accidents in electrical and electronics shops today. Remember, all moving machinery is dangerous. It’s not safe to lean against any machine that is started or may start to move. Don’t start a machine until you fully understand its operation. Treat a machine with respect; there is no need to fear it. Always follow the two basic safety precautions stated below: 1. Use the proper tool for its intended function, and use it correctly. 2. Keep all tools in working order and in a safe condition. Replace dulled cutting tools. Protect tools from damage while in use or in stowage. If a tool becomes worn, damaged, or broken, report it to the work center supervisor. Always clean, inspect, and account for all tools after you complete a job. Return all tools to their proper stowage place. Nonmagnetic Tools You can get tools made of nonmagnetic materials through normal supply channels. You will use them to maintain equipment that can be damaged if tools become magnetized. Normally, these tools are made from beryllium-copper or plastic. They aren’t as rugged as steel tools, and can be damaged easily. Use them properly; they will last longer and operate properly.

When working near compasses and other components containing permanent magnets, always use nonmagnetic tools. Magnetic susceptible tools could become magnetized and transfer this magnetic condition to other equipment. 2-9

WARNING

Unauthorized modification of tools may present hazards due to unknown insulating material limitations and properties. Insulated Tools Often, safety considerations require use of insulated tools. Authorized insulated tools are available through supply. Only utilize authorized tools; do not modify tools except as specifically prescribed by applicable instructions or maintenance manuals.

Power Tools Working as an AE, you will use shop machinery, such as a power grinder or drill press. In addition to the general precautions on the use of tools, there are a few other precautions to follow when working with machinery. Some of the precautions are as follows: 1. Never operate a machine with a guard or cover removed. 2. Never operate mechanical or powered equipment unless you know how to operate them. When in doubt, consult the appropriate instruction or ask someone who knows. 3. Always make sure that everyone is clear before starting or operating mechanical equipment. 4. Cut off the source of power before trying to clear jammed machinery. 5. Always keep everyone clear when hoisting heavy machinery or equipment by a chain fall. Guide the hoist with lines attached to the equipment. 6. Never plug in electric machinery without knowing that the source voltage is the same as that called for on the nameplate of the machine. Carefully inspect all portable power tools to be sure they are clean, well-oiled, and in working order before you use them. The switches should operate normally, and the cords should be clean and free of defects. Ground the casings of all electrically driven tools. Do not use sparking portable electric tools in any place where flammable vapors, gases, liquids, or exposed explosives are present. Check to make sure that power cords do not come in contact with sharp objects. Don’t let cords kink. Don’t leave them where they might be run over. Don’t let cords contact oil, grease, hot surfaces, or chemicals. When damaged, replace power cords. When unplugging power tools from receptacles, you should grasp the plug, not the cord. Soldering Iron The soldering iron is a potential fire hazard and a source of burns. Always assume a soldering iron is hot. Never rest the iron anywhere but on a metal surface or rack provided for that purpose. Keep the iron holder in the open to reduce the danger of fire from accumulated heat. Don’t shake the iron to get rid of excess solder. The hot solder may strike people, causing burns, or the equipment, causing a short circuit. Hold small soldering jobs with pliers or clamps. 2-10

When you are cleaning the iron, place the cleaning rag on a suitable surface and wipe the iron across it. Don’t hold the rag in your hand. Disconnect the iron when leaving the work area, even for a short time; the delay may be longer than planned. Significant lead exposures can occur during s oldering with torches. Lead has long been a recognized health hazard. Lead can damage the nervous system, kidneys and reproductive systems. Chronic lead exposure can initially damage the blood forming organs. Higher levels can result in reproductive dysfunction in both men and women, and it can cause peripheral nerve and central nervous system changes. Lead inhibits heme synthesis (red blood cell production) and at high levels leads to anemia. Grounding A poor safety ground, or one with incorrect wiring, is more dangerous than no ground at all. A poor ground is dangerous because it doesn’t offer full protection; it lulls you into a false sense of security. The incorrectly wired ground is a hazard because one of the live wires and the safety ground are transposed. When this happens the shell of the tool is hot the instant the plug connects into the outlet. You will get a shock unless the safety ground connects to the grounded side of the line on a single-phase grounded system, or no grounds are present on an ungrounded system. Today, a three-wire, standard, color- coded cord with a polarized plug and a ground pin is used. In all properly connected new tools, the green wire is the safety ground. This wire attaches to the tool’s metal case at one end and to the polarized grounding pin at the other end. It normally carries no current, and is in use only when the tool insulation fails. When this happens, the safety ground short-circuits the electricity around the user to ground and protects that person from shock. To check the grounding system resistance, you use a low-reading digital fluke meter to be certain the safety ground is adequate. If the resistance shows greater than 0.1 ohm, you should use a separate ground strap. Some old installations do not have receptacles that will accept the grounding plug. If you are working on these, use one of the following adaptations:  Use an adapter fitting.  Use the old type plug and bring the green ground wire out separately.  Connect an independent safety ground line. When you use an adapter, connect the ground lead extension to a good ground. Don’t use the center screw that holds the cover plate on the receptacle. Where separate safety ground leads are connected externally, connect the ground first, and then plug in the tool. Likewise, when disconnecting the tool, first remove the line plug, and then disconnect the safety ground. ALWAYS CONNECT THE SAFETY GROUND FIRST AND REMOVE IT LAST. Unscheduled Maintenance Unscheduled maintenance is the performance of a repair action without a set interval. Discrepancies found before, during, and after flight fall into this category. Some of the duties performed during unscheduled maintenance are shown below:  Visually check equipment for loose leads, improper connections, damaged or broken components, etc., before applying power to the equipment. This check applies particularly to new equipment preserved or stored for long periods and equipment exposed to the weather. 2-11

 Conduct a close visual inspection on O-rings, gaskets, and other types of seals when the equipment under check is a pressurized component. A visual inspection often reveals discrepancies that you can correct at that time with minimum labor and parts. Such discrepancies, if left uncorrected, might result in loss of an aircraft or a major maintenance problem. TROUBLESHOOTING TECHNIQUES You will spend most of your time troubleshooting the equipment in squadron aircraft. You will maintain many components and systems that are complex and difficult to troubleshoot. However, the most difficult troubleshooting job will become simple if you break it down into the following steps: 1. Review the system operation and components in the MIMs. 2. Analyze the symptom. 3. Detect and isolate the trouble. 4. Correct the trouble and test the work. Analyze If you don’t understand a system or component, you can’t find out what’s wrong with it. The following are tools that will help you analyze what’s wrong with a system or component: 1. MIMs 2. Schematics 3. Records on the equipment Detect and Isolate The first things to do when troubleshooting is to review the system in the MIM then visually inspect the component. If parts are obviously not in proper condition, correct these faults before going any further. The types of things to look for include open circuit breakers, improperly placed switches, burnt equipment, loose mountings, disconnected components, and dented equipment. If you can’t find anything after visually inspecting the equipment, check the condition of fuses and circuit breakers. Sometimes, a circuit protector will open. If this occurs reset or replace the circuit protector as required and apply power. If the protector opens again, secure the power because there is a probable circuit malfunction. Don’t apply power until after you correct the malfunction. Utilizing the applicable MIM, you can identify which pin in a connector provides the power input to the connector. You can use the MIM to close the correct circuit breakers and locate the switches to use for testing. If there is power to the component and it still does not work, then the fault is with the component itself. If a short, ground, or overload condition is not indicated, take power readings at the checkpoints shown in the MIMs. Common faults that interrupt power through a circuit are broken wiring, loose terminal or plug connections, faulty relays, and faulty switches. Look for these conditions when checking points along a circuit. If there is no power reaching the component, assume that the component is good, and start checking the power supply. Begin the check at the bus that supplies the power. If you find evidence of a short circuit or overload, secure the power. At this time, you need to make continuity and resistance checks. 2-12

NOTE Only Aviation Structural Mechanics (AMEs) work on oxygen lines. Continuity and Resistance Checks The process of fault detection often leads beyond visual inspection and power checks. You use a digital multimeter to find out if a power circuit is delivering power to the proper place. However, the digital multimeter volts selection only indicates the absence of voltage. The ohms selection on the digital multimeter can identify the type of fault, including opens, shorts, grounds, or incorrect resistance values. You use the schematic to trace circuits in components, part by part and wire by wire, until you isolate the trouble. You can use the MIM for the equipment to find the proper resistance readings. At intermediate maintenance activities, bench-test installations are used for off- equipment testing. In this type of installation, there is a complete and fully operational system with the components grouped closely together. This lets the troubleshooter remove any unit and replace it with the one for testing. Bench-testing components are used when the fault in the unit under test is hard to determine or when the unit is functioning, but does not measure up to minimum standards. Bench tests are highly specialized test installations, and only specific tests can be run on specific components. Normally, AEs troubleshoot using a meter and a schematic. Nonelectric Components So far, you have learned about fault detection in systems using electric power. However, a group of systems and components do not use electricity. This group includes such items as mechanical instruments, direct reading gauges, and mechanisms closely associated with electrical equipment. Such mechanisms include switch actuators, mounting assemblies, mechanical linkages, and any type of hardware that is part of an electrical system. Even though you cannot check this equipment with a digital fluke meter, the three basic rules for troubleshooting remain the same: analyze, detect, and correct. Correct and Test Many faults may occur, and for each fault you must perform a corrective action. There are rules that apply to practically all corrective actions. In some cases, the detection of a fault involves more time and labor than correcting it. Study the job and think through each step. Form a plan of attack, and decide which tools you need. In addition to hand tools, consider equipment, such as extension lights and cooling blowers. Refer to the MIM, which lists special tools you need and also describes how to gain access to a system component or area. Ask someone who has performed that particular job before; they can usually offer some very good pointers. As a troubleshooter, you should plan how to deal with special safety considerations you might meet during a job. Some of these are hydraulic oil drainage from disconnected lines, dangling and unprotected power cables, high-pressure air lines, and the handling of heavy components. Drain or bleed off gas or liquid lines, and handle heavy objects with special care.

If exposed power cables are unavoidably left loose, use some means to prevent their shorting out. You can do this by covering them with a temporary insulating shield. Place 2-13

a warning sign on the main power switch in the cockpit and on the external power receptacle. Where applicable, you may also disconnect the aircraft battery. In this case, place a sign in the cockpit stating the battery is disconnected. Complete all of the safety measures mentioned here before starting work. The nature of the job determines the procedures you will use to actually do the work. You will either repair or replace the component or system. Total job quality depends on the quality of work on smaller components/systems, such as soldering, replacing connecting devices, safety wiring, and using tools. Learn and practice good techniques. Working under the supervision of an experienced AE is one way to learn them. Another place to look for approved methods is in Installation and Repair Practices, NAVAIR 01-1A-505 (series). It contains the accepted way of performing many practical operations, and it is well illustrated. Often, parts replacement is the best way to correct a malfunction. However, never use parts replacement as a method of troubleshooting. Each year, thousands of dollars’ worth of instruments and electronic components are returned to overhaul activities with labels stating that they are faulty. Upon completion of test and check by the overhaul activity, many have no defect. This practice is wasteful and very expensive. The replacement of an entire unit when only a small part is at fault reflects poor maintenance practices. A typical example of such a case is the replacement of an entire compass amplifier when there is only a faulty fuse. When troubleshooting, remember you are working with equipment that is both expensive and scarce. Make decisions to replace equipment only after thoroughly testing it and making sure it is faulty. For the procedures to follow when replacing and testing equipment, refer to the removal, installation, and testing section of the MIM. Some of the topics covered in the MIMs are discussed in the following paragraphs. Removal – Remove equipment in such a way that you don’t damage it more and you don’t damage nearby structures. Pay attention so you don’t lose or misplace small parts. They may be easy to replace with new ones, but small items lost in the aircraft are a very real Foreign Object Damage (FOD) hazard. FOD kills! If you can’t finish installing the new part immediately, remove all tools and parts from the aircraft during the waiting period. You might forget and leave them in the aircraft. This precaution applies to any unfinished job involving a waiting period. Before you remove the defective part from the aircraft, pull and tag circuit protective devices and notify maintenance control. This prevents people from turning on fuel pumps when fuel lines are open. Also, it prevents power being applied to loose cables. Installation – Many of the rules for removal apply to installation. However, there are a few rules that apply almost exclusively to installing new equipment. Some units have special conditions while in shipment. Look for these conditions because they often require you to remove sealing plugs, locking devices, and other special equipment from the component before you can install it. If adjustments are necessary on new equipment, you might need to make the adjustments before completing the installation. For example, often the matching of a position transmitter to the position indicator is easier to do with only the electric wiring connected to the actuator. After calibration, complete the mechanical mounting. Again, the only guide for making adjustments to any system is the MIM. In the final installation, make sure that all mounting hardware is in place and properly tightened. Also, make sure all cable and line connections are secure and connected to the proper points. 2-14

Testing – The entire process of detecting and correcting a fault means nothing unless the system operates properly. Therefore, the final step of the job is to test the equipment. This test is usually an operational test. Where possible, simply operate the new component or system to see that it is doing its job properly. When a final check shows the job is complete, you must ensure the work is inspected by qualified personnel, maintenance/production control is notified and that all Work Orders or maintenance action forms are completed. Officially, the aircraft is down until the proper people are notified and the paperwork is complete. USE OF BASIC TEST EQUIPMENT Test equipment, like any other tool the AE uses, is susceptible to damage, misuse, and deterioration. Knowledge of the Navy calibration program is essential to ensure your shop or division has the most reliable test equipment. For more information on the c alibration program, refer to the latest edition of Aviation Maintenance Ratings Fundamentals, NAVEDTRA 10342-3, and Navy Electricity and Electronics Training Series (NEETS), Module 16, Introduction to Test Equipment, NAVEDTRA 14188. Also, when making tests on equipment, you should adhere to the following rules: 1. Always connect a digital fluke meter (current scale) in series —never in parallel. 2. Always connect a digital fluke meter (voltage scale) in parallel —never in series. 3. Never take resistance reading with energized circuitry. 4. On test meters, select the highest range first, and then switch to lower ranges as needed. 5. When you use a digital fluke meter, select a scale that gives a near midscale reading, since midscale is where the meter is most accurate. 6. Turn meter OFF when not in use. 7. Observe polarity when measuring dc voltage or current. 8. Do not place meters in the presence of strong magnetic fields. 9. Never measure the resistance of a meter or a circuit with a meter in it. The high current required for meter operation may damage the meter. This also applies to circuits with low-filament-current tubes and to some types of semiconductors. 10. When you are measuring high resistance, do not touch the test lead tips or the circuit. Doing so may cause body resistance to shunt the circuit, causing an erroneous reading. 11. Connect the ground lead of the meter first when making voltage measurements. Work with one hand whenever possible. Continuity Tests In open circuits, current flow stops. The cause may be a broken wire, defective switch, etc. To detect open circuits, you perform a continuity test, which will tell if the circuit is complete or continuous. To make a continuity test, you will use a digital fluke meter. Normally, you will make this test in a circuit where the resistance is low, such as the resistance of a copper conductor. If the resistance is very high or infinite, the resistance between the two points shows an open circuit. 2-15

Figure 2-2 — Continuity test.

Look at Figure 2-2. It shows a continuity test of a cable connecting two electronic units. You can see that both plugs are disconnected, and the digital fluke meter is connected in series with conductor A, which is under test. The power should be off. When checking conductors A, B, and C, the current from the digital fluke meter flows through plug 2 (female), conductor A, B, or C, to plug 1 (female). From plug 1, current passes through the jumper to the chassis, which is grounded to the aircraft structure. The aircraft structure serves as the return path to the chassis of unit 2 and completes the circuit through the series-connected digital fluke meter. The digital fluke meter digital display shows a low resistance (.001 Ohms) because no break exists in conductors A, B, or C. However, checking conductor D, as shown by the reading in the figure, reveals an open. The digital fluke meter shows maximum resistance because current cannot flow in an open circuit. With an open circuit, the digital fluke meter digital display reads 0.L megohms of resistance. If you can’t use the aircraft structure as the return path, use one of the other conductors known to be good. This technique will also reveal the open in the circuit.

Grounded Circuit Test Grounded circuits happen when some circuit part makes contact either with the aircraft metallic framework or with a metal object that is a ground. The most common cause of grounds is insulation fraying from a wire, allowing the conductor to come in contact with aircraft structure. When testing for grounds, you use a digital fluke meter or some other continuity tester. By measuring the resistance to ground at any point in a circuit, you can determine if the point goes to ground. Look at Figure 2-2. After removing the jumper from pin A of plug 1, you can test each cable conductor for grounds. To do this, connect one meter lead to ground and the other to each of the pins of one of the plugs. A low resistance shows 2-16

that a pin is shorting to ground. For this test you remove both plugs from their units; if you remove only one plug, a false indication is possible. Short Test A short circuit occurs when two conductors accidentally touch each other directly or through another conducting element. In a short circuit, enough current may flow to blow a fuse or open a circuit breaker. However, it is possible to have a short between two cables carrying signals and the short not blow the fuse. You use the resistance setting to check for a short. By measuring the resistance between two conductors, you may detect a short between them by a low resistance reading. Refer to Figure 2-2. If you remove the jumper and disconnect both plugs, you can make a short test. You can do this by measuring the resistance between the two suspected conductors. Shorts not only happen in cables, but also in many components, such as transformers, motor windings, capacitors, etc. The major method for testing such components is to make a resistance measurement and compare the indicated resistance with the resistance given on schematics or in maintenance manuals. Voltage Tests Voltage checks are made while the power is applied; therefore, you must follow the prescribed safety precautions to prevent injury to personnel and equipment damage. Voltage tests are important. You will use them to isolate malfunctions to major components and to maintain subassemblies, units, and circuits. For the applicable ac or dc voltages, the digital fluke meter is used to test voltages. When you use the digital fluke meter, make sure the meter selection is the correct one for the type of current (ac or dc) under test. Also, make sure it has a scale with a suitable range. Since defective parts in a circuit may cause higher than normal voltages to be present at the point of test, use the highest digital fluke meter range available first. Once you get a reading, you can determine if a lower scale is possible. If so, you use the lower scale as it provides a more accurate reading. Another consideration in the circuit voltage test is the resistance and current in the circuit. A low resistance in a high-current circuit may cause a considerable voltage drop. The same resistance in a low-current circuit may cause a minimal voltage drop. You can check abnormal resistance in part of a circuit with a digital fluke meter. Where practical, use resistance readings because the test is then on a dead circuit. Normally, you will work on low-current electronic circuits. Schematics show the voltages at various test points. If you suspect that a certain stage is defective, you may check the voltage by connecting a digital fluke meter from the test point to ground. If the suspected stage is good, the digital fluke meter readings will match the voltages given on the schematic. Some technical manuals contain voltage charts. These charts usually show the sensitivity of the meter (e.g., 20,000 ohms/volt) you should use to take the voltage readings for the chart. To get comparable results, you must use a digital fluke meter of the same sensitivity (or greater) as that specified. This is to be sure the digital fluke meter is not loading the circuit while taking a measurement. If the meter resistance is not considerably higher than circuit resistance, the reading will be markedly lower than true circuit voltage because of the digital fluke meter’s loading effect. 2-17

Ammeter The ammeter connects in series with the current path. Circuits requiring frequent current readings or adjustments provide current jacks for use with a plug-in meter. Some systems have a meter installed as part of the equipment. Ohmmeter The Navy does not have an instrument consisting solely of an ohmmeter. The ohmmeter and the voltmeter form a digital fluke multimeter. Therefore, you must determine the choice of ohmmeter by the resistance ranges available in the various digital fluke multimeters. Small multimeters have a high range of R x 100; larger m ultimeters, such as the AN/PSM-4, have a high range of R x 10,000. VTVM TS- 505A/D has a high range of R x 1,000,000. Wheatstone Bridge Resistance measurements taken using an ohmmeter are not always accurate enough. The cause of this inaccuracy is an error in meter movement and in the reading of the meter. The Wheatstone bridge (Figure 2-3) is widely used for precise resistance measurements. Resistors R1, R2, and R3 are precision, variable resistors. The value of Rx is an unknown value of resistance that you must determine. After properly balancing the bridge, you can find the unknown resistance by using a simple formula. The galvanometer (an instrument that measures small amounts of current) across terminals b and d shows the condition of balance. When the bridge is in balance, no difference in potential exists across terminals b and d; when switch S2 closes, the galvanometer reading is 0. When the battery switch (S1) closes, electrons flow from the negative terminal of the battery to point a. Here, the current divides as it would in any parallel circuit. Part of it passes through R1 and R2; the remainder passes through R3 and Rx. The two currents, I1 and I2, unite at point c and return to the positive terminal of the battery. The value of I1 depends on the resistance of R1 and R2, and the value of I2 depends on the resistances of R3 and RX. The current is inversely proportional to the resistance. Adjust R1, R2, and R3 so when S1 closes, and no current flows through G. When the galvanometer shows no deflection, there is no difference in potential between point’s b and d. All of follows the path, and all of follows the p ath. This means that voltage drop is the same as voltage drop . Similarly, the voltage drops across R2 and ( and ) are also equal. =

Figure 2-3 — Wheatstone bridge. 2-18

With this information, you can figure the value of unknown resistor . Divide the voltage drops across R1 and R3 by their respective voltage drops across R2 and as follows:

=

Simplify this equation:

=

Then, multiply both sides of the expression by to separate it:

For example, in Figure 2-3, you know that R1 is 60 ohms, R2 is 100 ohms, and R3 is 200 ohms. To find the value of , use the formula as follows:

The Wheatstone bridge is widely used to get precise resistance measurements. You can measure capacitance, inductance, and resistance for precise accuracy by using ac bridges. These bridges consist of capacitors, inductors, and resistors in a wide variety of combinations. RX = R2R3 R1 RX = R2R3 R1 RX = 100 x 200 60 RX = 20000 60 RX = 333.33 ohms 2-19

DC Voltmeter When selecting a dc voltmeter, you should consider the sensitivity of the meter movement and its effect on the circuit under test. NEETS, Module 16, Introduction to Test Equipment, NAVEDTRA 14188 discusses meter sensitivity. Use a multimeter having low sensitivity for quick, rough readings where approximations are adequate. When desiring a high degree of accuracy, use a meter having high sensitivity. Such a meter has wide application in the maintenance of medium- and high-impedance electronic circuits found in aviation electrical/instrument systems. A vacuum tube voltmeter, because of its high input impedance, is the ideal instrument for measuring low voltage in oscillators, automatic gain control, automatic frequency control, and other electronic circuits sensitive to loading. When measuring voltages of more than 500 volts, a multimeter having a sensitivity of 20,000 ohms per volt has an input impedance comparable to most vacuum tube voltmeters. The input impedance of most vacuum tube voltmeters is between 3 and 10 megohms. A 20,000-ohm-per-volt meter, when reading a voltage of 500 volts, has an input impedance of 500 x 20,000 or 10 megohms. Therefore, on the 500-volt scale, a multimeter of this sensitivity has an input impedance at least equal to the vacuum tube voltmeter. For voltage readings over 500 volts, a 20,000-ohms-per-volt multimeter offers input impedance higher than the vacuum tube voltmeter. You can use any multi-range voltmeter, though its sensitivity may not exceed 1,000 ohms per volt, to get fairly reliable readings in a dc circuit. If you do not know the impedance of the circuit under test, a comparison of two voltage readings will show if the meter is having a loading effect on the circuit. Take the voltage readings on the lowest usable range and the on the next higher range. If the two readings are approximately the same, the meter is not causing appreciable voltage variations, and you can accept the higher reading as the true voltage. If the two readings differ considerably, the true voltage may be found by the following formula:

E is the true voltage is the lower of the two voltage readings. is the higher of the two voltage readings. R is the ratio of the higher voltage range to the lower voltage range. As an example of how the formula works, make the following assumptions: 1. A reading of 22 volts was obtained between two terminals with the meter on the 0-30 volt scale. 2. A reading of 82 volts was obtained from the same terminals with the meter on the 0-300 volt scale.

E = E2 - E1 + E2 E1 R E2 2-20

The true voltage is found as follows:

Digital Multimeters The digital multimeter is an accurate, reliable measuring instrument used at the organizational, intermediate, and depot levels of maintenance. The instrument is easy to use and read. Some digital multimeters have a battery pack for the portable operation. The Fluke ® model 8808A digital multimeter (Figure 2-4, view A) is representative of a portable multimeter presently used in the fleet. It is a small, compact unit capable of measuring 200 mV to 750 Vac and dc voltages 200 mV to 1000 Vdc, resistance 200 ohms (Ω) to 100 MΩ, maximum current range up to 10A, continuity and diode test and frequency range 20hz to 1 Mhz. Standard features provided by the 8808A model include a 5

digit dual fluorescent display that allows two properties of an input signal to be displayed at the same time, measurement rates of 2.5, 20 and 100 samples/second, front-panel setup key for single key access to saved setups, a compare mode to determine if a measurement is within defined limits, and remote operation via the RS- 232 interface terminal. The 8808A digital multimeter uses three fuses, one fuse to protect the line-power input and two fuses to protect the current-measurement inputs. Accessories and options besides the premium test lead set are a RS-232 interface cable, a precision electronic probe set, a 2x4 wire ohms 1000 V test lead and switched ac/dc current shunts. During organizational level troubleshooting, you will often measure voltage, current, and resistance. Rather than using three or more separate meters for these measurements, you can use the Fluke ® digital multimeter 170 series (Figure 2-4, view B). The Fluke ® 170 series digital multimeter gives the operator true voltage and current measurements, 6000 count resolution, manual and automatic ranging and provides frequency capacitance, resistance, continuity and diode measurements. A digital fluke multimeter is often called a volt-ohm-milliammeter (VOM). One advantage of a VOM is that no external power source is necessary for its operation. Therefore, no warm-up is necessary. It is a portable, versatile meter that provides one-handed operation. The VOM does have two disadvantages: 1. The VOM can load the circuit under test. 2. The meter movement is easy to damage because of improper testing procedures. Safety is a vital concern when working around or near electrical circuits and cannot be overemphasized. To avoid possible electric shock or personal injury while using the VOM, follow these simple guidelines: E = 82 – 22 + 82 22 x 10 - 1 82 E = 117.7 volts 2-21

Figure 2-4 — Digital multimeters. 1. Do not use multimeters or test leads that appear damaged or not operating properly. 2. Check the test leads for continuity prior to use. 3. Always use proper terminal, switch position and range for measurements. 4. Verify the multimeter’s operation by measuring a known voltage. 5. Do not apply more than the rated voltage than the meter is capable of reading. 6. Disconnect circuit power supply and discharge all high-voltage before testing resistance, continuity, diodes, or capacitance. 7. Avoid working alone.

Megger Ordinary digital fluke multimeters cannot be used to measure multimillion ohm values of resistances, such as those in conductor insulation. To test for insulation breakdown, you use a much higher potential than that supplied by the battery of a digital fluke meter. This potential is placed between the conductor and the outside of the insulation. You will use a megger (megohmmeter) for these tests. 2-22

Figure 2-5 — Megger internal circuit and external view.

The megger (Figure 2-5) is a portable instrument consisting of two main elements. 1. The hand-driven dc generator, which supplies the voltage for making the measurement, and 2. The instrument portion, which shows the value of the resistance you are measuring. The instrument portion is of the opposed-coil type, as shown in Figure 2-5, view A. Coils a and b are on movable member c. A fixed angular relationship exists between coils; they are free to turn as a unit in a magnetic field. Coil b moves the pointer counterclockwise, and coil a moves it clockwise. Coil a connects in series with R3 and unknown resistance RX. The combination of coil a, R3, and RX forms a direct series path between the + and – brushes of the dc 2-23

generator. Coil b connects in series with R2. This combination also connects across the generator. Notice that the movable member (pointer) of the instrument portion of the megger has no restoring springs. Therefore, when the generator is not operating, the pointer will float freely and may come to rest at any position on the scale. The guard ring, shown in Figure 2-5, view A, shunts any leakage currents to the negative side of the generator. This prevents such current from flowing through coil a and affecting the meter reading. If the test leads are open, no current will flow in coil a. However, current will flow internally through coil b and deflect the pointer to infinity. This reading shows a resistance too large to measure. When you connect a resistance, such as RX, between the test leads, current also flows in coil a; this moves the pointer clockwise. At the same time, coil b moves the pointer counterclockwise. Therefore, the moving element, composed of both coils and the pointer, comes to rest at a position in which the two forces balance. This position depends upon the value of RX, which controls the current in coil a. Because changes in voltage affect both coils in the same proportion, the position of the moving element is independent of the voltage. If you short the test leads together, the pointer will come to rest at zero because the current in coil a is relatively large. Since R3 limits the current, the instrument is not damaged under these circumstances. Figure 2-5, view B, shows the external appearance of one type of megger. Most meggers you will use have a 500-volt rating. (Normally, meggers have friction clutches.) When you crank the generator faster than its rated speed, the clutch slips. This prevents the generator speed and output voltage from exceeding rated values. A 1,000-volt generator is available for extended ranges. When you want to measure an extremely high resistance (10,000 megohms or more), a high voltage is needed to cause enough current flow to operate the meter movement. When you use a megger, observe the following precautions:  When making a megger test, make sure the equipment is de-energized.  Observe all rules for safety in preparing equipment for test and in testing, especially when testing installed high-voltage apparatus.  Use well-insulated test leads, especially when using high-range meggers. After connecting the leads to the instrument and before connecting them to the component under test, operate the megger to make sure there is no leakage between leads. The reading should be infinity. Make certain the leads are not broken and the connections are good by touching the leads together while turning the crank slowly. The reading should be zero.  When using high-voltage meggers, take proper precautions against electric shock. There is enough capacitance in most electrical equipment to store up sufficient energy from the megger to give a very disagreeable and even dangerous electric shock. The megger has a high protective resistance, and its open circuit voltage is not as dangerous, but always be careful.  Discharge equipment having a considerable capacitance before and after making megger tests to avoid the danger of receiving a shock. You can do this by grounding or short circuiting the terminals of the equipment under test. 2-24

Oscilloscope An oscilloscope is a versatile instrument used in viewing wave shapes of voltages or currents. It is capable of giving information concerning frequency values, phase differences, and voltage amplitudes. It can also trace signals through electronic circuits to localize sources of distortion and to isolate troubles to particular stages. The majority of oscilloscopes used for test and other measurements contain a basic presentation unit (screen) and accessory units, such as amplifiers, synchronizing circuits, time-delay circuits, and sweep oscillators. These circuits are used to display a stationary pattern on the Cathode-Ray Tube (CRT). The sensitivity of an oscilloscope should be adequate for the smallest signals to have sufficient amplitude for screen display. NEETS discusses the CRT, the oscilloscope, and their operation in detail. Time-Domain Reflectometry Time-Domain Reflectometry (TDR) is a measurement concept widely used in the analysis of wideband systems. The art of determining the characteristics of electrical lines by observation of reflected waveforms is not new. For many years power- transmission engineers have located discontinuities in power-transmission systems by sending out a pulse and monitoring the reflections. Discontinuity is any abnormal resistance or impedance that interferes with normal signal flow. TDR is particularly useful in analyzing coaxial cables, such as those in fuel or oxygen quantity indicating systems. The amplitude of the reflected signal corresponds directly to the impedance of the discontinuity. You can find the distance to the discontinuity by measuring the time required for the pulse to travel down the line to the reflecting impedance and back to the monitoring oscilloscope. TDR analysis consists of inserting an energy step or pulse into a system and the subsequent observation, at the insertion point, of the energy reflected by the system. Several arrangements are possible, but the following procedure is used with the newer, specialized reflectometers (Figure 2-6). The pulse generator develops a fast (or incident) step. This step then passes through a TEE connector and goes into the sys tem under test. The sampling oscilloscope is also attached to the TEE connector, and the incident step, along with the reflected waveform, shows on the CRT. Analysis of the magnitude, duration, and shape of the reflected waveform shows the type of impedance variation in the system under test. When the fast-rise input pulse meets with a discontinuity or an impedance mismatch, the resultant reflections appearing at the feed point are compared in phase, time, and Figure 2-6 — Typical time domain reflectometer. 2-25

amplitude with the original pulse. Also, since distance relates to time and the amplitude of the reflected step directly relates to impedance, the comparison shows the distance to the fault, as well as an indication of its nature. Time-domain reflectometry shows both the position and the nature (resistive, inductive, or capacitive) of each line discontinuity. It also reveals the characteristic impedance of the line and shows whether losses are parallel or series losses. Phase Angle Voltmeter You can determine the overall accuracy of many electronic parts by measuring phase angles in computing transformers, computing amplifiers, and resolver systems. In the past, the most common method used for measuring phase shift or phase angles between signals was observing patterns on an oscilloscope. With this method, it was hard to determine small angles and difficult to translate various points into angles and sines of angles. When one of the signals contained harmonic distortion or noise, this interference limited the use of oscilloscope patterns. In any complex waveform containing a basic frequency and harmonics, measuring phase shifts presents problems. In most applications, interest lies in the phase relationship of the basic frequencies, regardless of any harmonics that are present. One requirement of a phase measuring device is measuring the phase difference between two discrete frequencies. It must accomplish this, regardless of phase and amplitude of other components of the waveform. The basic block diagram of a phase angle voltmeter is shown in Figure 2-7. You should refer to it while reading this section. There are two inputs, the signal and the reference. Both channels contain filters that pass only the fundamental frequency. Harmonics are highly attenuated. Each channel has a variable amplitude control and amplifiers to increase the variety of signals you can check. By placing a calibrated phase shifter into the reference channel, that channel signal can be phase shifted to correspond to the other channel. The phase detector will detect this action and it will also show on the meter. The calibrated phase shifter connects to a switch (whose position corresponds to the 0- degree, 90-degree, 180-degree, and 270-degree phase shift) and a potentiometer (whose dial is calibrated from 0 degree to 90 degrees). The total phase shift is the sum of the two readings. Figure 2-7 — Phase angle voltmeter, block diagram. 2-26

The phase detector is a balanced diode bridge-type demodulator. Its output is proportional to the signal amplitude times the cosine of the angle of phase difference between the signal input and the reference input. If the reference input is phase shifted until it is in phase or 180 degrees out of phase with the signal input, the output from the phase detector is proportional to the signal input amplitude (the cosine of the angle is unity). If the reference input is phase shifted until it is 90 degrees out of phase with the single input, the phase detector output is zero (the cosine of the angle is zero). The point where the two signals are in phase or 180 degrees out of phase is the point of maximum deflection on the meter. The difference between the in-phase and 180 degrees out-of-phase points is in the direction of needle swing, not the distance it swings. Upon approaching the point of maximum deflection, the rate of change of the meter reading decreases. This happens because the cosine has a smaller rate of change near 0 degree. This makes it difficult to read the point of maximum deflection. Most commercial voltmeters determine the point at which the signals are 90 degrees out of phase as quadrature. The cosine has a maximum rate of change as it approaches 90 degrees, thus it gives a better indication on the meter. When the voltmeter is set up for this point, there must be some way of converting the phase shifter reading. You have to convert the reading to show the correct amount of phase shift, rather than 90 degrees more or less than the actual amount. Some confusion exists in this area because different manufacturers have different methods of determining the signal quadrant. Manufacturers also differ on whether the final reading is a leading or a lagging phase shift. This means that you should be familiar with the type of phase angle voltmeter you are using. You can’t assume that the method used to determine phase angle on one type of meter can be used on another. AIRCRAFT WIRE AND CABLES An important part of aircraft electrical maintenance is determining the correct wire or cable (Figure 2-8) for a given job. For electrical installations, a wire is a stranded conductor, covered with an insulating material. The term cable, as used in aircraft electrical installation, includes the following:  Two or more insulated conductors contained in the same jacket (multiconductor cable)  Two or more insulated conductors twisted together  One or more insulated conductors covered with a metallic braided shield (shielded cable)  A single insulated center conductor with a metallic braided outer conductor (Radio Frequency (RF) cable)

2-27

Figure 2-8 — Cables commonly used in aircraft.

2-28

Wire Replacement Before you can replace wiring, you need to find information about the wire. First, consult the aircraft’s Maintenance Instructions Manual (MIM) for wire replacement since it normally lists the wire used in a given aircraft. When you cannot get this information from the manual, you must select the correct size and type of wire needed. The factors in determining correct wire size are in NEETS, Module 4, Introduction to Electrical Conductors, Wiring Techniques, and Installation and Repair Practices, NAVAIR 01-1A- 505 (series). The procedures specified in NAVAIR 01-1A-505 (series) are mandatory for the maintenance of naval aircraft. The information in these publications and in the latest Military Specification MIL-W-5088 should help you select the correct replacement wire. The data needed to determine correct wire for a given application is summarized below: 1. Current drawn by the load (Table 2-1) 2. Length of wire required to go from the source to the load 3. Allowable voltage drop between the source/point of voltage regulation and load 4. The maximum voltage applied to the wire 5. The approximate ambient air temperature of the wire installation location 6. Whether the conductor is a stranded wire in free air or one of a group of wires in a bundle or in a conduit

2-29

Table 2-1 — Current Rating of Wires CONDUCTOR MATERIAL WIRE SIZE Continuous duty current (amperes) (2) Wires in bundles, groups, or harnesses WIRE TEMPERATURE RATING 221 °F (105 °C) 302 °F (150 °C) 302 °F (200 °C) COPPER OR COPPER ALLOY 26(1) 2 3 4 24 3 4 5 22 4 6 7 20 5 8 10 18 7 11 14 16 8 12 16 14 11 17 22 12 15 22 29 10 19 30 38 8 26 39 50 6 35 53 68 4 48 72 93 2 66 100 130 1 78 115 148 1/0 89 135 173 2/0 106 159 203 3/0 125 186 241 4/0 146 220 287 ALUMINUM (2) 8 17 NOTE (1) The use of these wires requires procuring activity approval. (2) Rating is for 94 °F (70 °C) ambient, 33 or more wires in the harness with no more than 20% of harness current capacity being used, at an operating altitude of 60,000 feet. (3) Ratings are for copper conducting size 4/0 through size 22 & copper alloy for size 24 through 26. 6 23 4 30 2 43 1 50 1/0 57 2/0 68 3/0 76 4/0 86 2-30

Wire Selection and Characteristics After you have determined the wire size, consider insulation characteristics. The following are descriptions of the various military specifications of wires and cables. These military specification numbers are on the reel, spool, or shipping container. Always refer to NAVAIR 01-1A-505 (series) before selecting wires and cables for weight discrepancies. See below for various wires and cable descriptions:  MIL-W-22759. MIL-W-22759 is a fluoropolymer-insulated single conductor electric wire. It is made of tin-coated, silver-coated, or nickel-coated wires of copper or copper alloy. You may use this specification wire in combination with other insulating materials. Temperature and voltage ratings range from 302°F to 500°F and 600 to 1,000 volts, respectively, depending on part number and application.  MIL-W-25038. MIL-W-25038 is a high temperature, fire-resistant electrical wire. This wire is made of nickel-clad copper with insulation that will operate efficiently in ambient temperatures that exceed +500°F. Its insulation design assures emergency operation of electrical circuits subject to fires.  MIL-W-81044. MIL-W-8 1044 is a single conductor with tin, silver, and nickel- coated copper or copper alloy wires insulated with various poly materials. You may use these poly materials alone or in combinations, depending upon application and part number.  MIL-W-81381. MIL-W-81381 is a polyamide-insulated, single wire with silver and nickel-coated conductors made of copper and copper alloys. You may use polyamide insulation alone or in combinations with other insulating materials, depending on application and part number.  MIL-C-7078. MIL-C-7078 includes three types of electrical cable: 1. Unshield ed/unjacketed/twisted—two or more coded wires with no overall jacket or shield 2. Jacketed/twisted —two or more coded wires with no overall shield enclosed within a single jacket 3. Shielded/jacketed/twisted —one or more coded wires, twisted, with an overall shield enclosed within a single jacket  MIL-C-27500. MIL-C-27500 is electrical cable made up of two to seven-wire specifications. Cables of this type are included in the following groups: 1. Unshielded/unjacketed —color-coded wires, spirally laid without an overall jacket 2. Jacketed —color-coded wires, spirally laid with an overall jacket 3. Shielded/jacketed —one to seven color-coded wires, spirally laid with one or two shields within an overall jacket Aluminum Wire Properties The use of aluminum wire in aircraft is well established. Therefore, you need to know some of the unusual properties of aluminum. 2-31

CAUTION Never cut aluminum wire with tools that have reciprocating motion, such as a hacksaw. Reciprocating cutting action hardens aluminum. This will lead to broken/torn strands. Aluminum is softer than copper. It has an inherent property known as creep, which makes proper installation of a terminal extremely critical. Creep is the tendency of aluminum to flow away from the point of pressure. Aluminum is unusual because it forms an electrically resistant oxide film on all of its surfaces. The electrically resistant aluminum oxide film is always present. Therefore, you must either penetrate or remove it to guarantee a satisfactory electrical connection by using the specified compound to deal with this film. Refer to Cleaning and Corrosion Control, NAVAIR 01-1A-509 (series). Initially, tin-plated terminals and splices are supplied so no oxide film is present. Terminals become dirty when stored or excessively handled. You should clean them by wiping with a soft cloth. Never use a wire brush or any abrasive method to clean a tinned aluminum surface. One problem found in aluminum wire is corrosion by dissimilar metals. This problem occurs frequently when aluminum and copper come in contact. As soon as moisture collects between the two metals, an electrical differential is created, which produces a battery effect. This effect quickly results in corrosion. Use the following techniques to reduce the undesirable effects caused by aluminum’s softness, creep, and oxide film:  Select the proper size of wire and terminals/splices. Terminals and splices used with aluminum wire will have an AL or ALUM stamp.  Select the proper hand tool when crimping.  Handle aluminum wire carefully. Use the proper assembly procedure when preparing it for an electrical connection.  Thoroughly clean any terminals.  Be careful not to scrape or nick the wire when stripping.

Wire Identification Equipment contractors assign the wire identification code for wires and cables having function letters R, S, T, and Y. The block of wire numbers for each type of equipment starts with 1 and continues for as many numbers needed to identify all wires. For example, wires of the AN/APS-45 are identified as APS45-1A20-APS45-975C22, those of the AN/ARC-52A would be ARC52-1A22- ARC52-9C22, and the MX-94 would be the MX94-1A20, etc. If a type designation (AN nomenclature) is not assigned for a piece of equipment (such as commercial equipment), get a block of numbers from the procuring activity. You can get the wire identification code by using that portion of the military type designation (AN nomenclature) following the slash (/). You must exclude the hyphen and any suffix letters. Wire Marking You may stamp the identification code on wires either horizontally or vertically, as shown in Figure 2-9. The preferred method is to stamp the identification marking directly 2-32

Figure 2-9 — Examples of wire identification coding.

on the wire or cable with a hot foil stamping machine. Use this method wherever possible. If the wire insulation or outer covering won’t stamp easily, stamp lengths of insulating tubing (sleeves) with the identification marking. Then, install the sleeve on the wire or cable. The following types of wire usually have sleeve identification markings: 1. Unjacketed shielded wire 2. Thermocouple wires 3. Multi-conductor cable 4. High-temperature wire with insulation difficult to mark, such as Tetrafluoroethylene (TFE), fiber glass, etc.

Whatever method you use to mark the wire, make sure the marking is legible and the color contrasts with wire insulation or sleeving. Use black stamping for light-colored backgrounds and white on dark-colored backgrounds. Make sure that markings are dry so they don’t smear. Stamp wires and cables at intervals of not more than 15 inches along their entire lengths (Figure 2-10). Also, stamp wires within 3 inches of each junction (except permanent splices) and at each ending point. Stamp wires which are 3 to 7 inches long in the center. You don’t need to stamp wires less than 3 inches long.

CAUTION Do not use metallic markers or bands for identification. Do not use any method of marking that will damage or deform the wire or cable. 2-33

NOTE Stranded conductor wire is used for flexibility in installation and service. Wire sizes approximate American Wire Gage (AWG). However, they vary sufficiently that it is improper to refer to aircraft wire as AWG. For an in-depth study of wiring specifications, limitations, and repair, you should refer to NAVAIR 01-1A-505 (series). Figure 2-10 — Spacing of identification stamping on wire and cable.

Wiring and Cable Identification Codes To make maintenance easier, each of the connecting wires in an aircraft have identification. The identification is a combination of letters and numbers that identifies the circuit it belongs to, gauge size, and other information relating the wire to a wiring diagram. This marking is the cable identification code. The accompanying discussion explains the code used in aircraft wiring. You can find complete details in the Military Specification: Wiring, Aerospace Vehicle, MIL-W-5088.

The basic wire identification code for circuits is read from left to right, except circuit function letters R, S, T, or Y. Look at Figure 2-12 as you read this section. Unit number – Where two or more identical items of equipment are in the same aircraft, use prefix unit numbers 1, 2, 3, 4, etc., to differentiate between wires and cables. To make it easier to interchange items, identical wiring is located in left and right wings, nacelles, and major interchangeable structural assemblies, without using the unit number. Use the unit number for circuit functions R, S, T, and Y, only where complete 2-34

duplicate equipment is installed. Unit numbers don’t apply to duplicate components within a single complete equipment, such as duplicate indicators or control boxes. Circuit function letter (except R, S, T, and Y) – The circuit function letter identifies the circuit function (Table 2-2). When using a wire or cable for more than one circuit function, the circuit function letter of the predominant circuit applies. When functional predominance is questionable, use the circuit function letter for the wire or cable having the lowest wire number. Wire number – The wire number consists of one or more digits. It identifies the different wires in a circuit. A different number is used for wire not having a common terminal or connection, as shown below: 1. Wires with the same circuit function having a common terminal connection or junction have the same wire number but different segment letters. 2. Beginning with the lowest number, assign a number to each wire in numerical sequence, insofar as practical. Wire segment letter – A wire segment is a conductor between two termi nals or connections. The wire segment letter identifies different conductor segments in a particular circuit. Use a different letter for wire segments having a common terminal or connection. Wire segment letters are in alphabetical sequence. The letter A identifies the first segment of each circuit starting at the power source. If a circuit contains only one wire segment, this wire segment is A. Do not use the letters I and O as segment letters. Use double letters AA, AB, AC, etc., when there are more than 24 segments. Two permanently spliced wires do not require separate segment letters if the splice is for modification or repair. Table 2-2 — Wiring Circuit Function Code Circuit Function Letter Circuits A Armament B Photographic C Control surface D Instrument (other than instrument & flight) E Engine instrument F Flight instrument G Landing gear, wing folding H Heating, ventilating, & de-icing J Ignition K Engine control L Lighting M Miscellaneous (electrical) P Dc power (wiring in the dc power / power control system is identified by the circuit function letter “P”) Q Fuel & oil R Radio (pulse technique) S Radar (navigation & communication) T Special electronic U Miscellaneous (electronic) V Dc power cables & dc control cables for ac systems are identified by the circuit function letter “V”. W Warning & emergency X Ac power (wiring in the ac power system is identified by the circuit function letter “X”). Y Armament special systems Z Experimental circuits 2-35

Figure 2-11 — Wire stripping method. Wire size number – The wire size number identifies the size of the wire or cable. For coaxial cables and thermocouple wires, the wire size number is not included. For thermocouple wires, a dash (-) is used instead of the wire size number. Ground, phase, or thermocouple letter(s): 1. Ground cable letter N is a suffix to the wire identification code. It identifies any wire or cable that completes the circuit to the ground network. Such wires and cables connect to the ground network of aircraft electrical systems without causing any circuit malfunctions. For electronic systems with interconnecting ground leads, but only one segment actually grounded to structure, N identifies the segment actually grounded to the structure. 2. Phase letter A, B, or C is a suffix on the wire identification code. It identifies the phase or wires in the three-phase power ac distribution systems. The phase sequence is A→B→C. 3. Phase letter V is a suffix on the cable identification code. It identifies the ungrounded wire or cable in a single-phase system. 4. For thermocouple wire, use the following suffixes:  CHROM —Chromel  ALML —Alumel  IRON —Iron  CONS —Constantan  COP —Copper Suffix (when required) – When using aluminum wire, add ALUMINUM OR ALUM to the identification code. Cable Stripping and Heat Shrinkable Tubing Nearly all wire and cable electrical conductors have some type of insulation. When making electrical connections with wire, you must remove a part of this insulation, leaving the end of the wire bare. To help you remove insulation, use a wire and cable stripping tool similar to the one shown in Figure 2-11.

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The operation of this basic tool is efficient, effective and extremely simple. To operate it, insert the wire end in the proper direction to the depth to be stripped. Now position the wire so it rests in the proper groove for that size wire, and squeeze. Heat-shrinkable tu bing is a plastic-like tubing (similar to insulation sleeving) that will shrink to a smaller diameter with proper heating. Place the tubing over the joint, terminal, or part needing insulation. Now apply heat with a heat gun, oven, or other appropriate heat source. When the tubing reaches a specific temperature (shrink temperature depends upon the type of tubing), it quickly shrinks around the object, forming a snug jacket. In addition to being an insulator, the shrinkable tubing helps relieve strain and adds waterproofing. Figure 2-12 gives some of the typical uses of heat-shrinkable tubing. Soldering Wires are soldered to form a continuous and permanent metallic connection, having a constant electrical value. When soldering, strive for superior workmanship. If you are sloppy, you create problems and compound difficulties in system troubleshooting techniques. Study and refer to NAVAIR 01-1A-505 (series) when you are soldering. The most important part of soldering is the selection of the correct iron for the job. Soldering irons are available in wattage ranges from 20 to 500 watts. Use irons with wattage ratings of 60, 100, and 200 watts for general work in aircraft electrical wiring. Use pencil irons with a rating of 20 to 60 watts for soldering small parts. The soldering iron for printed circuit soldering is a lightweight, 55-watt iron with a 600°F (316°C) Curie point tip control. This iron has a three-wire cord, which prevents leakage currents that could damage the printed circuits. 2-37

Figure 2-12 — Typical heat shrinkable tubing. 2-38

When soldering, you should select a soldering iron with a thermal capacity high enough so the heat transfer is fast and effective. Refer to Table 2-3. An iron with excessive heat capacity will burn or melt wire insulation. One with too little heat capacity will make a cold joint in which the solder does not alloy with the work. A soldering iron should also suit the production rate. Do not select a small pencil iron where you need a high steady heat flow. AIRCRAFT ELECTRICAL HARDWARE You shouldn’t always reuse the same mounting parts that you removed from the installation. Before reusing the parts, inspect them to make sure they aren’t defective or damaged. Also check the instructions; some parts can't be reused. Then, and only then, may you use the removed parts. The hardware you should use when installing electrical equipment in aircraft is specified in the applicable MIM. You should always use proper parts. If you need to substitute, make sure the substitute item is satisfactory. General information about mounting parts and consumable components can be found in various Navy training manuals. Aircraft Structural Hardware for Aircraft Repair, NAVAIR 01-1A-8, and Installation and Repair Practices, NAVAIR 01-1A-505 (series), are sources of detailed information. If you can’t get the mounting parts specified by the applicable Illustrated Parts Breakdown (IPB), you may make a temporary installation using suitable substitute parts. Replace these parts with the proper items as soon as you receive them. Always check with the work center supervisor before you make any substitution. When making part substitutions, you should give special consideration to the following factors:  Corrosion – Pay attention to the chemical or metallic composition of the part. Choose a part that doesn’t contribute appreciably to the danger of corrosion.  Strength – The strength of the substitute part must be the same, or greater than the one prescribed. When determining the strength, consider the tensile, compression, and shear strength, as applicable to the specific use.  Size – Substitute nuts, bolts, and screws should be the same size as the prescribed item. In all cases, washers must have the same inner diameter as the prescribed item. A different outer diameter or thickness is acceptable.  Length – The length of substitute screws or bolts must be enough for the particular installation. However, length can’t be long enough to interfere with any moving part. Hardware shouldn’t come in contact with other aircraft items, such as electrical wiring, hydraulic lines, etc.  Magnetic properties – Equipment installed in specific areas of the aircraft shouldn’t cause distortion of the magnetic fields of the area. Examples of such items include the magnetic compass, magnetic anomaly detection equipment, Table 2-3 — Approximate Soldering Iron Size for Tinning Wire Size (AN Gage) Soldering Iron Size (Heat Capacity) #20 - #16 #14 - #12 #10 & #18 65 Watts 100 Watts 200 Watts 2-39

radio direction finder, or gyros. In areas containing these types of equipment, any substitute part must have the same magnetic properties and characteristics as the one prescribed.  Style – Most items of mounting hardware are available in various styles. It is usually easy to find screws and bolts that are the same in all respects except the head type. Use these parts as substitutes, provided they have all required special features.  Special features – If a bolt requires torque to a given value, a suitable torque wrench for that type part must be available. If the MIM calls for lockwire, the part must have suitable provisions.  Lubrication or coating – If specific instructions call for lubrication or coating of the parts, follow those instructions for the substitute part as well as for the prescribed part. Shock Mounts Electrical and electronic equipment must be protected from the effects of vibration in aircraft, as vibration is a major problem. Most amplifiers, instrument panels, and other fragile parts have shock mounting protection. Shock mounts are also known as vibration insulators. You can trace the failure of many systems to faulty shock mounts; therefore, you should periodically check shock mounts. If you find that they are defective, replace them before equipment is damaged. Figure 2-13, views A and B, shows two types of shock mounts used in naval aircraft. View A shows mounts that are individually replaceable. Each mount has a rod that extends into the vibration eliminating material. You can replace this type of mount by drilling out the mounting base rivets and riveting the replacement in position. The replacement must be of the same size and type as the mount that it is replacing. The weight of the unit being protected determines the type of mount you use. If you use a mount designed for a heavier unit, it won’t give to protect the unit. If you use a mount designed for a lighter unit, it can easily pull away from the base and damage the unit. The isolation washers in Figure 2-13, view B, are used with a particular piece of equipment. You replace these types of washers individually. The vibration isolator washers are made of hollow rubber and locked into place between the top and bottom of the isolator when installed. Inspect these isolators for cracks or splits. If you find damage, replace the complete set all at once. Shock absorbing materials commonly used in shock mounts are usually electrical insulators. These types of shock mounts are mounted so they have an electrical bond to the aircraft structure, preventing static electricity buildup. See Figure 2-13, view A. You need to inspect the bonding strap when inspecting the shock mounts. Replace or repair defective or ineffective bonding straps.

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Figure 2-13 — Typical shock mounts.

Electrical Connectors In this section, the word connector is used in a general sense. It applies to connectors with AN numbers and those with MS numbers. AN numbers were formerly used for all supply items cataloged jointly by the Army and Navy. Many items, especially those of older design, continue to carry the AN designator, although the supply system is shifting over to Military Specification (MS) numbers. Electrical connectors provide detachable coupling between major components of electrical and electronic equipment. These connectors are built to withstand the extreme operating conditions imposed by airborne service. They must make and hold electrical contact without excessive voltage drop despite extreme vibration, rapid shifts in temperature, and changes in altitude. Connectors consist of two portions, the fixed portion, called the receptacle, and the movable portion, called the plug. Plug assemblies may be straight or angled (usually 90 degrees). Receptacle assemblies may be of the wall-mounted, box-mounted, or integral-mounted types. MS numbers and letters identify the type, style, and arrangement of a connector. Connectors vary widely in design and application. A coupling nut or ring holds the two assemblies firmly together. The assembly consists of an aluminum shell containing an insulating insert, which holds the current-carrying contacts. The plug usually attaches to the cable end and is the part of the connector on which the coupling nut mounts. The 2-41

receptacle is the half of the connector to which the plug connects. The receptacle is usually mounted on a part of the equipment. In naval aircraft, connectors with crimp-type contacts are widely used. Maintenance is easier because you can remove the contact from the connector. If the connector is damaged, you can remove the contacts and replace the connector shell. If just a connector pin is damaged, you can remove and replace the pin. This is a considerable advantage over the solder-type connector, both in convenience and time savings. A discussion of the special tools you need to remove and insert crimped contacts is contained in Installation and Repair Practices, NAVAIR 01-1A-505 (series). Some common types of subminiature connectors are shown in Figure 2-14. They are used on instruments, switches, transformers, amplifiers, relays, etc. Fabrication of Cables Occasionally, you will have to make a cable using connectors. The type of connector you will use is specified in the MIM for the particular aircraft. The following steps outline the procedure you should use to make a cable: 1. Disassemble the connector to allow access to the terminals. Devise a way to hold the connector so both hands are free. 2. Cut the cables to the correct length. 3. Strip the wire end with a wire stripper; every effort must be taken to avoid cutting or nicking the wire strands. Tin the bare wire end. 4. Run the wires through the connector assembly and coupling nuts. 5. Make sure all surfaces are clean. 6. Flow rosin-core solder into the connector terminals. 7. Hold the tip of the soldering iron against the terminal. As the solder melts, push the wire into the cavity. Hold the wire steady while the solder cools. When you solder, be careful not to injure the connector insulation with the soldering iron. Follow a prearranged sequence (Figure 2-15). The recommended sequence is to start from the bottom connection and work left to right, moving up a row at a time. After soldering the connections, the shields, if used, are soldered to a common terminal or ferrule. Then lace the cable and reassemble and moisture-proof the connector, if necessary. Fabricating instructions are contained in NAVAIR 01-1A-505 (series). Figure 2-14 — Subminiature connectors.

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Figure 2-15 — Connector soldering sequence.

Figure 2-16 — Installation of cable terminals on terminal block.

Moisture proofing Present Navy practice is to use potted connectors (moisture-proof or environment-proof connectors). All jet and carrier-type aircraft have potted connectors. On other aircraft, use moisture-proofing sealant on electrical connectors in areas where a chance of failure exists. All connectors in wheel wells, wing fold areas, engine areas, engine nacelles, or cockpit decks have a high chance of failure and are sealed. In addition, moisture-proof all connectors that interconnect flight/basic navigation equipment. Vibration and lateral pressure fatigue wires at the solder cup. Moisture-proofing reduces electrical connector failures by reinforcing the wires against vibration and lateral pressure. The sealing compound also protects electrical connectors from corrosion and contamination by excluding metallic particles, water moisture, and aircraft liquids. One result of the connector’s better dielectric characteristics is the reduced chance of arc-over between pins. Terminal Blocks Terminal blocks, made from an insulating material, support and insulate a series of terminals from each other, as well as from ground. They provide a way to install terminals within junction boxes and distribution panels. The two methods of attaching cable terminals to terminal blocks are shown in Figure 2-16. View A uses a standard non- locking nut. In this installation method, the use of a lock washer is necessary. Figure 2- 16, view B shows the preferred method. 2-43

Figure 2-17 — Aircraft junction box. When using an anchor nut, or self-locking nut, you omit the lock washer. The use of anchor nuts is especially desirable in areas of high vibration. In both installations, a requirement exists for the use of a flat washer, as shown in the drawing. Each terminal board in the aircraft electrical system is identified by the letters TB followed by the number of the individual board. Each stud on the terminal board is identified by a number. The lowest number in the series starts at the end nearest the terminal board identification number. The identification number is on the structure to which the terminal board attaches. It mounts on any identification strip cemented to the structure under the terminal board. When replacing a terminal board, don’t remove the identification marking. If the identification marking is damaged, replace it with one that is the same as the original. Junction Boxes Junction boxes accommodate electrical terminals or other equipment. Individual junction boxes are named according to their function, location, or the equipment with which they are associated. Junction boxes, except boxes labeled “vapor tight” , have a drain hole, at the lowest point to drain water, oil, condensate, or other liquids. Figure 2-17 shows a representative junction box for housing and securing a multitude of relays.

When you install a junction box, make sure the screw or bolt heads are inside the box. The junction box has fasteners or knobs to secure the assembly; make sure they are tight and the assembly attaches to the radio frequency ground contact strips, if provided. Don’t install attaching hardware so the threaded part of the screw or bolt protrudes inside the junction box. The sharp thread edges of protruding hardware may damage wire insulation.

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Figure 2-18 — Cable clamps. Figure 2-19 — Routing cables through lightening holes Support Clamps Clamps provide support for conduit and open wiring and serve as lacing on open wiring. Clamps usually have a rubber cushion, or they are of all-plastic construction. When used with shielded conduit, the clamps are of the bonded type (Figure 2-18, view A); that is, there is a provision for electrical contact between the clamp and conduit. Use unbent clips for the support of open wiring. A strap-type clamp (Figure 2-18, view B) or an AN 742 (Figure 2-18, view C) is used for long cable runs between panels. The preferred method for supporting cable runs of all types is using AN 742 clamps. MS 25281D plastic clamps are for use where the maximum temperature does not exceed 250°F. When using the strap-type clamp, you must make sure the clamps hold the cable firmly away from lines, surface control cables, pulleys, and all movable parts of the aircraft. Use these clamps only as a temporary measure. Replace with a permanent installation as soon as possible. When cables pass through lightening holes, the installation should conform to the examples shown in Figure 2-19. Lightening holes are openings cut into a strengthening member that decreases its weight without significantly reducing its strength. These are useful in aircraft application when routing wire and cable bundles. In each case, the MS21919 cable clamp holds the cable firmly. Route the cable well in the clear of the edges of the lightening hole to avoid any chance of chafing the insulation. If any wire is closer than 1/4 inch to the edge of the lightening hole, use a grommet (a rubber cushion) to protect the wires. Protect wire bundles from the following:  High temperature  Battery acid fumes, spray, or spillage  Solvents or fluids  Abrasion in wheel wells where exposed to rocks, ice, or mud  Damage due to personnel using the bundle as handholds or footsteps  Damage due to shifting cargo wire 2-45

Figure 2-20 — Routing cables.

Never support any wire or wire bundle from a plumbing line carrying flammable fluids or oxygen. Use clamps on these lines only to ensure separation of the wire bundle from the plumbing line. Whenever possible, route wires and bundles parallel with or at right angles to the stringers or ribs of the area involved. See Figure 2-20. Don’t install single wires or wire bundles with excessive slack. Slack between support points, such as cable clamps, should not normally exceed 1/2 inch. This is the maximum you should be able to deflect the wire with moderate hand force. You may exceed this slack if the wire bundle is thin and the clamps are far apart. The slack must never be so large that the wire bundle can touch any surface. Allow a sufficient amount of slack near each end for the following reasons:  To permit ease of maintenance.  To allow replacement of terminals at least twice.  To prevent mechanical strain on the wires, cables, junctions, and supports.  To permit free movement of shock and vibration mounted equipment.  To permit shifting of installed equipment for purposes of maintenance. 2-46

Conduit and Fittings In many aircraft, the use of conduit is limited. This is a good practice because it saves weight and ensures wide separation of cables, making the electrical system less vulnerable to gunfire. However, some current aircraft, especially those with limited space for wire routing, use conduit. Conduit comes in two basic types, flexible and rigid. Its chief functions are to act as radio shielding and as a support and protection for wires. Conduit fittings attach either flexible or rigid conduit to junction boxes and other equipment, and usually include ferrules and coupling nuts. Various forms of both are in use along with special designs of locknuts, box connectors, and coupling adapters. Couplings are straight or angular in design so they fit all needs. Ferrules are bushings or flanges applied to the ends of the conduit to give greater strength and support to the coupling nuts. They are either crimped or swaged on by the use of crimping or swaging tools. SAFETY WIRE There are three types of safety wire: 1. Lockwire is a heavy wire used to secure parts against accidental opening. Use lockwire in all areas of high vibration such as an aircraft’s compartment. Electric connectors are lockwired in high-vibration areas that are inaccessible for periodic maintenance and inspection. 2. Shear wire is a lighter wire used to secure parts subject to periodic disconnection, maintenance and inspection, or for parts you need to remove quickly. 3. Seal wire is a thin, easily breakable wire used as a seal on fire-extinguishing systems, oxygen regulators, and other emergency devices that need a quick release for use. You use seal wire to show whether these devices have been used or tampered with. Bonding and Bonding Devices A bond is a union that exists between two metallic objects and results in electrical conductivity between them. Aircraft electrical bonding is the process of obtaining electrical conductivity between aircraft metallic parts or between the aircraft structure and installed equipment. An isolated conducting part or object is one physically separate from the aircraft structure and other conductors bonded to the structure. A bonding connector provides the electrical conductivity between metallic aircraft parts not in electrical contact. Bonding jumpers and bonding clamps are examples of bonding connectors. Purpose An aircraft can become highly charged with static electricity while in flight. If the aircraft doesn’t have a proper bond, all metal parts won’t have the same charge. A difference of potential will then exist between various metal surfaces. Neutralization of the charges flowing in paths of variable resistance produces electrical disturbances (noise) in the radio receiver. If the resistance between isolated metal surfaces is large enough, charges can collect, causing a spark, which is a fire hazard. If lightning strikes the aircraft, a good conducting path lessens severe arcs and sparks that could damage the aircraft and possibly injure its occupants. 2-47

The aircraft structure is also the ground for the radio. For the radio to function properly, a proper balance is necessary between the aircraft structure and the radio antenna. This means the surface area of the ground must be constant. Control surfaces, for example, may become partially insulated from the remaining structure. This is caused by a film of lubricant building up on the hinges. This would affect radio operation if bonding did not take care of the condition. Bonding also provides the necessary low- resistance return path for single-wire electrical systems. The reasons for bonding are summed up as follows:  To reduce radio and radar interferences by equalizing static charges that accumulate.  To reduce the fire hazard by preventing static charges from accumulating between two isolated members and creating a spark.  To minimize lightning damage to the aircraft and its occupants.  To provide the proper ground for proper functioning of the aircraft radio.  To provide a low-resistance return path for single-wire electrical systems.  To provide a means of bringing the entire aircraft to the earth’s potential and keeping it that way while it is grounded to the earth. Parts Requiring Bonding Current naval aircraft design keeps the number of bonding jumpers to a minimum. As a result, jumpers are very important and need replacing whenever necessary to keep them in good condition. Some of the aircraft parts that require bonding are identified below:  Control surfaces. Each control surface should have at least two bonding jumpers. This doesn’t apply to trim tabs.  Engine mounts. Use at least four bonding jumpers across each engine mount support. This provides a current path between the engine mount and aircraft structure.  Engine cowling. Use at least four symmetrically placed bonding jumpers to bond the engine ring cowling to the engine across the rubber mounts at the front end of the cowling.  Equipment mounts. Place bonding jumpers across shock mounts supporting electrical and radio equipment and the instrument panel. Methods and Materials You should install bonding connections so they won’t break or loosen. This prevents a variation in the resistance during movement. One primary objective for bonding is to provide an electrical path of low dc resistance and low RF impedance; therefore, it is important that the jumper be a good conductor of ample size. Make the bonding wire as short as possible. Bond parts directly to the basic aircraft structure rather than through other bonded parts, insofar as practical. Install bonding jumpers so they do not interfere with aircraft movable components. See Figure 2-21.

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Contact of dissimilar metals in the presence of an electrolyte, such as salt water, produces an electric action (battery action), which causes corrosion in the connection. The intensity of this electric action varies with the kinds of metals. Bonding frequently causes the direct contact of dissimilar metals. In such cases, the metals used are of the kind that produces minimum corrosion. You should make connections so that if there is co rrosion, it’s in replaceable elements such as jumpers, washers, or separators. Don’t use self-tapping screws for bonding purposes. Jumpers shouldn’t be compression-fastened through plywood or other nonmetallic material. When you are performing a bonding operation, clean the contact surfaces of insulating finishes or surface films before assembling. After installation, refinished the assembly with a suitable protective finish. You should refer to NAVAIR 01-1A-505 (series) for detailed information about bonding procedures. Cable Lacing and Tying Wire groups and bundles are laced or tied to provide ease of installation, maintenance, and inspection. This process is commonly referred to as spot tying. Spot tying keeps cables neatly secured in groups and avoids possible damage from chafing against equipment or interference with equipment operation. You tie a group or bundle of wires by using individual pieces of cord tied around the group or bundle at regular intervals. You lace a bundle or group of wires by securing them together inside enclosures by a continuous piece of cord. You use this cord to form loops (half hitches) at regular intervals around the group or bundle. Cotton, nylon, or fiber glass cord is used to tie or lace. The cotton cord has a wax coating to make it resist moisture and fungus.

Figure 2-21 — Bonding methods. 2-49

Sleeving Sleeving helps protect connections from accidental shorting and moisture and to lengthen the arc-over path between contacts. Don’t use insulating sleeves on connections or connectors that require moisture-proofing. Don’t use sleeving on connectors that have a sealing grommet that covers the soldered connection. Refer to the NAVAIR 01-1A-505 (series) for all sleeving applications. Sleeving has many applications in naval aviation. Some of the applications are for covering soldered connections, open bus bars, and permanent splices. For general-purpose wiring, use either clear or opaque flexible vinyl sleeving. For high temperature applications (160°F to 400°F), you should use silicone rubber or fiber glass sleeving. Where resistance to synthetic hydraulic fluids or other solvents is necessary, use nylon sleeving, either clear or opaque. Bus bars normally have panels or junction boxes to protect them against shorting. If the bus bars aren’t enclosed, use of protective coating is desirable. You can use sleeving by slitting a piece of vinyl tubing and wrapping it around the bar after making all connections. Be careful to choose a tubing that has a large enough diameter to permit a generous overlap when tied in place (Figure 2-22, view A). You can protect open and over braided harness wires with damaged shielding with an Electromagnetic Interference (EMI) shielded wrap. This two layer sleeving provides a wraparound repair application and is for single shielded wire harnesses only (Figure 2- 22, view B). The external layer is Nomex and provides mechanical protection with a nickel plated copper internal layer, which provides EMI protection. A blue tracer differentiates this from the unshielded version. In preparation, this must be cut to length, trimmed, and tightly installed on harness using split rings (not shown) and a 90 degree overlap past the starting point. The external layer is secured with lacing tie approximately every 1-2 inches. Terminate the sleeving within 0.5 inches of the ends using metal bands. To cover a permanent splice resulting from a damaged single conductor wire (Figure 2- 22, view C), slip the sleeving on the wire before the splicing operation. After completing the splice, center the sleeving to cover the finished splice, and then heat from middle to the ends until sealant melts and begins to flow out of ends using approved compressed air/ nitrogen heating tool with a small termination sleeve reflector. You will use a slightly different arrangement of sleeving for continuous chafe protection using a tube-like Expando sleeve. Expando sleeving (Figure 2-22, view D) is usually installed at the initial build of the harness. However, while harness is in service, connector de-pinning (size best suited for application) allows the Expando sleeve to slide over the harness wires and terminates under the clamp fingers and applying teflon barrier tape to the ends and securing with tie string.

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Figure 2-22 — Typical use of sleeving.

M AINTENANCE OF MOTORS AND GENERATORS Since most motors and generators are not accessible during flight, the first sign of motor trouble in many instances is complete failure. Generators and motors are mechanically similar, so the preventive maintenance you perform is basically the same. You should be sure that the motor and/or generator mount is secure. Inspect the mechanical linkage for proper alignment and security. Check that the electrical connections are secure and clean. Check for signs of corrosion. Make sure of proper voltage. You should inspect the exposed portion of the windings for evidence of overheating. If the insulation on the windings or leads is cracked and brittle, replace it. If, you smell burning insulation, or notice smoke or excessive noise, replace the motor. 2-51

When dismantling a motor, you should carefully remove the bearings and wipe them clean. Be sure to wrap the bearings in clean oil paper until needed during reassembly. Replace bearings showing pronounced stickiness or bumpy operation. During inspection of bearing assemblies, check for the presence of cracks and pitted surfaces. Also, check for any physical damage present in the bearing elements. Most ac motors are pre-lubricated at the time of manufacture and have sealed bearings. The bearings require a visual inspection only. If the bearings are damaged, turn the motor in for overhaul. Before reassembling the motor, wipe the rotor and stator clean with a lint-free cloth. During cleaning operations, use the correct cleaning solvents. As a final check, conduct an operational inspection of the motor, using the applicable MIM. Corrosion control is a particularly important factor when dealing with rotating machinery. Corrosion inspection includes the following:  An adequate cleaning program  Thorough periodic lubrication  Detailed inspection for corrosion and failure of protective systems  Prompt treatment of corrosion and touchup of damaged paint areas Effective protection begins with a comprehensive cleaning schedule. There is no single cleaning agent or process that will clean all parts. A cleaning agent that will clean one set of parts may not clean another. The cleaning agent may not be usable because it attacks the alloys or metal making up the part. Therefore, different cleaning agents are necessary. The selection of these agents will vary for different surfaces and equipment. Refer to Cleaning and Corrosion Control, NAVAIR 01-1A-509 (series) manual for information on cleaning agents and their use. In addition to approved dry-cleaning solvents, you can use trichloroethylene and inhibited methyl chloroform to clean electrical and electronic equipment. Inhibited methyl chloroform (trichloroethane) does present hazards to personnel and insulation. For information on such hazards, refer to Naval Ships’ Technical Manual (NSTM), Chapter 9600. Insulation can be affected by methyl chloroform. Insulation exposed to methyl chloroform deteriorates proportionally with the time of exposure to the solvent. After a 5- minute exposure, methyl chloroform noticeably softens varnishes and reduces dielectric strength and abrasion resistance. A continuous immersion of 1 hour completely destroys the properties of most insulating materials and varnishes used in armatures, field coils, and similarly constructed electrical components. Glass melamine and laminated phenolic, however, are only slightly affected by such immersion. Methyl chloroform itself is nonflammable, but after 90 percent evaporation, the residue contains a high percentage of the flammable inhibitor. When the solvent evaporates from containers, you must recognize the flammability of the residue. Some of the special precautions you should observe when using methyl chloroform include the following:  Avoid prolonged or repeated breathing of vapor or contact with the skin. Do not take internally.  Prevent contact with open flame, as this may form highly toxic phosgene.  Immerse electrical equipment less than 5 minutes, as prolonged immersion destroys most insulating materials. 2-52

WARNING

When using cleaning solvents; avoid burning the skin or inhaling fumes. Use solvents in a well-ventilated room only.

 Test the solvent on a portion of the component. This ensures that it will not destroy, blister, or otherwise damage the material.  DO NOT use on oxygen equipment —the inhibiting agent is flammable.  DO NOT use on hot equipment — accelerated evaporation will increase the toxic hazard. There are a variety of reasons for generator failures or apparent failures. Before removing a generator that isn’t delivering its rated voltage or current, determine that the trouble is not a problem in the control, feeder, or regulating circuits. You should refer to the applicable overhaul instruction manuals as a guide in determining exact voltage and frequency tolerance during these tests. Periodic inspection of generators includes the following:  Check for defects in the mounting flange.  If you find defects, remove and replace the generator. Inspect mounting studs, nuts, and safety wire for security.  Check removable units such as ventilation tube, air blast cover, brush inspection band, terminal box, end shield, and conduit for tightness and general mechanical condition. Repair any minor dents. Replace any parts with cracks.  Check the generator and cable connections for overheating. Discoloration and a burnt odor show overheating. Clean or replace the metal parts. Repair or replace damaged cables.  If you find defects you cannot repair with the generator installed, remove it for repair or overhaul. The maintenance and testing techniques used to maintain ac generators apply to dc generators. For detailed instructions on a specific generator, refer to the applicable manuals that are provided for the aircraft and component. Corrosion control of generators is the same as for motors /inserters. Generators are not likely to be overloaded. A defective load (one that is highly inductive and not the equivalent of resistance load) may result in low voltage, excessive field current, oversaturation, and continued low voltage. Any condition resulting in abnormal field current may result in overheating. This condition is dangerous to the affected generator and the entire system, dc as well as ac. PRINTED CIRCUITS The move toward replaceable units has led to several new methods of electronic equipment construction. The printed circuit is an example of this construction. This design or type of circuit gives speed and economy to maintenance, as well as saving space and weight. 2-53

Figure 2-23 — Printed circuit module. Circuit Construction Photo etching is one method used to manufacture printed circuits. In this method, a plastic or phenolic sheet with a thin layer of copper coating is used. The exposure of the printed circuit is similar to a photographic exposure. The copper coating is covered with light-sensitive enamel and a template of the circuit that appears when the sheet is exposed to light. The entire sheet is then exposed to light. The exposed area of copper reacts to the light. An etching process then removes this area. The enamel on the unexposed circuit protects the copper from the etching bath that removes the exposed copper. After the etching bath, the enamel is removed from the printed circuit. This leaves the surface ready for soldering parts and connections. Some manufacturers use machinery to mount standard parts like capacitors, resistors, and transistors, further speeding manufacture. Circuits produced by this method operate as well as conventional circuits and are easy to repair. Look at Figure 2-23. From the troubleshooters’ standpoint, it shows an improved type of construction that consists of a removable modular subassembly. The modules have numerous internal and external test points to make troubleshooting easier. Most test racks have plug-in extensions that let you raise the module so all parts are accessible for test and repair. The module is not expendable, but it is easy to repair, since all parts are of conventional design. Miniature and subminiature parts are so common in today’s electronic equipment that they are considered to be conventional.

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Printed Circuit Maintenance Although troubleshooting procedures for printed circuits are similar to those for conventional circuits, repair of printed circuits requires considerably more skill and patience. According to The Naval Aviation Maintenance Program (NAMP), COMNAVAIRFORINST 4790.2 (series), only Fleet Readiness Centers (FRCs) being certified as having the capability may perform micro-miniature circuit repair. Activity certification means that the FRC has individually-certified technicians assigned and NAVAIR-approved, operable repair equipment available, and does not include any specific authorization to repair. Initial repair technician certification is granted only upon successful completion of the appropriate Center for Naval Aviation Technical Training (CNATT) course or equivalent contractor training course, approved and provided by NAVAIR. Each individual must be recertified periodically to assure continued quality of performance. TEST EQUIPMENT You can classify test equipment as either common or peculiar. In some cases, test equipment may be peculiar to a specific system, but that system may be used on several different aircraft. Test equipment can also be classified by where it is used. You use line test equipment on the aircraft to determine a system’s operational readiness and to isolate malfunctions to a particular part. Normally, this type of equipment is painted yellow to distinguish it from the aircraft equipment. Bench test equipment is used in the shop to isolate malfunctions to a particular module, subassembly, or to the exact resistor, capacitor, etc. Occasionally, you will use line test equipment with bench test equipment to perform complete checks of a part. Whenever you use test equipment, follow the instructions in the MIM or other operating instructions for the specific test equipment, aircraft, or system. Never try to troubleshoot equipment from memory! By using test equipment instructions and the references for equipment under test, you prevent false findings. You also prevent damage to test equipment and systems and the loss of time. Aircraft Engine Component Test Stand (AECTS) The Aircraft Engine Components Test Stand (AECTS), (Figure 2-24) is an integrated test system which provides testing capability of aircraft engine-driven accessories, such as generators and generator drive systems, which will be discussed in Chapter 5. It provides testing for a wide variety of aircraft electrical components. The AECTS design allows for tailoring to meet space constraints aboard aircraft carriers and will be deployed both afloat and ashore at Fleet Readiness Centers (FRC) and Marine Aviation Logistics Squadrons (MALS) for validating ready-for-issue status of components, verifying operation after a repair action, and troubleshooting and fault-isolating generator system components. It will provide horsepower, shaft speed, and electrical loading requirements to test all aircraft power generating system components. The AECTS, with necessary adapters, is a modular system consisting of four major assemblies; a Variable Speed Drive (VSD) System, Test Set Instrumentation, Hydraulic Cooling System, and Load Banks. The AECTS is capable of interfacing with the 440 vac, 60 Hertz, three-phase, 4-wire electrical system. The total area does not exceed 2-55

Figure 2-24 — Aircraft Engine Components Test Stand (AECTS). eighty square feet and facilitates shipboard installation. A means of isolating the AECTS from shipboard power is provided by an internal isolation transformer.

Variable Speed Drive The variable speed drive is capable of operating at any speed up to 31,000 RPM with a 150 horsepower output. Test Set Instrumentation The test set instrumentation provides the controlling, selecting, and monitoring of test parameters of the Unit Under Test (UUT). It displays both the selected and actual parameters of the VSD output speed and the ac and/or dc load of the UUT. The control console, which is part of the test set instrumentation, is approximately 1800 pounds. 2-56

Hydraulic Cooling System The hydraulic cooling system provides conditioned oil to the gearbox, the constant speed drive, and any oil-cooled component under test. The hydraulic assembly is 36 inches high, 34 inches deep and 35 inches wide. It weighs approximately 500 pounds. Load Bank The load bank is capable of providing loads of infinite resolution, and has a shock load of a preset value for all existing electrical system components. It is capable of balanced phased loading, or for loading each phase separately on both resistive and reactive loads. A cooling fan maintains the load elements at a safe operating temperature. The load bank is 71 inches high, 63 inches deep and 42 inches wide, and weighs approximately 2,500 pounds. Semiconductor Testing Since semiconductors have replaced vacuum tubes, the testing of semiconductors is vital. In this section, three basic types of equipment are discussed, the Huntron Tracker 1000, Huntron Tracker 2000, and the Automatic Transistor Analyzer Model 900 in-circuit transistor tester. Huntron Tracker 1000 You will test components with a Huntron Tracker 1000 using a two-terminal system, where two test leads attach to the leads of the component under test. The 1000 tests components in-circuit, even when there are several components in parallel. The following types of devices are tested using the Huntron Tracker 1000:  Semiconductor diodes  Bipolar and field effect transistors  Bipolar and MOS integrated circuits (both analog and digital)  Resistors, capacitors, and inductors The 1000 is used on boards and systems with ALL voltage sources in a power-off condition. A 0.25 ampere signal fuse (F1) connects in series with the channel A and B test terminals. Accidentally contacting test leads to active voltage sources (e.g., line voltage, powered-up boards or systems, charged high voltage capacitors, etc.) may cause this fuse to open, making replacement necessary. When the signal fuse blows, the display shows open circuit signatures, even with the test leads shorted together.

The line fuse (F2) should only open when there is an internal failure inside the instrument. Therefore, you should always locate and correct the problem before replacing F2. CAUTION The device to be tested must have all power turned off and have all high-voltage capacitors discharged before connecting the 1000 to the device. 2-57

The front panel of the 1000 makes function selection easy. The 1000 uses interlocking pushbutton switches for range selection. A toggle switch is used for channel selection, and integral LED indicators show the active functions. The CRT displays the signatures of the parts under test. The display has a graticule consisting of a horizontal axis that represents voltage, and a vertical axis that represents current. The horizontal axis is divided into eight divisions, which lets you estimate the voltage at which signature changes occur. This is mainly useful in determining semiconductor junction voltages under either forward or reverse bias. Push in the power on/off switch. The 1000 should come on line with the power LED illuminated. Before you can analyze signatures on the CRT, you must focus the 1000. To do this, turn the intensity control to a comfortable level. Now, adjust the focus control (back panel) for the narrowest possible trace. Aligning the trace is important in determining the voltages at which changes in the signature occur. With a short circuit on channel A, adjust the horizontal control until the vertical trace is even with the vertical axis. Open channel A, and adjust the vertical control until the horizontal trace is even with the horizontal axis. Once set, you should not have to adjust these controls during normal operation. Turn the power off by pushing the power switch in. When you turn the power on again, the same intensity setting will be present. The 1000 has three impedance ranges, low, medium, and high. To select these ranges, press the appropriate button on the front panel. Always start with the medium range; then you can adjust for other ranges. If the signature on the CRT is close to an open (horizontal trace), try the next higher range for a more descriptive signature. If the signature is close to a short (vertical trace), try the next lower range. There are two channels (channel A and channel B) that you can select by moving the toggle switch to the desired position. When using a single channel, plug the red probe into the corresponding channel test terminal. Then plug the black probe into the common test terminal. When testing, connect the red probe to the positive terminal of the device (i.e., anode, +V, etc.). Connect the black probe to the negative terminal of the device (i.e., cathode, ground, etc.). By following this procedure, the signature will appear in the correct position on the CRT display. The alternate mode of the 1000 provides automatic switching back and forth between channel A and channel B. This allows easy comparison between two devices or the same point on two circuit boards. You select the alternate mode by moving the toggle switch to the ALT position. The alternate mode is useful when comparing a known good device with the same device whose quality is unknown. The signal section applies the test signal across two terminals of the device under test. The test signal causes current to flow through the device and a voltage drop across its terminals. The current flow causes a vertical deflection of the signature on the CRT display. The voltage across the device causes a horizontal deflection of the signature on the CRT display. The combined effect produces the current-voltage signature of the device on the CRT display. An open circuit has zero current flowing through the terminals and a maximum voltage across the terminals. In the LOW range, a diagonal signature from the upper right to the lower left of the CRT (Figure 2-25, view A) represents an open circuit. In the HIGH and MEDIUM ranges, an open circuit shows as a horizontal trace from the left to the right (Figure 2-25, view B). When you short the terminals together, the maximum current 2-58

flows through the terminals, and the voltage at the terminals is zero. A vertical trace from the top to the bottom of the CRT graticule in all ranges shows this short (Figure 2- 25, view C). The CRT deflection drivers boost the low-level outputs from the signal section to the higher voltage levels needed by the deflection plates in the CRT. The HORIZONTAL and VERTICAL controls on the front panel adjust the position of the trace on the CRT display. You use three other CRT controls to adjust the brightness and clarity of the trace— INTENSITY, FOCUS, and ASTIGMATISM. The front panel intensity control is the primary means of adjusting the visual characteristics of the trace. The focus control is on the back panel and is operator adjustable. The astigmatism trim pot is inside the 1000 on the main printed circuit board. The pot is factory adjusted to the correct setting. Huntron Tracker 2000 The Huntron Tracker 2000 (Figure 2-26) is a versatile troubleshooting tool having the following features:  Multiple test signal frequencies (50/60 Hz, 400 Hz, 2000 Hz)  Four impedance ranges (low, medium 1, medium 2, high)  Automatic range scanning  Range control: High Lockout  Adjustable rate of channel alteration and/or range scanning  Dual polarity pulse generator for dynamic testing of three terminal devices  LED indicators for all functions  Dual channel capability for easy comparison  Large CRT display with easy to operate controls Figure 2-25 — Circuit signatures: View A – Low-range open circuit; view B – medium and high range open circuit; and view C – all ranges short circuit.

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Figure 2-26 — Huntron Tracker 2000.

General Operation You will test components using the 2000 two terminal system. It also has a three terminal system when using the built-in pulse generator. When using this system, you place two test leads on the leads of the component under test. The 2000 tests components in-circuit, even when there are several parts in parallel. Use the 2000 only on boards and systems with all voltage sources in a power-off condition. A 0.25 ampere signal fuse connects in series with the channel A and B test terminals. Accidental contact of the test leads to active voltage sources, such as line voltage, powered-up boards or systems, and charged high voltage capacitors, may cause this fuse to open, making replacement necessary. When the signal fuse blows, the 2000 displays short circuit signatures even with the test leads open.

The line fuse should only open when there is an internal failure inside the instrument. Always locate the problem and correct it before replacing this fuse.

CAUTION The device under test must have all power turned off and all high-voltage capacitors discharged before connecting the 2000 to the device. 2-60

Front Panel The front panel of the 2000 makes function selection easy. All push buttons are the momentary action type. Integral LED indicators show which functions is active. Look at Figure 2-27 and Table 2-4 for details about each item on the front panel.

Back Panel Secondary controls and connectors are located on the back panel (Figure 2-28 and Table 2-5).

Figure 2-28 — Back panel.

Figure 2-27 — Front panel.

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Table 2-4 — Front Panel Controls and Connectors Item No. Name Function 1 HORIZ Control Controls the horizontal position of the CRT display. 2 VERT Control Controls the vertical position of the CRT display. 3 INTENSITY Control & Power On / Off Switch Controls the intensity of the CRT display. Power Switch: Rotate clockwise to turn on. 4 TRACE ROTATE Control Controls the trace rotation of the CRT display. 5 CRT Display Display the component signatures produced by the 2000. 6 Range Selectors Push button that select one of four impedance ranges: low, medium1, medium 2, and high. 7 AUTO Switch Push button that initiates automatic scanning of the four ranges from low to high. The scanning speed is determined by the RATE control (see item #4). 8 Channel A Switch Push button that causes channel A to be displayed. 9 Channel A Test Terminal Fused test lead connector that is active when channel A is selected. All test lead connectors accept standard banana plugs. 10 ALT Switch Push button that causes the 2000 to alternate between channel A & channel B at a speed determined by the RATE control (see item #4). 11 COM Test Terminal Test lead connector that is instrument common & the common reference point for both channel A & channel B. 12 Channel B Switch Push button that causes channel B to be displayed. 13 Channel B Test Terminal Fused test lead connector that is active when channel B is selected. 14 RATE Control Controls the rate of channel alternation &/or range scanning. 15 G1 & G2 Terminals Pulse generator output test lead connectors. 16 WIDTH Control Controls the duty cycle of the internal pulse generator. 17 LEVEL Control Controls the amplitude of the internal pulse generator. 18 Pulse Generator Selectors Push buttons that select various output modes of the pulse generator: +. -, PULSE/DC. 19 Frequency Selectors Push buttons that select one of three test signal frequencies: 50/60Hz, 400Hz, 2000Hz. 20 HIGH LOCKOUT Switch Push button that activates a mode where it is not possible to enter the high range either by manual or automatic range selection.

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Table 2-5 — Back Panel Controls and Connectors Item No. Name Function 1 Accessory Output Connector Connector which provides the power and clock to the Huntron Switcher Model HSR410. 2 Power Chord Connector IEC standard connector that mates with any CEEE-22 power cord. 3 FOCUS Control Controls the focus on the CRT display.

CRT Display The signature of the part under test is displayed on the CRT. The display has a graticule consisting of a horizontal axis that represents voltage, and a vertical axis that represents current. The axes divide the display into four quadrants. Each quadrant displays different portions of the signatures. The displays are:  Quadrant 1 displays positive voltage (+V) and positive current (+I)  Quadrant 2 displays negative voltage (-V) and positive current (+I)  Quadrant 3 displays negative voltage (-V) and negative current ( –I)  Quadrant 4 displays positive voltage (+V) and negative current ( –I) The horizontal axis divides in eight divisions, which allows the operator to estimate the voltage at which changes in the signature occur. This is useful in determining semiconductor junction voltages under either forward or reverse bias. Operation of Panel Features Turn the Power/Intensity knob clockwise. The 2000 comes on with the LEDs for power, channel A, 50/60 Hz, low range, and pulse/DC illuminated. Focusing the 2000 display is an important part of analyzing the test signatures. First adjust the intensity control to a comfortable level. Then, adjust the focus control (back panel) for the narrowest possible trace. Aligning the trace is important in determining which quadrants the portions of a signature are in. With a short circuit on channel A, adjust the trace rotation control until the trace is parallel to the vertical axis. Adjust the horizontal control until the vertical trace is even with the vertical axis. Open channel A, and adjust the vertical control until the horizontal trace is even with the horizontal axis. Once set, you should not have to readjust these settings during normal operation. Range Selection – The 2000 has four impedance ranges, low, medium 1, medium 2, and high. You select these ranges by pressing the appropriate button on the front panel. Start with one of the medium ranges, i.e., medium 1 or medium 2. If the signature on the CRT display is close to an open (horizontal trace), select the next higher range for a more descriptive signature. If the signature is close to a short (vertical trace), select the next lower range. The High Lockout feature, when activated, prevents the instrument from entering the high range. This feature works in either the manual or auto mode. 2-63

The auto feature scans through the four ranges, three with the HIGH LOCKOUT activated at a speed set by the RATE control. This feature allows you to see the signature of a part in different ranges while freeing your hands to hold the test leads. Channel Selection – There are two channels on the 2000-channel A and channel B. You select a channel by pressing the appropriate front panel button. When using a single channel, plug the red probe into the corresponding channel test terminal. Plug the black probe into the common test terminal. When testing, connect the red probe to the positive terminal of the device; i.e., anode, +V, etc. Connect the black probe to the negative terminal of the device; i.e., cathode, ground, etc. Following this procedure should assure that the signature appears in the correct quadrants of the CRT display. The ALT mode is a useful feature of the 2000. It lets you compare a known good device with a device of unknown quality. In this test mode, you use common test leads to connect two equivalent points on the boards to the common test terminal. The ALT mode of the 2000 allows you to automatically switch back and forth between channels A and channel B, so you can easily compare two devices. You may also compare the same points on two circuit boards. Select the ALT mode by pressing the ALT button on the front panel. You may vary the alternation frequency by using the RATE control.

When using the alternate and auto features simultaneously, each channel is displayed before the range changes. Figure 2-29 shows the sequence of these changes. Frequency Selection – The 2000 has three test signal frequencies, 50/60 Hz, 400 Hz and 2000 Hz. You can select these by pressing the appropriate button on the front panel. In most cases, you should start with the 50/60 Hz test signal. Use the other two frequencies to view small amounts of capacitance or large amounts of inductance. Pulse Generator – The built-in pulse generator of the 2000 allows dynamic, in-circuit testing of certain devices in their active mode. In addition to using the red and black probes, you use the pulse generator. The output of the pulse generator connects to the control input of the device under test with one of the blue micro clips provided. The pulse generator has two outputs, G1 and G2, so you can test three terminal devices in the alternate mode. NOTE The black probe plugs into the channel B test terminal. Figure 2-29 — Auto/alternate sequence.

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A variety of output waveforms is available using the pulse generator selector buttons. First select the pulse mode or the dc mode using the PULSE/DC button.  In the pulse mode, the LED flashes at a slow rate.  In dc mode, the LED is continuously on. Then select the polarity of output desired using the positive (+) and negative (–) buttons. All three buttons function in a push-on/push-off mode, and only interact with each other to avoid the NOT ALLOWED state. After selecting the specific output type, set the exact output using the LEVEL and WIDTH controls. The LEVEL control varies the magnitude of output amplitude from zero to 5 volts (peak or dc). During pulse mode, the WIDTH control adjusts the duty cycle of the pulse output from a low duty cycle to 50 percent maximum (square wave). The start of a pulse is triggered by the appropriate zero crossing of the test signal. This results in the pulse frequency being equal to the selected test signal frequency. The WIDTH control setting that selects the duty cycle determines the end of a pulse. The WIDTH control has no effect when in the dc mode. Troubleshooting Tips You will use the Huntron Tracker 1000 and the Huntron Tracker 2000 to test various types of devices and circuits. Some troubleshooting tips are:  Perform most tests using the medium or low range.  Use the high range only for testing at a high impedance point, or if higher test voltages are required (i.e., to test the Zener region of a 40-volt device).  Sometimes, component defects are more obvious in one range than another.  If a suspect device appears normal for one range, try the other ranges.  Use the low range when testing a single bipolar junction, such as a diode, a base-emitter junction, or a base-collector junction. It offers the best signature.  Use a higher range to check for reverse bias leakage.  When performing in-circuit testing, do a direct comparison to a known good circuit.  The 1000 test leads are not insulated at the tips. Be sure to make good contact to the device(s) under test. This tip pertains to the 1000 only. When you troubleshoot, try relating the failure mode of the circuit under test to the type of defect the 1000 shows. For example, expect a catastrophic printed circuit board failure to have a dramatic signature difference from that of a normal device of the same type. A marginally operating or intermittent board may have a failed part that shows only a small pattern difference from normal. If you cannot relate a system failure to a specific area of the printed circuit board, begin by examining the signatures at the connector pins. This method of troubleshooting shows all the inputs and outputs. It will often lead directly to the failing area of the board. Devices made by different manufacturers, especially digital integrated circuits, are likely to produce slightly different signatures. This is normal and may not show a failed device. Remember, leakage current doubles with every 10-degree Celsius rise in temperature. Leakage current shows up as a rounded transition, where the signatures show the 2-65

NOTE The 2000 test leads are conductive only at the tips. Be sure to make good contact with the device(s) under test. change from zero current flow to current flow, or by causing curvature at other points in the signatures. Leakage current causes curvatures due to its nonlinearity. Never begin the testing of an integrated circuit using the low range. If you initially use the low range, confusion can result from the inability of this range to display the various junctions. Always begin testing using the medium range. If the signature is a vertical line, switch to the low range. Here you can check for a short or low impedance (less than 500 ohms). Switch to the low range if the device is suspect and appears normal in the medium range. This will reveal a defective input protection diode not evident when using the medium range.

When testing analog devices or circuits, use the low range. Analog circuits contain many more single junctions. Defects in these junctions show more easily when using the low range. Also, the 54-ohm internal impedance in the low range makes it less likely that parts in parallel with the device under test will sufficiently load the tester to alter the signature. When testing an op amp in-circuit, compare it directly to a known good circuit. This is because the many different feedback paths associated with op amps can cause an almost infinite number of signatures. Often when checking a Zener diode in-circuit, it will not be possible to examine the Zener region due to circuit leakage. If you must see the Zener region under this condition, unsolder one side of the diode to eliminate the loading effects of the circuit. Huntron Tracker 1000 – Bipolar integrated circuits containing internal shorts produces a resistive signature (a straight line). This line begins in the 10 o’clock to 11 o’clock position. It ends in the 4 o’clock to 5 o’clock position on the display when using the low range. This type of signature is always characteristic of a shorted integrated circuit, as it results from a resistive value of 4 to 10 ohms. A shorted diode, capacitor, or transistor junction always produces a vertical (12 o’clock) straight line using the low range. Huntron Tracker 2000 – Bipolar integrated circuits containing internal shorts produce a resistive signature (a straight line) beginning in the 1 o’clock to 2 o’clock position. This signature ends in the 7 o’clock to 8 o’clock position when using the low range. This type of signature is characteristic of a shorted integrated circuit. This results from a resistive value of 4 to 10 ohms. A shorted diode, capacitor, transistor junction, etc., always produces a vertical (12 o’clock) straight line when using the low range. Jet Ignition System Tester The jet ignition system tester detects and isolates faults in the jet engine ignition system. You can use the tester to make the following checks:  Operational check of the ignition system. You are able to check the engine through the ignition unit to the spark plug.  Operational check of the spark plugs.  Check of the ignition unit output in sparks per second. 2-66

 Power input to the ignition unit. The panel of the ignition system tester is shown in Figure 2-30. If you want the procedures for using the tester, refer to the tester’s operation and service instruction manual and the engine ignition system maintenance manual.

Tachometer Indicator Generator Test Set TTU-27/E, Part Number 8401B The tachometer indicator-generator test set, TTU-27/E, is used to test aircraft tachometer generators and indicators on both piston and jet engine aircraft. You can use the tachometer indicator-generator test set to make the following tests:  Test tachometer indicators from the master generator  Test tachometer generators both on and off-aircraft  Test tachometer generators and indicators as a unit Figure 2-31 shows a TTU-27/E tester completely self-contained in a light-weight aluminum case designed for portability. All the operating controls, switches, and indicators required for the tester are on the panel assembly. The variable speed drive pad, which is varied from 0-5000 rpm by a panel mounted control knob, accommodates the tachometer generator for testing. The master indicator provides a precision readout of tachometer generator speed with scales of RPM x 2, RPM x 1, and PERCENT RPM. The tester uses 115- volt, 60-400 hertz ac power.

Figure 2-30 — Jet ignition system test panel.

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Figure 2-31 — Tachometer Indicator-Generator Test Set TTU-27/E. A master tachometer generator is mounted inside the tester and be used to test the tachometer indicators. You can check indicator calibration by the master tachometer generator by using the INCREASE SPEED control and comparing to the MASTER INDICATOR readings with those of the indicator under test. Internal load banks are selected to load the generator under test, and, while under load, the phase to phase voltage may be selectively displayed on the panel mounted voltmeter by selecting the A-B, B-C and A-C TERM CONNECTION switch.

When testing the generator and indicator as a unit, the generator is mounted to the drive pad and the indicator to the GEN OUTPUT connector. The LOAD IN OHMS is set to off during this test. When the output selector switch is in the MASTER GEN position, the indicator under test is driven by the internal master generator, as is the speed of the MASTER INDICATOR. When the switch is in the TEST GEN position, the indicator under test is being driven by the generator under test, as is also the speed MASTER NOTE Don’t try to operate the tachometer tester until after you make the ground connection of the power lead. 2-68

INDICATOR. For additional information on the operation and service instruction of the TTU-27/E tester, refer to NAVAIR 17-15CM-4. Jet Calibration (JETCAL) Analyzer Of the many factors affecting jet engine life, efficiency, and safe operation, the most important are exhaust gas temperature and engine speed. If the exhaust gas temperature is just a few degrees higher than it should be, the turbine blade life is reduced as much as 50 percent. Abnormally high exhaust gas temperature resulting from excess engine speed can cause premature engine failure. If the exhaust gas temperature is too low, jet engine efficiency and thrust is reduced. Either of these two conditions makes engine operation dangerous. By using the JETCAL analyzer, you can detect or prevent these conditions. You can more accurately check signs of fuel system trouble, tailpipe temperature, and RPM with the JETCAL analyzer than with aircraft gauges. With proper use of the JETCAL analyzer, you can detect malfunctions in these systems. The JETCAL analyzer (Figure 2-32) is a rugged, portable instrument made of aluminum, stainless steel, and plastic. The major components of the analyzer are the thermocouples, RPM and EGT indicators, resistance and insulation check circuits, and the overheat detection test circuits.

Figure 2-32 — JETCAL analyzer. Figure 6-9 – Grounding procedure. 2-69

On the EGT system functional test and the thermocouple and harness checks, the JETCAL analyzer has an accuracy of ± 4°C at the test temperature. The test temperature is usually the maximum operating temperature of the jet engine. To find the maximum engine operating temperatures, you should check the applicable MIMs. The first test you perform is a functional test of the EGT system. In this test, you heat the engine thermocouples in the tail cone of the engine to test temperature. This heat comes from the JETCAL heater probes through the necessary cable connections. To measure the temperature of the heater probes, thermocouples are embedded in the probes. You may read the heater probe temperature on the JETCAL potentiometer. At the same time, you can read the aircraft thermocouple temperature on the aircraft EGT indicator. You can then compare the readings for accuracy. On engines that have a balancing type of thermocouple system, you must remove the balancing thermocouple from the circuit. Then, you can check the remaining thermocouples individually or together. Check the balancing thermocouple using a single probe. Read the output of the balancing thermocouple on the JETCAL potentiometer, and compare it to the heater probe thermocouple reading. Gas temperature is critical to engine operation; therefore, use the JETCAL for nozzle scheduling. During temperature adjustments, make all temperature readings on the JETCAL potentiometer. This is necessary because you must read engine temperature accurately to ensure the engine is operating at optimum engine conditions. When checking and adjusting the engine exhaust gas temperature, install a switch box in the EGT circuit before starting the test. You will use the switch box to switch the cockpit indicator into the circuit, or to switch the temperature indication of the engine thermocouple and harness to the JETCAL potentiometer. You can also read temperature readings from the engine thermocouple and harness on the JETCAL analyzer by making the necessary connections. Incorporated in the analyzer is the TAKCAL unit (RPM indicator) check circuit. This circ uit reads engine speed with an accuracy of ±0.1 percent. The TAKCAL unit check can also be used to troubleshoot the aircraft’s tachometer system. After testing the exhaust gas temperature and engine speed systems, you may use selected portions of the JETCAL analyzer circuits to establish the proper relationship between exhaust gas temperature and engine speed. The JETCAL analyzer requires a power supply of 95 to 135 volts, 50 to 400 hertz. It will operate in temperatures from –55°C to +70°C. You must use a 95- to 135-volt ac power supply for the TAKCAL unit check and the thermocouple check. You may perform all other operations using emergency batteries when an external power supply is not available. You actuate the batteries by a push-button switch. When the switch is released, the batteries are out of the circuit. However, you can depress the switch when using ac without damaging the batteries. To preserve the life of the emergency batteries, use an external ac power supply whenever possible. The JETCAL analyzer has the following primary and separate functions:  To check the entire jet aircraft exhaust gas temperature system for error without running the engine or disconnecting the wiring.  To check individual thermocouples before placing them in the aircraft.  To check each engine thermocouple for continuity.  To check the thermocouples and harness for accuracy of output. 2-70

Figure 2-33 — Synchrophaser test set.  To check the resistance of the EGT circuit, without the EGT indicator, to assure allowable limits.  To check the insulation of the EGT circuit for shorts or grounds.  To check the EGT indicators.  To check engine thermocouples and harness on the engine with the engine removed from the aircraft.  To read engine RPM to an accuracy of ±0.1-percent engine run- up.  To use the RPM check (TAKCAL) and potentiometer to establish the proper relationship between exhaust gas temperature and engine speed during tabbing. (Tabbing is the procedure to adjust fixed or variable exhaust gas tail cone areas during normal aircraft checks every 30 to 50 hours.)  To check aircraft fire detector, overheat detector, and wing anti-icing systems by using tempcal probes. Refer to the MIM of a particular aircraft for the proper procedure when you check and adjust the variable nozzle systems. You can get detailed instructions for adjusting the EGT with the JETCAL analyzer from the applicable MIM. Synchrophaser Test Set The synchrophaser test set (Figure 2-33) tests propeller synchrophaser electronic units. The test set generates all pulses (dc and ac) required to test the synchrophaser electronic unit independent of its associated components. The test set is completely solid state. Each switch position programs a particular test. It does this by connecting the appropriate inputs to the synchrophaser and the meter required to verify synchrophaser performance. The test set generates a master pulse and a slave pulse. The pulse circuits closely simulate the synchrophaser pulse input under dynamic flight conditions. You program different phase and speed relationships of master and slave pulses, under test specifications, by setting the selector switches for a particular test.

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The output of synchrophaser-associated components is simulated by the test set. It provides a simulated tachometer signal and the resistance of synchrophaser controls for certain tests. When programmed, these signals help measure the synchrophaser dynamic response to off-speed, off-phase, speed reset, resync, and throttle anticipation signals. Control adjustments are minimal. Only a few of the automatically programmed tests require adjustment of either of the feedback gain potentiometers. The accuracy of control adjustment is optimized by designing test set circuits that require few adjustments. This adjustments will be to null or zero settings on center-scale zero meters. A high degree of accuracy and repeatability is a feature of the gain measuring circuits within the tester. Conventional gain circuits require application and measurement of incremental voltages to the synchrophaser for comparison with resultant output voltages. The gain measurement circuit has a galvanometers, demodulator, dc power supply, and calibrated potentiometer in a bridge circuit. When the gain test is programmed, rotate the calibrated potentiometer to null the galvanometers. With the galvanometers nulled, the potentiometer calibration gives a direct readout of amplifier gain. This circuit avoids inaccuracies of input settings and error amplification of incremental gain comparison. It also avoids the chance of error in calculating the gain factor. For complete instructions on this tester, you should refer to the current Operation and Service Instruction Manual, NA 17-15CFA-2. Propeller Synchronizer Tester (TTK-512), Part Number 1383AS200-1 The propeller synchrophaser system is tested using the propeller synchronizer tester (Figure 2-34) and is capable of testing solid state propeller synchrophaser systems on P-3, C-130, and E2/C2 type aircraft. The tester contains a microcomputer which senses switch positions during operation, and configures a test circuit which activate relays on a board. The microcomputer selects the proper signal conditioning path and conversion algorithm and displays performance data on the display panels and indicators. The tester is a portable flight line unit, completely solid state, and consists of a panel and chassis assembly housed in a carrying case. Additionally, there is an accessory case contains the accessory cables and wiring harnesses. The panel and chassis assembly contain all of the operating controls, indicators, components and electronic circuitry. The tester is approximately 15 inches long, 14 inches high, and 15 ½ inches wide, and weighs 50 pounds. When power is first applied to the tester, it initiates an automatic internal self- test of the components. When a zero appears after the countdown in DISPLAY # 2, the tester passed the internal self-test. If a number other than “0” appears, the tester is malfunctioning and should be rejected to the Intermediate Maintenance Activity (IMA). The synchronizer test set uses aircraft electrical power, 28vdc or external 28vdc, and monitors voltage inputs and outputs from the synchrophaser to determine if the synchrophaser or components of the propeller synchrophaser system and aircraft electrical wiring can be suspected of malfunction.

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Figure 2-34 — Propeller synchronizer tester (TTK-512).

Generally, initial testing of the synchrophaser system consists of resistance and megger checks. Test data should be recorded while performing these tests. You want to ensure resistance and megger tests are completed prior to performing the voltage tests. These tests are usually performed with the engines not running. Additional tests are performed to evaluate operation, such as voltage tests on ground with the engines shut down or with the engines running at 100% RPM. Tests can also be performed in flight with engines running at 100% RPM. Follow testing and troubleshooting procedures for current aircraft when using the propeller synchronizer test set. Procedures differ for each aircraft. For complete instructions on this tester, you should refer to the current maintenance instruction manual, AG-240AB-OMP-000.

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Test Set, Air Data TS-4508/U, Part Number ADTS-405-8325-M4 The Air Data Test Set (Figure 2-35) is used to verify proper operating parameters of pitot and static systems of aircraft. You can also conduct dynamic tests, quantitative calibration tests, and pneumatic-system leak tests using this tester. The Air Data Test Set consists of a rugged rack mounted on wheels within a weatherproof flight line case with a pump unit which provides regulated pressure and vacuum supplies to pitot and static pressure to related instruments, air data systems, and other auxiliary equipment. It is supplied with a standard hand terminal for one man operation which can be utilized by the operator during remote operating conditions. All the controls, indicators, switches, electrical, and pneumatic connections are on the front panel of the test set. The components are on the underside of the control panel. The test set simulates airspeed, altitude and rate of climb information, which is displayed on the front panel and hand held terminal. Refer to the NAVAIR 17-15CA-62 for additional information. See Table 2-6 for test set particulars. The test set has the following capabilities and features: 1. Turn-on procedures provide a self-test. 2. Provides three aircraft safety protection methods:  Limit checking of user entered set points.  Automatic regain of control if leak rate is over limit during leak testing.  Automatic regain of control if a leak takes the system pressures outside of limits. 3. Remote and local operation of unit. 4. The TS-4508/U can measure the following parameters:  Altitude measured in aeronautical units measured in ft or m.  Calibrated airspeed and true airspeed measured in kts or km/hr  MACH measured in units of math  Pressure Static (Ps) and Pressure total ( Pt) measured in pressure units.  Difference in pressure between Ps and Pt ( Qc) measured in pressure units.  Rate of change of altitude.  Rate of change of airspeed.  Rate of Ps.  Rate of Qc.  Rate of Pt.

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Figure 2-35 — Test set, TS-4508/U

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Table 2-6 — Test Set, Air Data TS-4508/U POWER REQUIREMENTS AC Input Consumption 103 - 127 Volts, 47 - 420 Hz 500 VA DIMENSIONS Depth Width Height Weight 12.1 inches 30.0 inches 18.1 inches 82 lbs ENVIRONMENTAL Warm-up period 15 minutes Operating temperature range 14 to 131 degrees F Storage temperature range -60 to 160 degrees F CONTROL STABILITY Altitude Greater of +/- 3 ft or +/-0.02% of simulated altitude Speed +/- 1 kt ACCURACY ALTITUDE Range -1500 to +65,000 feet Resolution 1 ft Accuracy Greater of +/- 10ft or +/- 0.1% reading ALTITUDE RATE Range 0-30,000 ft/min Resolution 1.0 ft/min Accuracy +/- 10.0% reading SPEED Range 20 to 1,000kts Resolution 0.1 kts Accuracy +/- 2.0kts

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Figure 2-36 — TTU-378 A/E fuel quantity test set.

Capacitive-Type Liquid Quantity Test Set 361046-001, Model TTU 378 A/E The test set is used to test and adjust aircraft fuel quantity gage probes, indicators, and complete gaging systems, with the equipment under test installed in the aircraft or removed for bench tests. The test set is housed within an aluminum combination case (Figure 2-36). The body of the combination case is divided into two sections. The lower section houses the test set internal mechanism and circuitry; the upper section of the case and the test set cover provides storage space for the accessory cables. The cover is secured to the case with two removable hinges at the rear and two trunk-style latches at the front of the case. The test set is operated within the combination case with the cover fully opened or removed.

CAUTION Dangerous voltages are present at system connectors. Ensure power is OFF prior to connecting or disconnecting cables. 2-77

Fuel Quantity System Check An aircraft's fuel quantity system is checked on board by connecting the test set between the indicator and tank units. In this configuration, the test set can be used to verify system operation, adjust the indicator to the tank units, or determine whether the indicator or probes are defective. When used to check the system, the test set can be used with any known amount of fuel in the tanks; It can modify the capacitance signal sent to the indicator for full and empty indicator adjustment. The test set can also measure the capacitance of the tank units for comparison against the expected capacitance. The testing mode for system check is controlled by the front panel CAPACITANCE FUNCTION and DISPLAY select switches. In the system check mode, only the capacitance circuitry of the test set is used. Insulation resistance is not tested in system check mode. Test Set Adjustment and Test The test set is adjusted and tested using standard resistors and capacitors in the set's internal circuitry. Capacitance circuitry adjustment is enabled by setting DISPLAY SELECT to CAP (PF) and CAPACITANCE RANGE (PF) to HIGH ADJ or LOW ADJ. To adjust the capacitance measuring circuitry, CAPACITANCE HIGH ADJ and LOW ADJ front panel pots are set so that the display reads the values marked on the CAP STANDARDS placard. Resistance measuring circuitry is enabled by setting DISPLAY SELECT to RES (MEG) and RESISTANCE RANGE (MEG) to HIGH TEST or LOW TEST. When in the HIGH TEST or LOW TEST positions, internal standard resistors are applied to the resistance measuring circuitry. The reading on the display is compared with the RESISTANCE TEST placard on the front panel. Refer to the NAVAIR 17-15BD-46 for additional information. When using this test set, you should follow the instructions in the applicable MIM for the system under test.

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Figure 2-37 — Fuel Control Test Set (FCTS) TTU-597/E. F uel Control Test Set (FCTS) TTU-597/E Many types of capacitive fuel quantity testers are in use in naval aviation. All operate on the same basic principle, that of a variable capacitor. The Fuel Control Test Set (FCTS) is designed for flight line use at the Organizational level, both afloat and ashore, to accurately and reliably test aircraft fuel quantity systems (Figure 2-37). The FCTS is compatible with all aircraft types which have a capacitive fuel content system. Adapter cable assemblies are provided to enable the FCTS to be used with the various aircraft configurations. The FCTS is a portable test set comprised of a test unit, transit case and accessory kit. The test unit is designed to operate from an internal battery pack, aircraft 28 vdc input cable, or an external ac/dc Power Converter Pack (PCP). An internal battery management system monitors the presence of an installed battery pack.

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The test unit is housed in a seamless, sealed aluminum case equipped with an automatic pressure release valve (breather valve). The case is comprised of the display compartment, electronics compartment and battery compartment. The case assembly (display compartment), is housed in the top portion of the test unit, which is hinged to the electronics compartment. The display is comprised of a 13-inch, 640 X 480 pixel electroluminescent screen, bezel and gasket. The display screen is connected to the electronics compartment by the display cable. The base assembly (electronics compartment) forms the center section of the test unit and houses the solid state electronic circuitry, F1 fuse and battery pack mating connector. All external connectors for operating the Test Unit are located on the top panel. The battery compartment is attached to the bottom of the electronics compartment. The interior of the battery compartment contains foam inserts that are formed to house the battery pack. The inserts provide a cushion surface, which prevents movement of the battery pack. The battery pack is comprised of the battery box, battery holder, mating connector cable (for connecting the battery box to the Test Unit), a 250V 5A, Type F fuse, a charge inhibit switch and twelve nickel metal hydride (NiMH) batteries, which are wired in series to provide a nominal 14.4 vdc. As an alternate, non-rechargeable type D dry cell batteries may be used in place of the NiMH batteries; however, the current capacity reduces the time the test unit can be powered. The following provides a description of the battery pack components:  Accessory kit – The accessory kit is comprised of the accessory bag, dc input cable, RS-232 cable and the ac/dc PCP.  Adapter Cables Assemblies – The aircraft adapter cable assemblies provide a means of connecting the FCTS to individual aircraft fuel systems or fuel system components (bench test). Instructions for connecting and using the adapters and boxes are provided by the “on screen” menus of the FCTS. The peculiar adapter cable assemblies will be packaged in their own transit case and shipped along with the FCTS. TTU-597/E Operation The FCTS can be used on aircraft with ac or dc capacitance fuel systems to measure capacitance in the range 0 to 30,000pF. The test unit is equipped with a battery management system which monitors the presence of an installed battery pack. Depending on level of charge, it requires about 1 1/2 to 2 hours to fully charge the battery pack. Fully charged NiMH batteries will provide operation for more than 4 hours of continuous use. The FCTS is designed to operate between a temperature range of -40° to +55° C and in an environment normal to carrier flight decks. The unit has adequate shielding and filtering to withstand hazardous electromagnetic fields, both conducted and radiated. The FCTS provides the following functions:  AC Capacitance Measurement – The method used is to apply a known voltage at a fixed frequency to the capacitor under test and measure the current flowing through the capacitor. This current is directly proportional to the value of the capacitor. The capacitor under test is fed from a low impedance source which eliminates the effect of any stray capacitance in the system. The measurement 2-80

signal out of the capacitor is fed into a virtual earth amplifier which, again, eliminates the effects of stray capacitance. This signal is then fed to a synchronous rectifier which also acts as a narrow band digital filter. The output from the synchronous rectifier is amplified and processed through a high resolution Analog-to-Digital Converter (ADC). The ADC converts the signal to a digital format which can be processed by the test unit software to a true capacitance value which is displayed on the screen. For stability reasons, the measurement electronics are calibrated against a very accurate internal capacitor on a continual basis to alleviate any uncertainties due to drift.  DC Capacitance – The method used for dc capacitance is very similar to the ac method. An ac voltage is applied to the capacitor under test, but in this case there are two diodes in series with the capacitor arranged so that the positive half cycles of the ac voltage pass through one diode and the negative half cycles pass through the other diode. In the case of negative dc c apacitance probes, the positive current goes to ground and the negative current goes into an integrating virtual earth amplifier whose output is a dc voltage directly proportional to the value of the capacitance. This voltage is fed to the same ADC as in the ac case with the same results. The positive dc capacitance probe case is the mirror image of the negative dc case.  Resistance Measurement – Resistance measurement is achieved by using three ranges. The lowest range from 0.1 ohms to 100 ohms uses a constant current method where the voltage across the unknown resistance is measured and the resistance calculated from knowledge of the applied current. The other two ranges use a method whereby a voltage is applied to the resistance under test via a known resistance and the voltage at the junction of these resistors measured, from this voltage the unknown resistor is calculated. In all three cases a 16 bit ADC is used for the measurement.  Voltage Measurement – Voltage measurement is achieved in the FCTS by dividing down an input voltage of ± 50 V to give 0 ± 10 V into the 16 bit bipolar ADC via a buffering op-amp of unity gain.  Distance To Fault – Distance to fault measurement is achieved by measuring the capacitance to ground of the Lo Z wire which is a shielded lead from the aircraft. Knowing the capacitance per foot of this lead and entering it into a display window on the tester, the resulting capacitance measurement will be displayed as the distance to fault in feet.  Capacitance Simulation – Capacitance simulation is achieved by using a bank of very accurate and stable capacitors arranged parallel and selected under software control to provide 11 to 10,000pF for TANK and 11 to 1000pF for COMP. The nominal value of 11pF is the result of the stray capacitances experienced within the wiring of the FCTS. The TANK and COMP capacitance simulators are independent and thus may be used simultaneously. TTU-597/E Operating Modes Power-On Self- Test (POST) – When the test unit is first powered up, a POST is conducted during the start-up sequence. Following a successful POST, the test unit performs self-calibration after which the user is prompted to choose either auto or manual testing. 2-81

Figure 2-38 — Angle-of-Attack Test Set AN/PSM-17A. Automated Mode – In the automated mode, the user can choose to perform a test on a selected aircraft fuel system or on an aircraft fuel system component (bench test). On screen instructions will guide the operator through the testing procedure. The on screen instructions list associated equipment and cables required prior to displaying the set-up procedure. Testing will not be initiated until the set-up procedure has been accomplished. Manual Mode – The manual mode allows the test unit to be used directly to measure insulation resistance, bonding or capacitance. A capacitor simulation output can also be provided. Angle-Of-Attack Test Set AN/PSM-17A, Part Number SLZ9060 The angle-of-attack test set lets you test, adjust, calibrate, simulate, and monitor the angle-of-attack indicating system. The test set also provides a means for you to test the aircraft approach lights, cockpit index lights, and the stall warning system. The test set (Figure 2-38) consists of a control panel enclosed in a case. It also includes the cables and components to interconnect and test the angle-of-attack system and associated components without their removal from the aircraft. The control panel contains a microammeter, a differential pressure gauge, a potentiometer control, a bellows assembly, indicator lamps, and electrical connectors. Also, the panel has various toggle and rotary switches to select and control circuits that are within the test set. A radiometer system makes up the largest portion of the test set. This system consists of the nullmeter, the SIM OR NULL potentiometer control, and the NULL switch. It lets you test the angle-of-attack (AOA) transmitter by placing switches in various positions and using the potentiometer dial to simulate known inputs to the indicators. The pressure system consists of a differential pressure gauge, test pressure connector, bellows assembly, pressure control, and surge chamber (test set case). It dynamically tests the angle-of-attack transmitters. An air pressure or vacuum transmits through a hose to parts of the AOA transmitter probe by positioning the control on the bellows assembly. This slight pressure causes the probe to rotate. Thus, the pressure system of the test set simulates conditions corresponding to various aircraft angles of attack.

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Figure 2-39 — Air-conditioning test set AN/PSM-21A. A series of indicator lamps (three indexer, three approach, and a stall warning) are on the test set control panel. They simulate the action of the aircraft indexer and approach lights and the stall warning vibrator. Two additional lamps show when the test set has ac and dc power. The power requirements for the test set are 28-volt dc and 110-to 120-volt, 400 hertz, single-phase ac. Air-Conditioning Test Set AN/PSM-21A The PSM-21A (Figure 2-39) is for flight line checkout and troubleshooting of electrical components in the cabin, pressure suit, and equipment air-conditioning systems. To accomplish checkout of these systems, apply external electrical power to the aircraft. Checkout of the system under test involves simulation of sensor and limiter inputs by the test set. You can connect external test equipment to test points on the test set to measure resistance, voltage, or waveforms to determine system operation. You can also determine system operation by visual monitoring. For example, with a known electrical input into the air-conditioning system, air- conditioning valves should move to a known position. Temperature Control System Test Set BR61-103 The Temperature Control System (TCS) Test Set, BR61-103 (Figure 2-40) consists of a panel of integrally mounted instruments and components, a cable assembly, a placard and self-test assembly all contained within a combination case. The combination case is divided into two sections connected by quick disconnect fasteners. The panel’s

components are used to operate, simulate, and monitor the aircraft cabin or flight station temperature control system. The removable lid of the combination case provides storage for the cable assembly, placard, and self- test assembly. The cable assembly is used to interface the test set with the self- test assembly and the aircraft cabin or flight station temperature control station. The test set is sixteen inches in length, fifteen inches in height, and fourteen inches wide and weighs thirty-four pounds. Power requirements are 115 vac, 400Hz. The test set is used for flight line checkout and troubleshooting of electrical components in equipment temperature control systems. To accomplish checkout of these systems, apply external electrical power to the aircraft. 2-83

Figure 2-40 — Temperature Control System (TCS) test set, BR61-103.

Modes of Operation Temperature control of the TCS can be attained either manually or automatically. With the test set SELECTOR INDICATOR switch set to ACFT or SIM, the desired temperature and mode of operation are chosen through use of a push-pull knob on either the TCS or test set selector indicator dial. An annunciator flag on the dial displays either MAN or AUTO according to the selection. The MODE MAN or MODE AUTO indicator comes on to further display and verify the selected mode. The desired temperature is selected by setting the index tab on the periphery of the dial which is positioned by rotating the push-pull knob. A pointer at the center of the dial is a program position indicator (PPI) that indirectly indicates the TCS A, B, and C valve positions are programmed to achieve or maintain the selected temperature. For example, if the pointer should be aligned to the two-dot position, the temperature would be approximately 23.3 ° C - program voltage would be 5 volts, and the related nominal degree of valve positions would be: A valve, 40 degrees; B valve, 80 degrees; and C valve, 20 degrees. The actual program voltage is displayed on the PROGRAM VOLTAGE meter and the VALVE POSITION A, B AND C meters indicate each of the TCS valve positions. For complete instructions on this test set, you should refer to the current maintenance instruction manual, AG-513SA-OMP-000. 2-84

ELECTROSTATIC DISCHARGE (ESD) The sensitivity of electronic devices and components to ESD has recently become clear through use, testing and failure analysis. The construction and design features of current microtechnology have resulted in devices being destroyed or damaged by ESD voltages as low as 20 volts. The trend in this technology is toward greater complexity, increased packaging density, and thinner dielectrics between active elements. This trend will result in devices even more sensitive to ESD. Various devices and components are susceptible to damage by electrostatic voltage levels commonly generated in production, test, and operation, and by maintenance personnel. These devices and components include the following:  All microelectronic and most semiconductor devices, except for various power diodes and transistors  Thick and thin film resistors, chips and hybrid devices, and crystals All subassemblies, assemblies, and equipment containing these components/devices without adequate protective circuitry are ESD sensitive (ESDS). You can protect ESDS items by implementing simple, low-cost ESD controls. Lack of implementation has resulted in high repair costs, excessive equipment downtime, and reduced equipment effectiveness. The operational characteristics of a system may not normally show these failures. However, under internal built-in-test monitoring in a digital application, they become pronounced, For example, the system functions normally on the ground, but, when placed in an operational environment, a damaged PN junction might further degrade, causing its failure. Normal examination of these parts will not detect the damage unless you use a curve tracer to measure the signal rise and fall times, or check the parts for reverse leakage current. Static Electricity Static electricity is electrical energy at rest. Some substances readily give up electrons while others accumulate excessive electrons. When two substances are rubbed together, separated or flow relative to one another, such as a gas or liquid over a solid, one substance becomes negatively charged and the other positively charged. An electrostatic field or lines of force emanate between a charged object to an object at a different electrostatic potential, such as one with more or less electrons, or ground. Objects entering this field will receive a charge by induction. The capacitance of the charged object relative to another object or ground also has an effect on the field. If the capacitance is reduced, there is an inverse linear increase in voltage, since the charge must be conserved. As the capacitance decreases, the voltage increases until a discharge occurs via an arc.

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Causes of Static Electricity Generation of static electricity on an object by rubbing is known as the triboelectric effect. Table 2-7 lists substances in the triboelectric series. The size of an electrostatic charge on two different materials is proportional to the separation of the two materials. Typical prime charge generators commonly encountered in a manufacturing facility are shown in Table 2-8. Electrostatic voltage levels generated by nonconductors can be extremely high. However, air will slowly dissipate the charge to a nearby conductor or ground. The more moisture in the air the faster a charge will dissipate. Table 2-9 shows typical measured charges generated by personnel in a manufacturing facility. Note the decrease in generated voltage with the increase in humidity levels of the surrounding air.

Table 2-7 — Triboelectric Series POSITIVE (+) ACETATE GLASS HUMAN HAIR NYLON WOOL FUR ALUMINUM POLYESTER PAPER (Small positive charge) COTTON (No charge) 0 STEEL (No Charge) WOOD (Small negative charge) ACETATE FIBER NICKEL, COPPER, & SILVER BRASS & STAINLESS STEEL RUBBER ACRYLIC POLYSTYRENE FOAM POLYURETHANE FOAM SARAN POLYETHYLENE POLYPROPYLENE PVC (VINYL) KEL-F TEFLON NEGATIVE (-) NOTE The triboelectric series is arranged in an order that when any two substances in the list contact one another and are separated, the substance higher on the list assumes a positive charge. 2-86

Table 2-8 — Typical Charge Generators Work Surfaces Formica (waxed or highly polished) Finished wood Synthetic mats Floors Wax finished Vinyl Clothes Common clean room smocks Personnel garments (all textiles except virgin cotton) Nonconductive shoes Chairs Finished wood Vinyl Fiberglass Packaging & Handling Common polyethylene (bags, wraps, envelopes, etc.) Common bubble pack, foam Common plastic trays, plastic tote Boxes & vials Assembly, Cleaning, Test, & Repair Spray cleaners Common solder suckers Common solder irons Solvent brushing (synthetic bristles) Cleaning & drying Temperature chambers Table 2-9 — Typical Measured Electrostatic Voltages Means of Static Generation Voltage Levels @ Relative Humidity Low (10-20%) High (65-90%) Walking across carpet 35,000 1,500 Walking over vinyl floor 12,000 250 Worker at bench 6,000 100 Vinyl envelopes for work instructions 7,000 600 Common poly bag picked up from bench 20,000 1,200 Work chair padded with urethane foam 18,000 1,500 2-87

Effects of Static Electricity The effects of ESD are not recognized. Failures due to ESD are often misanalysed as being caused by electrical overstress due to transients other than static. Many failures, often classified as other, random, unknown, infant mortality, manufacturing defect, etc., are actually caused by ESD. Misclassification of the defect is often caused by not performing failure analysis to the proper depth. Component Susceptibility All solid-state devices (all microcircuits and most semiconductors), except for various power transistors and diodes, are susceptible to damage by discharging electrostatic voltages. The discharge may occur across their terminals or through subjecting these devices to electrostatic fields. Latent Failure Mechanisms The ESD overstress can produce a dielectric breakdown of a self-healing nature when the current is unlimited. When this occurs, the device may retest good but contain a hole in the gate oxide. With use, metal will eventually migrate through the puncture, resulting in a shorting of this oxide layer. Another structure mechanism involves highly limited current dielectric breakdown from which no apparent damage is done. However, this reduces the voltage at which subsequent breakdown occurs to as low as one-third of the original breakdown value. ESD damage can result in a lowered damage threshold at which a subsequent lower voltage ESD will cause further degradation or a functional failure. ESD Elimination The heart of an ESD control program is the ESD-protected work area and ESD- grounded work station. When you handle an ESD-sensitive device outside of its ESD protective packaging, provide a means to reduce generated electrostatic voltages below the levels at which the item is sensitive. The greater the margin between the level at which the generated voltages are limited and the ESDS item sensitivity level, the greater the probability of protecting that item. Prime Generators All common plastics and other prime generators of static electricity should be prohibited in the ESD protected work area. Carpeting should also be prohibited. If you must use carpet, it should be of a permanently anti-static type. Perform weekly static voltage monitoring where carpeting is in use.

Personal Apparel and Grounding An essential part of the ESD program is grounding personnel and their apparel when handling ESDS material. Means of doing this are described in this section.

CAUTION Anti-static cushioning material is acceptable; however, the items cited shall be of conductive material to prevent damage or destruction of ESDS devices. 2-88

Smocks Personnel handling ESDS items should wear long sleeve ESD protective smocks, short sleeve shirts or blouses, and ESD protective gauntlets banded to the bare wrist and extending toward the elbow. If these items are not available, use other anti-static material (such as cotton) that will cover sections of the body that could contact an ESDS item during handling. Personnel Ground Straps Personnel ground straps should have a minimum resistance of 250,000 ohms. Based upon limiting leakage currents to personnel to 5 mill amperes, this resistance will protect personnel from shock from voltages up to 125 volts RMS. The wrist, leg, or ankle bracelet end of the ground strap should have some metal contact with the skin. Bracelets made completely of carbon impregnated plastic may burnish around the area in contact with the skin, resulting in too high impedance to ground. ESD Protective Materials There are two basic types of ESD protective materials: 1. Conductive – Conductive materials protect ESD devices from static discharges and electromagnetic fields. 2. Anti-static – Anti-static material is a static inhibiting material. Other than not generating static, anti-static material offers no other protection to an ESD device. Conductive ESD Protective Materials Conductive ESD protective materials consist of metal, metal-coated, and metal- impregnated materials (such as carbon particle impregnated, conductive mesh or wire encased in plastic). The most common conductive materials used for ESD protection are steel, aluminum, and carbon-impregnated polyethylene and nylon. The latter two are opaque, black, flexible, heat sealable, electrically conductive plastics. These plastics are composed of carbon particles, impregnated in the plastic, which provides volume conductivity throughout the material. Anti-Static ESD Protective Materials Anti-static materials are normally plastic-type materials, such as polyethylene, polyolefin, polyurethane, nylon, which are impregnated with an anti-static substance. This anti-static substance migrates to the surface and combines with the humidity in the air to form a conductive sweat layer on the surface. This layer is invisible and, although highly resistive, is amply conductive to prevent the buildup of electrostatic charges by triboelectric, or rubbing methods in normal handling. Simply stated, the primary asset of an anti-static material is that it will not generate a charge on its surface. However, this material won’t protect an enclosed ESD device if it comes into contact with a charged surface. This material is of a pink tint, a symbol of its being anti-static. Anti-static materials are for inner-wrap packaging. However, anti-static trays, vials, carriers, boxes, etc., are not used unless components and/or assemblies are wrapped in conductive packaging. 2-89

Hybrid ESD Protective Bags Lamination of different ESD protective material (that is, conductive and anti-static) is available that provides the advantage of both in a single bag. ESDS Device Handling The following are general guidelines applicable to the handling of ESDS devices:  Make sure that all containers, tools, test equipment, and fixtures used in ESD protective areas are grounded before and/or during use, either directly or by contact with a grounded surface.  Personnel handling ESDS items must avoid physical activities in the vicinity of ESDS items that are friction-producing, for example, removing or putting on smocks, wiping feet, sliding objects over surfaces, etc.  Personnel handling ESDS items must wear cotton smocks and/or other anti- statically treated clothing.  Avoid the use or presence of plastics, synthetic textiles, rubber, finished wood, vinyl, and other static-generating materials (Table 2-9) where ESDS items are handled out of their ESD protective packaging.  Place the ESD protective material containing the ESD item on a grounded work bench surface to remove any charge before opening the packaging material.  Personnel must attach personnel grounding straps to ground themselves before removing ESDS items from their protective packaging.  Remove ESDS items from ESD protective packaging with fingers or metal grasping tool only after grounding and place on the ESD grounded work bench surface.  Make periodic electrostatic measurements at all ESD protected areas. This assures the ESD protective properties of the work station and all equipment contained there have not degraded.  Perform periodic continuity checks of personnel ground straps (between skin contact and ground connection), ESD grounded work station surfaces, conductive floor mats, and other connections to ground. Perform this check with a megohmmeter to make sure grounding resistivity requirements are met. ESDS Device Packaging Before an ESDS item leaves an ESD-protected area, ensure the following precautions are taken:  Ensure shorting bars, clips, or non-corrective conductive materials are correctly inserted in or on all terminals or connectors.  Package ESD items in an inner wrap, of type II material conforming to MIL-B- 81705, and an outer wrap of type I material conforming to MIL-B-81705. You may use a laminated bag in instead of the above provided such meets the requirements of M-B-81705. Cushion-wrap the item with electrostatic free material conforming to PPP-C-1842, type III, style A. Place the cushioned item into a barrier bag fabricated from MIL-C-131 and heat-seal closed, method 1A-8. Place the wrapped, cushioned, or pouched ESDS item in bags conforming to MIL-B-117, type I, class F, style 1. 2-90

Figure 2-41 — ESDS markings.

NOTE Mark the packaged unit with the ESD symbol and caution. (Figure 2-41) 2-91

Testing/Repair Before you work on ESDS items, make sure the following precautions/procedures are met:  Be sure that work area, equipment, and wrist strap assembly have a proper ground.  Attach wrist strap and place metal tools, card extractors, test fixtures, etc., on grounded bench surface.  Place conductive container on the bench top. Remove component/assembly from packaging. Remove shorting devices, if present. Handle components by their bodies and lay them on conductive work surface or test fixtures.  Test through the connector or tabs only.  Do not probe assemblies with test equipment.  You must use a high input impedance meter such as a Fluke Multimeter to test parts or assemblies.  After testing, replace shorting devices and protective packaging.  Dielectric strength tests are not permitted.

2-92

End of Chapter 2 ELECTRICAL MAINTENANCE AND TROUBLESHOOTING Review Questions 2-1. What is one rule you should always follow when working on electric or electronic equipment?

A. Power source B. Safety first C. Location of equipment D. Required tools

2-2. What should you do to the power supply when working on an electrical system?

A. Close and tag the circuit breakers and main switches B. Open and tag the circuit breakers C. Open and tag the circuit breakers and main switches D. Open and tag the main switches

2-3. What is the manual that you should refer to for an illustration of the danger areas for your aircraft?

A. MIM B. IPB C. RTM D. NEETS

2-4. What exterior light is required to be on during engine operation?

A. Wing lights B. Cockpit light C. Tail light D. Anti-collision light

2-5. When working on a circuit, you need to have a person who is fully qualified in what standing by?

A. Search and rescue B. CPR C. The operation of equipment being tested D. Microminiature circuit repair

2-6. Before touching a capacitor, you should be sure it is discharged by doing what?

A. Turning off power B. Opening the circuit breaker C. Short-circuiting the terminals D. Install the capacitor safety shield 2-93

2-7. What is the preferred fire extinguishing agent for electrical fires?

A. PKP B. H20 C. AFFF D. CO2

2-8. What are the two basic rules you should follow when working with hand tools?

A. Use the proper tool and ensure it is properly checked out from tool issue B. Ensure it is properly checked out from tool issue C. Use the proper tool and keep it in working order D. Ensure it is properly checked out from tool issue and is the proper tool

2-9. List the steps that will make your troubleshooting easier?

A. Review system operation, analyze symptom, detect and isolate the trouble B. Review system operation, clean and inspect, detect and isolate the trouble C. Review system operation, analyze symptom, and perform clean/inspect D. Analyze symptom, clean and inspect, detect and isolate the problem

2-10. What test equipment is used when looking for shorts, grounds, opens, and wrong resistances?

A. Megger B. TDR C. Voltmeter D. Digital multimeter

2-11. What process should never be used as a troubleshooting method because it is expensive and does not always fix the fault?

A. Seeking advice from FRC B. Working overtime C. Parts replacement D. Component isolation

2-12. After removing a part, you should always look for small items from the work to ensure they do not become what type of a hazard?

A. Eye hazard B. Foreign Object Damage (FOD) C. Choking hazard D. Tripping hazard

2-13. Which of the following tools would help you analyze a system?

A. MIMs, schematics, and records on the equipment B. MIMs, schematics, and Illustration Parts Pubs C. Illustration Parts Pubs, and equipment records D. Illustration Parts Pubs, schematics, and equipment records

2-94

2-14. What is the first action to take when beginning to troubleshoot a piece of equipment?

A. Analyze the symptom B. Detect and isolate the trouble C. Perform resistance checks D. Review the system

2-15. What types of common faults interrupt power through a circuit?

A. Intact wiring, loose terminals, faulty relays, and faulty switches B. Broken wiring, loose terminals, faulty relays, and faulty switches C. Broken wiring, tight terminals, faulty relays, and faulty switches D. Intact wiring, tight terminals, faulty relays, and faulty switches

2-16. Which reference is referred to for approved maintenance procedures?

A. IPB B. NATOPS C. MIMs D. NAVEDTRA

2-17. Which reference is used for procedures when removing and replacing equipment?

A. MIMs B. NAVEDTRA C. NATOPS D. IPB

2-18. What is the purpose of testing a piece of equipment after it has been repaired?

A. Ensure technician skills are adequate B. Ensure future maintenance will be not required C. Ensure operators can properly operate equipment D. Ensure system works properly

2-19. How would you connect a digital multimeter to a circuit that is being tested for voltage?

A. Connect in series B. Connect in parallel C. Connect in parallel - series D. Connect in phase

2-20. When using a digital multimeter, you should start at what range?

A. Lowscale B. Lowscale - midscale C. Midscale D. Highscale

2-95

2-21. A circuit in which current no longer flows is known as a ________.

A. short circuit B. open circuit C. faulty relay D. broken switch

2-22. What test do you perform to find an open circuit?

A. Continuity test B. Resistance check C. Voltage check D. Ohm test

2-23. Blown fuses and open circuit breakers are usually an indication of what type of fault?

A. Open circuit B. Faulty relay C. Open switch D. Short circuit

2-24. List the two disadvantages of the volt- ohm-milliammeter (VOM)?

A. Can unload the circuit under test and damage meter movement B. Can load the circuit under test and damage meter movement C. Can load the circuit under test and damage equipment D. Can unload the circuit under test and damaged equipment

2-25. What type meter should you use to test for insulation breakdown?

A. Megger B. Ohmmeter C. Digital multimeter D. Ammeter

2-26. What test equipment shows you the wave shape of current or voltages?

A. Megger B. Multimeter C. Voltmeter D. Oscilloscope

2-27. What is a discontinuity?

A. Normal resistance or impedance that ensures normal signal flow B. Abnormal resistance or impedance that interferes with normal signal flow C. Abnormal resistance or impedance which in turn ensures normal signal flow D. Normal resistance or impedance that interferes with normal signal flow

2-96

2-28. What type of wire should you use to carry 600 to 1,000 volts with a temperature rating between 302°F to 500°F?

A. MIL-W-25038 B. MIL-W-22759 C. MIL-W-81044 D. MIL-C- 27500

2-29. What type of metal used in wire forms a resistant oxide film on all its surfaces?

A. Steel B. Copper C. Aluminum D. Brass

2-30. When stamping wire identification numbers, at what interval should you stamp the wire?

A. 15 inches B. 24 inches C. 30 inches D. Total length of wire divided by 4 (measured in inches)

2-31. In wire identification numbers, the suffix N means the wire completes the circuit to?

A. Photographic circuits B. Ground C. Loaded equipment D. Miscellaneous equipment

2-32. What is the most important item to consider when soldering?

A. Selection of the correct solder B. Selection of the correct power supply C. Selection of the correct flux D. Selection of the correct iron

2-33. What publication contains information on mounting hardware for aircraft parts?

A. NAVEDTRA 14308 B. NAVAIR 01- 1A-8 C. NAVAIR 01-1A-50 D. NATOPS 01- 1A

2-34. When substituting hardware, what should you consider before making the substitution?

A. Location B. Strength C. Design D. Manufacturer 2-97

2-35. What is the only reason for clamping a wire bundle to a plumbing line?

A. Separation of the wire bundle from the plumbing line B. Combining of the wire bundle with the plumbing line C. Combining of individual wires with the plumbing line D. Separation of individual wires from the plumbing line

2-36. List the three types of safety wire:

A. Lockwire, sheen wire, & seal wire B. Lockwire, shear wire, & seal wire C. Lockwire, shear wire, & seam wire D. Lockwire, sheen wire, & seam wire

2-37. What is the purpose of bonding?

A. To provide a low-resistance return path for single-wire electrical systems B. To pro vide a means of masking the entire aircraft to the earth’s potential C. To increase lightning damage to the aircraft and its occupants D. To provide for a decrease in grounding

2-38. List the manuals to which you should refer for the correct cleaning agents and procedures?

A. NATOPS 01- 1A-50 B. NATOPS 01- 1A-505 C. NAVAIR 01- 1A-8 D. NAVAIR 01- 1A-509

2-39. For a technician to repair a printed circuit board, what certification must they meet?

A. First aid B. Microminiature circuit repair C. CPR D. Electronic repair

2-40. What are the two types of test equipment?

A. Common and uncommon B. Common and unusual C. Common or particular D. Common or peculiar

2-41. What test equipment is used to test aircraft engine driven accessories, such as generators and generator drive systems?

A. AECTS B. UUT C. VSD D. TTU-27/E

2-98

2-42. The Huntron Tracker 1000 and 2000 are for use on circuit boards and systems with all voltage sources in what condition?

A. Power off B. Power on C. Standby D. Safe

2-43. What type of signal display does the Huntron Tracker 1000 and 2000 show when the signal fuse is open and the test leads shorted together?

A. Voltage phases B. Closed circuit signatures C. Open circuit signatures D. Voltage amplitude

2-44. When using the Huntron Tracker 2000, why must you make good contact with the test leads?

A. Test leads are conductive only at the tips B. Materials used for leads are not very conductive C. Test leads are conductive only at the base of stem D. Materials used for leads prevent transfer of current

2-45. Which piece of test equipment would you use to test an engine ignition unit output in sparks per second?

A. TTU-27/E B. JATCAL C. Jet Ignition System Tester D. Hunton Tracker 2000

2-46. When testing a tachometer generator, which switch is used to select the individual phase-to-phase voltage displayed on the voltmeter?

A. RPM x 2, RPM x 1, and PERCENT RPM B. A-B, B-C and A-C TERM CONNECTION C. LOAD IN OHMS D. TEST GEN INPUT

2-47. What component of the TTU-27/E is used to test and calibrate the tachometer indicator?

A. Drive pad B. Master indicator C. Master tachometer generator D. Jet engine tachometer generator

2-99

2-48. What is the minimum voltage for using the TAKCAL unit of the JETCAL Analyzer?

A. 5 volts B. 33 volts C. 75 volts D. 95 volts

2-49. What initial test should be performed when using the TTK-512 propeller synchronizer tester on the synchrophaser system?

A. Voltage tests B. Tests on ground with engines running C. Inflight tests D. Resistance and megger tests

2-50. What piece of equipment is used to test pitot-static systems?

A. TS-4508/U B. T TU-378 A/E C. AN/PSM- 17A D. AN/PSM- 21A

2-51. The TTU-378 A/E test set can be used to verify system operation, adjust the indicator to the tank units, or determine whether the ________ are defective.

A. unit insulation B. accessory cables C. indicator or probes D. insulation resistance

2-52. The FCTS is compatible with all types of aircraft which have what type fuel system?

A. Reactive B. Inductive C. Electromagnetic D. Capacitive

2-53. Which operating mode of the TTU-597/E allows the test unit to be used directly to measure insulation resistance, bonding or capacitance and provides a capacitor simulation output?

A Automated Mode B. Manual Mode C. Power -On Self-Test D. Capacitance Mode

2-100

2-54. The AN/PSM-17A test set is used to test what system?

A. Navigation B. Ignition C. Fuel D. Angle- of-attack

2-55. When a desired temperature is selected when operating the BR61-103, a pointer at the center of the dial is a program position indicator (PPI) that indicates what?

A. TCS A, B, and C valve positions are programmed to achieve or maintain the selected temperature. B. TCS A, and B valve positions must be programmed for selected temperature C. A valve, 40 degrees; B valve, 60 degrees; and C valve, 80 degrees. D. A valve, 40 degrees; B valve, 80 degrees; and C valve, 80 degrees.

2-56. What are the dimensions for the BR61-103 test set?

A. 16 inches in length, 15 inches in height, and 14 inches wide B. 17 inches in length, 14 inches in height, and 14 inches wide C. 16 inches in length, 15 inches in height, and 14 inches wide D. 16 inches in length, 15 inches in height, and 34 inches wide

2-57. What is the lowest voltage that will destroy or damage an ESD-sensitive device?

A. 10 volts B. 20 volts C. 30 volt D. 25 volts

2-58. The generation of static electricity on an object by rubbing is known as the ________.

A. electrostatic charge B. dielectric effect C. triboelectric effect D. prime charge

2-101

RATE TRAINING MANUAL – USER UPDATE CNATT makes every effort to keep their manuals up-to-date and free of technical errors. We appreciate your help in this process. If you have an idea for improving this manual, or if you find an error, a typographical mistake, or an inaccuracy in CNATT manuals, please write or email us, using this form or a photocopy. Be sure to include the exact chapter number, topic, detailed description, and correction, if applicable. Your input will be brought to the attention of the Technical Review Committee. Thank you for your assistance. Write: CNATT AV Rate Training Manager 230 Chevalier Field Avenue Pensacola, FL 32508 COMM: (850) 452-9700 utilize voice directory for AE/AT Rate Training Manager. DSN: 922-9700 utilize voice directory for AE/AT Rate Training Manager. E-mail: Refer to NKO AE rate training web page for curent contact information.

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2-102

CHAPTER 3 POWER GENERATION AND CONTROL SYSTEMS Aviation Electrician's Mates (AEs) operate and maintain various modern naval aircraft systems. As an AE, you must know the electric power systems of these aircraft. The electric power requirements and the electric system components of aircraft vary widely according to the size and application of the aircraft. You must understand the component parts of the electrical systems and the power distribution systems of modern naval aircraft. Alternating Current (ac) generators supply the electrical energy for operating aircraft avionics equipment. A generator is a machine that converts mechanical energy into electrical energy by electromagnetic induction. Navy Electricity and Electronics Training Series (NEETS), Module 5, Introduction to Generators and Motors, NAVEDTRA 14177 contains a detailed discussion on generator theory. You should study this module and refer to it during your study of this chapter. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Identify various electrical power sources used on aircraft and recognize their functions, construction, and operating characteristics. 2. Explain the reasons for and means of protecting electrical circuits through use of voltage and frequency control circuits and associated circuits. AIRCRAFT ELECTRICAL POWER SOURCES AC Generators Modern ac power systems provide for better aircraft equipment design and use. Older electronic equipment, powered by direct current, had an inverter for ac power and a dynamotor for supplying higher voltage Direct Current (dc) power. These components are very heavy compared to their relative power outputs. They are not reliable, and they increase maintenance. The same ac-powered equipment obtains various ac voltages and dc power by using simple transformers and transformer-rectifiers. These components are lightweight, simple, and reliable devices. Modern naval aircraft use the three phase, 120-/208-volt, 400-hertz ac power system in order to meet increasing aircraft power requirements. The number of magnetic poles and rotor Revolutions per Minute (RPM) determines the voltage frequency of the generator. Constant frequency requires constant rotor RPM when the number of poles are a fixed quantity. The ac generator rotating field has 12 poles with adjacent poles being of opposite polarity. Each pair of poles produces one cycle per revolution; therefore, each revolution produces six cycles. The output frequency of the generator varies in direct proportion to the engine drive speed. A generator operating at 6,000 RPM is operating at 100 revolutions per second or at 600 hertz. NEETS Module 5 contains a detailed discussion of frequency.

3-1

Figure 3-1 — Three-phase ac generator output. The 120-/208-volt, 400-hertz, three-phase, ac power system has many advantages over the 28-volt dc system. It requires less current than the 28-volt dc system because of h igher voltage and a ground neutral system. The current required is a fraction of that required for the same power in a 28-volt dc system. This permits the use of smaller aircraft wiring, saving weight. The ac generator and many of the system's control and protection components are lighter. Twelve kilowatts is the practical limit to the size of an aircraft dc generator. Aircraft now have ac generators with ratings up to 90 kilovolt ampere (kVa). Types of AC Generators Aircraft ac generators range in size from the tachometer instrument generator up to the 90,000 volt-ampere generators. Regardless of weight, shape, or rating, practically all of these generators have the following common characteristics:  The stator (stationary armature winding) provides the ac output.  The ac generator field (rotor) is a rotating magnetic field with fixed polarity.  Regulating the RPM of the rotating magnetic field controls the voltage frequency.  Controlling the strength of the magnetic field is the method of voltage regulation. Present military specifications require that the basic aircraft ac power system produces voltage with a value of 120 and 208 volts. A three-phase generator is actually three separate power sources enclosed in one housing (Figure 3-1, view A). External connections form a wye (Figure 3-1, view B) to produce the required 120-/208-volt output. Each output winding produces 120 volts as measured from n to a, b, or c (phase voltage). The voltage is 1.73 times the single-phase voltage when measuring two separate phase voltages together (line voltage).

3-2

Figure 3-2 — Brush-type, three-phase ac generator. The line voltage found in a three-phase, wye­connected system is the vector sum of the voltages generated by two separate phase windings. Voltages reach their peak amplitudes at different times because a 120-degree phase difference exists between them. Due to this phase difference they must be added vectorially and not directly. In the four-wire, grounded-neutral, wye-connected system, the neutral wire attaches to the frame of the aircraft (ground). The three-phase wires run to buses, which supply power to various loads. The connections for loads requiring 120 volts are between one of the buses and the aircraft frame. The load connections requiring 208 volts are between two of the buses (phases). BRUSH TYPE – Figure 3-2 shows a brush type ac generator. It consists of an ac generator and a smaller dc exciter generator as one unit. The output of the generator su pplies ac to the load. The only purpose for the dc exciter generator is to supply the direct current required to maintain the ac generator field. Figure 3-2, view B, is a simplified schematic of the generator.

Refer to Figure 3-2 as you read this section. The exciter is a dc, shunt-wound, self- e xcited generator. The exciter field (2) creates an area of intense magnetic flux between its poles. Voltage is induced in the exciter armature windings when the exciter armature (3) rotates in the exciter-field flux. The output from the exciter commutator (4) flows through brushes and slip rings (5) to the generator field. Having already been converted 3-3

Figure 3-3 — Disassembled brushless ac generator. by the exciter commutator, the dc current always flows in one direction through the generator field (6). Thus, a fixed-polarity magnetic field is maintained in the generator field windings. When the field winding rotates, its magnetic flux passes through and a cross the generator armature windings (7). The ac in the ac generator armature windings flows through fixed terminals to the ac load. The stationary member of the generator consists of the ac armature and the dc exciter field. Both ac and exciter terminal boards are easily accessible. All brush rigging is on the generator and has a brush cover. The slotted-hole mounting provides for ease in attaching to the engine pad. The capacitors connected between the exciter armature terminals and ground suppresses radio noise. BRUSHLESS TYPE – Most naval aircraft are using brushless generators for voltage generation. The advantage of a brushless generator over a brush type is its increased reliability and the greater operating time between overhaul. Figure 3-3 is an expanded view of the main assembly of a brushless generator. It shows those items that you will find important. The brushless generator shown in Figure 3-3 is a salient eight-pole, 6,000 RPM, ac generator. It has a 12-pole ac exciter and a three-phase, half­wave diode rectifier rotating with the exciter armature and main generator field assembly. The exciter rotor is a hollow frame assembly with the main ac field mounted on the inside and connected to a common drive shaft. A single-phase Permanent Magnet Generator (PMG) furnishes control voltage and power for the voltage regulator. Three, half-wave rectifiers are on the exciter rotor and connected to the exciter armature windings. A generator shaft shear section prevents possible damage to the engine or drive unit if the generator seizes. A fan at the drive end of the generator provides cooling airflow for the rotor and stator windings and the drive bearings.

3-4

Figure 3-4 — Sectional schematic of a brushless ac generator. Some aircraft have oil-cooled generators. The oil enters the generator through an inlet port and leaves through an exit port in the mounting flange of the generator. As the oil passes through the generator, it absorbs the heat from the rotor and stator. At the same time, it cools the rotating seals and lubricates and cools the bearings. Aircraft engine oil cools the generator and is used for the constant speed drive operation. As the generator shaft rotates (Figure 3-4), the PMG supplies single-phase, ac voltage to the voltage regulator and other protective circuits. PMG power is rect ified and supplied to the exciter field. The electromagnetic field, built by the excitation current flowing in the exciter, induces current flow in the rotating three-phase exciter rotor. This current is half-wave rectified by rotating rectifiers. The resultant dc goes to the rotating field winding of the ac generator. The rotating electromagnetic field induces ac voltage in the three-phase, wye- connected, output winding of the generator stator. Varying the strength of exciter stationary field accomplishes voltage regulation. Brushes within the generator aren't required when an integral ac exciter is used.

The absence of brushes minimizes radio noise in other avionics equipment. Two, three-phase differential transformers provide protection against shorts in the feeder lines b etween the generator and the bus (called feeder fault). One transformer is on the generator (Figure 3-4). Its coils sense the current flow through each of the legs that connect the ground side of the generator stator to ground. The other transformer is at the main bus and senses current flow through the three feeder lines. A short in the feeder line would cause the transformers sensing a difference in current to trip the generator off line. A generator mechanical failure warning device is incorporated in the generator. It consists of a soft copper strip embedded in and insulated from the generator stator assembly. A bearing beginning to fail allows the rotor to rub against the copper strip, co mpleting a warning light circuit to ground. 3-5

Prime Movers A prime mover is a device, such as an aircraft engine, that provides the driving force for a generator. Early attempts to control the rotor speed of ac generators using variable- pitch propellers or slipping clutches were unsuccessful, and ac generator power was of variable frequency. If a constant frequency power for the equipment were available, savings in weight and improved performance would be possible. From the weight and performance standpoint, as power requirements grew, it became necessary to furnish ac power at a constant frequency. The constant frequency was found in a hydromechanical, Constant Speed Drive unit (CSD). Other constant speed units are air or gas turbines and the constant RPM turboprop engine. The air turbine gets its air supply by using bleed air from the jet engine compressor or from a separate compressor. The hydromechanical, CSD unit converts variable engine speed to a constant speed output. It holds the frequency steady, to within a few hertz, of the desired 400 hertz. Load and fault transients limits are within a 380- to 420-hertz range. Air or gas-turbine drives are somewhat smoother in operation and hold steady-state frequencies to within ± 10 hertz. The constant RPM characteristic of the turboprop engine gives good frequency stability to the ac generator output. The propeller mechanical governor will hold the generator frequency to 400 ± 4 hertz. Inverters Inverters are an emergency source of ac power when normal ac power fails. The backup system of the F/A-18 aircraft is an example of this type. The standby attitude indicator receives power from the right, 115-volt ac bus. If the aircraft's generators fail to supply power to this bus, the standby attitude indicator receives power from an inverter. Because of a wide variety of inverters in use on aircraft, only one is discussed in this Rate Training Manual. Inverters consist of a speed-governed dc motor, an armature and brush assembly, and a permanent magnet inductor-type ac generator in one unit. The armature and the permanent magnet rotor mount on a common shaft. The standard inverter is a 120-volt, three-phase, four-wire, 400-hertz ac system. The four-wire system is better than the three-wire system. It allows a greater choice of single-phase circuits, improves phase load balance, decreases vulnerability to power failure, and gives better frequency and voltage control. The dc armature and the ac generating field windings are on the same rotor shaft in most inverters. The dc motor field and generator output (armature) windings are on the stator. A control box on the inverter contains the necessary devices to control the inverter's operation. These devices consist of the operating relays, voltage regulator and rectifier, filtering units, and smaller circuit components. The dc motor of most aircraft inverters is essentially a shunt-wound motor. High starting currents and a low rate of acceleration (because of low torque at starting) are characteristic of shunt­ wound inverters. The larger inverters have a series-starting winding to help avoid the effects of these undesirable characteristics. When the inverter reaches rated speed relays disconnect the series-starting winding and connect the dc input to the dc motor armature and the shunt winding. Then, the inverter operates as a shunt-wound motor having desirable constant-speed characteristics. Others use small compensating and commutating pole windings in series with the motor armature. These windings have no effect on the shunt-motor action. 3-6

Figure 3-5 — Typical aircraft inverter. The dc motor converts electrical energy into mechanical energy to drive the generator. The dc load current drawn by the motor depends on the ac load on the generator. A speed governor, or pulsating dc current through the field windings, controls motor speed. A solid-state, ON-OFF switching circuit provides the pulsating direct current. The speed of a dc motor is inversely proportional to the strength of the field. Therefore, as the motor speeds up, more current flows in the shunt windings, reducing the speed. Less current flows in the shunt-field windings and the motor speeds up when the motor speed falls below its normal value. The generator ac voltage is proportional to the speed of the rotor and the strength of the generator rotor field flux. The controlled frequency of the ac output is usually 400Hz. This frequency is a function of the number of poles in the generator field and the speed of the motor. The number of sets of generator stator windings determines the number of independent voltages, or phases, in the output. Some inverters supply both three-phase and single-phase outputs. Figure 3-5 shows a typical inverter.

The rating of aircraft inverters varies, depending on the equipment that it supplies. For e xample, an aircraft may carry a number of inverters. One may supply 120-volt, three­ phase ac to an essential bus during emergencies. Another supplies 120 -volt, single- phase ac power while another furnishes 120-volt, three-phase power to a specified bus or equipment. Figure 3-6 shows a cutaway view of an inverter. Controlling the dc excitation current in the generators rotating field maintains the inverter o utput voltage at a constant value. Demand variations on the inverter output determine the strength of the dc rotating field.

3-7

Figure 3-6 — Cutaway view of an E1616-2 inverter. Inverters operate on the same electrical principles as dc motors and ac generators. NEETS, Module 5, discusses these principles. For more information on the theory of voltage generation and regulation, you should refer to the NEETS, Module 2, Introduction to Alternating Current and Transformers, NAVEDTRA 14174 and Module 5. For detailed information about a particular inverter, refer to the manuals covering that inverter. Transformers A transformer, by itself, is not a true electrical power source. A-true electrical power so urce can produce electrical energy from another type of energy, such as chemical or mechanical. Transformers take electrical energy in the form of ac voltage and convert it to a different usable ac voltage. If you feel you need to study transformer construction and theory, you should refer to the NEETS, Module 2, before continuing this chapter. Transformer-Rectifiers Currently ac-powered equipment is more efficient than larger, heavier dc-powered e quipment. So ac generators now power naval aircraft, but dc power is needed for lighting and for controlling ac-powered equipment. The most common device now used to provide the necessary dc voltage is the Transformer Rectifier (TR). TRs have no moving parts, other than a cooling fan. They provide high reliability and ruggedness unmatched by most other avionics equipment. A separate voltage regulator 3-8

Figure 3-7 — Schematic diagram of a typical transformer rectifier. is not necessary so long as the ac input voltage maintains reasonable limits. The dc current capability is high and is largely dependent on the cooling available. Figure 3-7 shows an electrical schematic of a typical transformer-rectifier. You should refer to it as you read this section. It requires a 120-/208-volt, three-phase, four-wire input at 400 hertz. It has an output capability of 200 amperes at 25.5 to 29.5 volts.

The input ac voltage enters through pins B, F, and I and enters a Radio Frequency (RF) filter. The filter reduces noise interference to other avionics equipment in the aircraft. Power then connects with the wye-connected primary of a step-down power tran sformer. The ac output of the wye, delta-connected secondaries is rectified by diodes CR1 through CR12. The output goes through interphase transformer T2, and a f ilter network consisting of L2, C4, and C5 to the load. Interphase transformer T2 has an adjustable center tap to balance the two delta transformers for equal current output. The f ilter network reduces the 4,800-hertz ripple voltage to nearly straight-line dc voltage. Fan motor B1 is connected in parallel with the power transformer primary. This fan m otor is essential to proper operation and provides the only moving parts of the TR. Thermostat (S1) provides detection of excessively high temperatures and, in conjunction with external circuits, turns on an overheat warning light and automatically disconnects the input. 3-9

Figure 3-8 — A 2:1 ratio autotransformer. Autotransformers The autotransformer is like an ordinary transformer, except that it has one winding that is common to both primary and secondary. Within the limits of its application, it offers savings in both size and cost over conventional units. These savings are greatest when the turn ratio is less than 2 to 1 (either step-up or step- down). Savings diminish to insignificance when the turn ratio increases beyond 8 to 10. There is no isolation between primary and secondary positions of the circuit, a feature that is sometimes objectionable. Figure 3-8 shows a 2:1 step-down autotransformer circuit. Refer to this figure as you read this section. The tap at point B divides the winding into two equal parts. With a load of 5 ohms connected as shown, compute the load current using the formula I =

or

= 10 amperes. The power in the load equals EI (50 x 10) or 500 watts. Just like a regular transformer, this power comes from the primary by the magnetic field. Disregarding losses, the primary must take 500 watts from the line. Therefore, the primary current would be

(

) or 5 amperes. Only the difference between these two currents, 5 amperes, flows in the common portion C to B (shown by the arrow). The current in both sections of the winding is the same when the turn ratio is 2:1. This saves the cost and weight of an entire winding.

3-10

Figure 3-9 — A 1.33:1 ratio autotransformer. The autotransformer in Figure 3-9 has a turn ratio of 1.33:1. It connects to a load that draws 20 amperes. This represents a secondary power of EI = (90 x 20) or 1,800 watts. The primary current, neglecting losses, equals

(

) or 15 amperes. The current in the winding from B to C common to both circuits is the difference between the primary and secondary line currents, or 5 amperes. The saving here is obvious. A conventional transformer with the same characteristics requires a 120- volt, 15-ampere primary and a separate 90-volt, 20-ampere secondary. Here, the requirement is a 30-volt, 15-ampere winding in series with a 90-volt, 5-ampere winding. Thus, a 0.45-kVA auto­transformer supplies the 1.8-kVA load. There are many interesting uses for autotransformers. An autotransformer with a continuous variable tap is sold under the name VARIAC. It is used for many purposes where a continuous control from zero to full (or even above) line voltage is necessary. In this case, the core is toroidal (ring shaped). The winding is usually in the form of a single layer covering almost the entire surface. A control shaft carries an arm and a brush that makes contact with each turn of the winding as the shaft rotates. The setting of the shaft determines the turn ratio. One end of the winding goes to both line and load, and the other end goes to the line. The brush connects to the other side of the load. To obtain voltages higher than line voltage, the primary connects to a tap about 10 percent down from the end of the winding. (Voltages higher than line voltage compensate for a bnormally low line voltage.) This provides secondary control from zero to full line voltage, even though the actual line voltage is as much as 10 percent below normal. Instrument Transformers Usually, meters are not connected directly to high-voltage and high-current ac circuits. Instrument transformers connect meters to these circuits. These transformers are of two general types— the current and the potential. They permit the use of standard low- voltage meters for all high-voltage or high-current ac circuits. They also protect the operating personnel from the high­voltage circuits. For more information on instrument transformers, you should study NEETS, Module 2. It covers transformers in detail.

3-11

Figure 3-10 — Typical electronic power supply. Electronic Power Supplies In high performance aircraft, avionics systems help the pilot communicate, navigate, or f ire missiles. Other systems, such as radar and autopilot Automatic Flight Control System (AFCS), ease the pilot's w orkload. Each of these systems requires precision voltage inputs for proper operation. For example, an inertial navigation system may require the voltages shown in Table 3-1. The voltages required by an AFCS in the same aircraft are shown in Table 3-2. Obviously, one simple electrical power source won't provide all the needed power for 20 or 30 avionics systems. Normally, each avionics system has its own power supply. Figure 3-10 shows the power supply for a typical autopilot system. The power supply requirements (Table 3-2) are for a 120-/208- volt, three-phase, four-wire, 400- hertz electrical power input.

Table 3-1 — Electrical Requirements for an Inertial Navigation System DC VOLTAGES AC VOLTAGES +45 V transistor bias 26 V, 400 hertz -45 V transistor bias 140 V, 400 hertz, three phase +28 V unfiltered and unregulated 90 V, 375 hertz, three phase +28 V transistor bias 12.6 V, 400 hertz +28 V unregulated, relay excitation 6.3 V, 400 hertz +20 V transistor bias +10 V reference -10 V reference

3-12

Figure 3-11 — Developing ac voltages for the autopilot. Table 3-2 — Electrical Requirements for a Typical Transistorized Autopilot System Phase A Phase B Phase C DC 120 VAC 120 VAC 120 VAC 28 V filtered 45 VAC 26 VAC 15 VAC 28 V unfiltered 26 VAC 15 VAC 10 VAC 19 VAC 15 VAC 7 VAC

Look at the schematic shown in Figure 3-11. Autotransformers T4, T5, and T6 produce the majority of the output voltages. The autotransformers have taps from each transformer winding at the proper position to produce the required voltage. No further voltage regulation is necessary under fairly constant load conditions.

3-13

Figure 3-12 — Full-wave rectifier and filter network. Figure 3-13 — Precision dc voltage developer. Full-wave rectifiers (Figure 3-12) produce dc voltages. Each pair of rectifiers (either CR7 and CR10 or CR8 and CR9) conducts during alternate half cycles of the ac input from the secondary of step-down transformer T7. The unfiltered dc provides power to operate lights and relays for internal operation of the system and feeds a filter network. Also, filtered dc supplies transistor bias to the electronic amplifiers in the autopilot system.

The circuit shown in Figure 3-13 develops precision dc voltages. Diode CR1 is a Zener diode that develops a constant dc voltage at the input of amplifier A1, regardless of input voltage fluctuations. CR1 will conduct harder and the excess voltage drops across R1 if the dc input voltage at the top of R1 increases. If the voltage decreases, CR1 conducts less and less voltage drops across R1. This maintains the voltage at the anode of CR1 at a constant, precision potential. If no current flows through R2, the same potential present at the input of amplifier A1 and on the anode of CR1 is the same. Feedback voltage through resistors R3 and R4 control the gain of amplifier A1. The potential of the output voltage and the anode of CR1 are the same when the combined resistance values of R3 and R4 are the same as the resistance of R2. Variable resistor R4 provides fine tuning of the output voltage to the desired level. Isolation amplifier A1 prevents changes in the load current from being felt at the Zener diode. 3-14

Figure 3-14 — Typical aircraft storage battery with quick disconnect. There are many methods of providing both dc and ac precision voltages. NEETS, Module 6, Introduction to Electronic Emission, Tubes, and Power Supplies, NAVEDTRA 14178 discusses several of these methods. EMERGENCY POWER SOURCES Naval aircraft have backup (emergency) electrical power if primary sources of electrical power fail. The various ways of supplying this emergency power are aircraft storage batteries, hydraulic motor-driven generators, and ram air­ driven turbine generators. The following paragraphs discuss each of these systems. Batteries Aircraft storage batteries provide an emergency source of electrical power for operating e lectrical systems of an aircraft. The ac generator and transformer-rectifier combination supply electrical energy and maintain the battery in a charged state during normal aircraft operation. The battery supplies power to the aircraft only when the generating systems are unable to supply power. Figure 3-14 shows an aircraft storage battery with a quick disconnect. The battery is the emergency power source fo r the aircraft. As such, you should maintain the battery in perfect condition at all times. Never use the battery for starting engines or servicing equipment if another power source is available. Doing so shortens the battery’s life. The service life of the aircraft battery depends upon the frequency and quality of care it receives. The most common aircraft batteries used today are lead-acid, nickel-cadmium, and sil ver­ zinc batteries. For detailed information on batteries, refer to NEETS, Module 1. It covers the basic principles of b atteries. Another reference on batteries is Naval Aircraft Storage Batteries, NAVAIR 17-15BAD-1. Most wet-cell batteries emit some type of gas when being charged or discharged. This is especially true of lead-acid batteries and, to a lesser degree, the nickel-cadmium and silver-zinc batteries. A vent in the filler plug allows each cell to vent gas and moisture into the void of the battery. Allowing the moisture to stand in the battery void could cause shorting of the cells and corrosion, so openings at each end of the battery provide ventilation of the void area. In a vent system (Figure 3-15), the void above the cells and beneath the sealed cover is subject to differential pressure through the vent nozzles. The higher of the two vent nozzles connects to a rising vent tube exposed to positive pressure on the aircraft surface. This provides definite pressure on the battery while in flight. It acts as a chimney for light hydrogen gas when the aircraft is at rest. The lower of the two vent nozzles connects a tube exposed to negative pressure on the aircraft surface. This tube allows battery acid to escape without injury to the aircraft. 3-15

Figure 3-15 — Battery vent system. In the battery drain sump, the negative pressure tube from the battery connects to a jar sump and extends 1-inch into the jar. The exhaust tube from the sump jar is cut at a 30-degree angle. It extends into the sump jar for one-third its depth. T he tube then runs to the aircraft surface. Normally, the sump jar contains a felt pad. The lead-acid battery is moistened with a concentrated solution of sodium bicarbonate for neutralizing gases and excess battery solution. You should refer to the aircraft's MIM for specific directions concerning the maintenance of the vent-sump system for your aircraft. AC/DC Hydraulic Motor-Driven Generators The ac/dc hydraulic motor-driven g enerators are emergency power so urces. They consist of the following components: a hydraulically-driven ac/dc g enerator, a motor-generator control unit, and a control solenoid. The motor- generator provides 115-/200-volt ac and 28-volt dc power to essential electrical circuits if normal power fails. The kVA rating of this emergency generator is much lower than the primary generator(s). Hence, the emergency generator powers a limited number of cir cuits. AC/DC GENERATOR – As the rotor turns (hydraulic motor running), the permanent magnet induces power into the PMG. This power energizes the co ntrol and regulation circuits and the four essential power transfer relays. Regulated and rectified PMG output power goes to a stationary control field within the motor­generator. The motion of the rotor assembly induces PMG power into the windings of the exciter alternator. This power is rectified and, in turn, induced into the output ac winding and the dc winding. The motor-generator control unit monitors ac output. This unit adjusts the regulation to maintain the ac output at 115 volts per phase ± 1.50% when operating under full hydraulic system pressure. The dc output from the sta tionary rectifier goes to the dc transfer relay contacts for distribution to the essential dc buses. The hydraulic motor converts 3,000 PSI of hydraulic pressure to constant-speed rotation, maintaining generator output frequency at 400 Hz (Figure 3-16). The motor- generator is cooled by hydraulic fluid from the same source that drives the hydraulic m otor. 3-16

Figure 3-16 — Motor-generator.

MOTOR-GENERATOR CONTROL – The motor-generator control provides voltage regulation for, and detection of, motor-generator output. When the motor-generator is o perating, PMG power flows to the rectifier, where the three-phase ac is rectified to a dc signal. The dc signal is for control panel and relay control power. Then, the signal flows to the voltage regulator section of the motor­generator control unit. The voltage regulator supplies field excitation to the motor-generator and monitors the output voltage. Monitoring protective circuits in the motor-generator control prevent out-of- tolerance power from being connected to the essential bus system. MOTOR-GENERATOR SOLENOID CONTROL VALVE – The motor-generator solenoid control valve (Figure 3-17) controls the operation of the emergency electrical power system. The valve drives electrically to the closed position when primary electrical power is available. The valve de-energizes and opens, routing hydraulic pressure to the hydraulic motor, driving the generator when primary electrical power fails. The entire operation is completely automatic upon primary electrical power failure.

3-17

Figure 3-17 — Motor-generator solenoid control valve.

Ram Air Turbine Generator Some naval aircraft have a Ram Air Turbine (RAT) emergency generator, hence the name emergency generator, provides emergency electrical power in the event of main electrical power failure. Different types of emergency power generating systems are available, and their installation depends upon aircraft type/model/series. In the EA-6B installation, when the RAT is deployed, it extends into the airstream. This occurs when the pilot pulls upward on the ELEC-AIR TURBINE T-handle assembly, a lever, which mechanically links the turbine manual selector valve, ports hydraulic fluid to the extend port to hydraulically cause the RAT to protrude in the airstream. The ram air of flight (caused by the aircraft moving through the air) provides the turning power for the turbine blades. This, in turn, rotates the generator's armature. The system's design prevents the emergency generator from powering the bus until its armature is up to speed. Figure 3-18, view A shows a typical three-phase emergency generator. This generator has a wye-connected output capacity of 2.5 kVA, 120/208 volts at 400 hertz. Two variable-pitch turbine blades, (Figure 3-18, view B) a pitch adjusting mechanism, and speed governing mechanism all compose the governor assembly to assist the RAT to maintain a constant output voltage frequency. Ground testing of some types of emergency generators may be performed while the generator is on the aircraft. The generator is driven by funneling compressed air from a gas-turbine compressor onto the turbine blades during ground testing. 3-18

Figure 3-18 — (A) Emergency generator; (B) typical emergency generator installation.

AUXILIARY POWER UNITS Some aircraft have Auxiliary Power Units (APUs). APUs furnish electrical power when engine-driven generators are not operating, external power is not available, or the engine­driven generator fails. Using the pneumatic starting system, the gas-turbine APU provides compressed air to start engines and for air conditioning. The aircraft is made independent of the need of ground power units to carry out its mission. There are many types and configurations of gas-turbine units. Because of their similarity in construction and operation, only one is described in the following paragraphs. GTCP-95 Unit – The GTCP-95 is a gas-turbine power plant unit (referred to as an APU). It is capable of furnishing electricity, starting air, and air conditioning while on the ground by supplying air for the air-cycle cooling systems (Figure 3-19). The gas-turbine engine of the APU requires only the aircraft battery and fuel for starting. Shaft power at the main output drive pad powers the generator. Pneumatic power is available as clean, compressed air at the output end of the engine bleed load control and air shutoff valve. The engine is composed of two main sections and four main systems. The two main sections include an accessory assembly and a compressor and turbine assembly. The fo ur main systems consist of an electrical system, a fuel control system, a bleed-air sy stem, and a lubrication system. 3-19

Figure 3-19 — Gas turbine power plant unit (GTCP-95). The engine develops power by compressing ambient air with a two-stage centrifugal compressor. Compressed air, mixed with fuel and ignited, drives a radial, inward-flow turbine wheel. The rotating shaft of the turbine wheel drives the compressor, the accessories, and the output shaft for the ac generator.

3-20

Figure 3-20 — GTCP-95-2 gas-turbine engine electrical schematic. Ga s-Turbine Engine Electrical System – The gas-turbine electrical system ( Figure 3- 20) provides automatic actuation (in proper sequence) of the various circuits that control fuel, ignition, engine starting, acceleration, and monitoring. The electrical system consists of the following components: holding relays, oil pressure switch, centrifugal switch assembly, hour meter, and harness assembly. The ignition portion consists of an exciter and ignition plug controlled by the multiple centrifugal switch. Ignition is only required during starting and automatically cuts out at 95-percent engine RPM.

3-21

Figure- 3-21 — Centrifugal switch assembly. CENTRIFUGAL SPEED SEQUENCE SWITCH – This component (Figure 3-21) controls the sequence of operation of various electrical components. The input drive shaft turns a knife-edged fulcrum (flyweight support) and a pair of flyweights pivots on the knife edges. Each flyweight has a toe that lies under the outer race of a ball bearing on the actuating shaft. As the centrifugal switch turns, centrifugal force causes the flyweights to pivot, moving the actuating shaft to the right against the lever arm. The three electrical switches (Figure 3-20) actuate at 35 percent, 95 percent, and 106 percent of turbine speed. These switches are fine adjusted by applying spring tension to the lever arm with the three adjustment screws shown in Figure 3-21. The functions of the switches are as follows:  35 percent: turns off starter motor  95 percent: arms load control circuits, starts hour meter, turns off ignition  106 percent: stops unit (overspeed protection)

The flyweights move outward as the input shaft rotates. This action forces the actuating shaft to move the lever arm, actuating the switches. The lowest percent speed adjustment spring acts on the lever arm during the actuating of all switches. Changing its setting affects the setting of the 95- and 106-percent switches. A drift in setting of the 106-percent switch affects only the106-percent switch. 3-22

Checking an overspeed switch in a gas turbine requires that the unit operate above its governed speed. The centrifugal switch assembly on the engine incorporates a lever, which can be manually positioned to operate the three switches. The lever is spring loaded so it doesn't interfere during normal operation. When manually actuating the lever, it rotates on a pivot, making the centrifugal switch lever arm actuate the switches. This check may be performed with the unit operating or stopped. Actuation of the switch cuts off the fuel flow to the combustion chamber and stops the unit. STARTER MOTOR – The starter provides initial power for rotating the components of the gas turbine to self-sustaining speeds. It rotates the compressor to a speed high enough for correct airflow for combustion. Also, the starter assists acceleration after light-off, preventing excessive turbine temperature at low speeds. The starter motor rating is 1.5 hp at 14 volts at 5,000 RPM. The starter has a duty cycle of 1 minute ON and 4 minutes OFF. The starter motor armature shaft is splined and pinned to the clutch assembly. The starter clutch assembly performs two functions. 1. As a friction clutch, it prevents excessive torque between the starter and accessory drive gears to protect both. 2. As an overrunning clutch, it provides the means of automatically engaging the starter with the gear train for starting. The clutch automatically releases it when the unit has reached a condition allowing it to accelerate and run without assistance. The friction clutch section provides overtorque protection. The assembly will slip at 135- to 145-inch pounds of torque. Because of the inertia the engine offers when the starter motor pawls first engage with the gearbox ratchet, the overrunning clutch flange and splined clutch plates remain stationary. The motor, clutch housing, and keyed clutch plates rotate when in this state. Slippage occurs until engine and starter speeds have increased enough to develop less than the specified torque value. The starter is normally de-energized by the centrifugal switch at 35 percent. If the switch does not cut out at this speed, the starter may fail from overheating or it may fail mechanically from overspeed. Mechanical overspeed failure of the starter results when the overrunning clutch does not release properly. EXTERNAL POWER SOURCES Naval aircraft accept electrical power from an external source. This source provides ground crews with electrical power for servicing, fueling, and performing maintenance actions. Aside from fuel costs and engine wear, it is unsafe and highly impractical to turn up aircraft in the hangar or hangar bay to provide electrical power. Aircraft design features make it impossible to have both aircraft generator power and external power applied to the buses simultaneously. To protect the aircraft, monitoring circuits, ensure voltage, frequency, and phasing of external power are correct before the aircraft accepts power from an external source. These monitoring circuits are an integral part of the aircraft's electrical system. In principle, they operate as a supervisory panel (discussed later in this chapter).

3-23

NOTE Some brushless generators use permanent magnets in the exciter circuits. CONTROLS AND CIRCUIT PROTECTION The first part of this chapter dealt with the various devices used to provide electrical power in naval aircraft. This part deals with methods that regulate the output voltages of ac generators. To understand voltage regulation, you should be familiar with the principles of ac and dc power generation. NEETS covers these principles in detail. Before continuing with this chapter, you should review the appropriate NEETS modules. Two methods of voltage regulation have become popular in recent years. The most common method in power-generating systems is varying the current to the generator exciter winding (sometimes called field winding). This, in turn, changes the size of the magnetic field, which changes the voltage output of the generator. The second method of voltage regulation is to maintain a constant load on the generator. This method uses a permanent magnet on the generator rotor in place of exciter windings, which simplifies generator construction. This type of regulation must, however, be used with systems that supply constant loads and have a limited capacity. For example, an inverter or an electronic power supply uses this type of voltage regulation. The regulator varies the resistance of a parallel resistor, so total resistance remains constant regardless of the load resistance. This type regulator is for use with both ac and dc power sources. AC GENERATOR CONTROL When magnetic fields of alternating polarity pass across the armature windings, ac voltage induction occurs. The voltage induced into the windings depends on three things. All of the following three things can control the voltage induced into the ac generator windings: 1. The number of turns of conductor per winding 2. The speed of the magnetic field passing across the winding (generator RPM) 3. The magnetic field strength The number of turns per winding and the number of windings is set during generator manufacture. The frequency of the output voltage depends on the speed of the generator. The strength of the magnetic field controls the level of output voltage. In some cases, as in tachometer generators, a permanent magnet field maintains the load at a constant value. In today's aircraft, electrical and electronic equipment operate at exact frequencies and voltages. Systems exposed to extreme overvoltages or off-frequencies not only destroy themselves, but may start a fire during an emergency. All ac generator control systems must contain circuits to protect against under voltage and overvoltage, under frequency and over frequency, and improper phase sequence. The generators shown in Figure 3-22, views A and B, use an electromagnetic field rather than a permanent magnet-type field.

3-24

Figure 3-22 — AC generators. (A) Brush type; (B) brushless type. The current flowing through the field controls electromagnetic field strength. Varying the voltage applied across the field helps to control the field strength. By varying the dc o utput voltage from the exciter armature, you control the ac generator field strength. The value of the generated ac voltage depends directly on the size of the exciter input. This relationship allows a small dc voltage to control a much larger ac voltage. The rotating three­phase rectifiers on the brushless generator help change the ac output of the exciter to dc. Then, the dc feeds the main ac generator rotating field, eliminating the use of brushes.

As you have already learned, controlling the strength of the magnetic field controls the generated voltage. A voltage regulator controls the magnetic field strength. Current g enerating the magnetic field is known as excitation current. The auxiliary dc generator (called the exciter) or a rotating three-phase rectified ac exciter generator supplies this cu rrent. The exciter is on the same shaft as the ac generator to make it an integral part of the generator. The present military specification for aircraft ac generators states that they should be self-supporting. All dc exciter units are integrated into the ac generators to meet this requirement. The chief advantage of exciter units is that each generator has its own ind ependent source of excitation. No external source of electric power is necessary for generator operation. In a multigenerator installation, failure of one generator exciter 3-25

Figure 3-23 — Solid-state voltage regulator. does not make the complete system inoperative. This would happen if a generator system had a common external excitation system. Internal excitation makes it u nnecessary to transmit excitation power, which reduces the chances of losing excitation from an open or short-circuited wire. In contrast to dc generators, the magnetic field coils in most aircraft ac generators rotate. This induces the ac voltage into the stationary windings. A solid-state regulator is a type of voltage regulator that has no mechanical moving parts (except the exciter control relay). The ac generator output flows to the voltage regulator, which compares it to a reference voltage. The difference supplies the control amplifier section of the regulator (Figure 3- 23). If the output is too low, regulator circuitry increases the field strength of the ac exciter. It reduces the field strength if the output is too high.

The power supply for the bridge circuit is CR1. CR1 provides full-wave rectification of the three-phase output from transformer T1. The dc output voltages of CR1 are proportional to the average phase voltages. The negative anode of CR1 supplies power th rough point B, R2, point C, Zener diode CR5, point D, and to parallel­ connected V1 and R1. Takeoff point C of the bridge is located between resistor R2 and the Zener diode. The other leg of the reference bridge (resistors R9, R7, and temperature compensating resistor RT1) connects in series with V1 and R1 through points B, A, and D. The output of this leg of the bridge is at point E. As voltage changes occur, voltage across R1 and V1 (once V1 starts conducting) remains constant, leaving the total voltage change occurring across the bridge. Because v oltage across the Zener diode remains constant (once it starts conducting), the total v oltage change occurring in that leg of the bridge is across resistor R2. The voltage change across the resistors is proportional to their resistance values in the bridge’s remaining leg.

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For this reason, the voltage change across R2 is greater than the voltage change at point E. If the generator output voltage drops, point C is negative with respect to point E. Conversely, if the generator voltage output increases, the voltage between the two points reverses polarity. The bridge output taken between points C and E connects between the emitter and the base of transistor Q1. With the generator output voltage low, the voltage from the bridge is negative to the emitter and positive to the base. This is a forward bias signal to the transistor, and the emitter to collector current increases. The voltage across emitter resistor R11 increases with the increase of current. This increase, in turn, applies a positive signal to the base of transistor Q4, which increases emitter to collector current and increases the voltage drop across emitter resistor R10. This gives a positive bias on the base of Q2, which increases its emitter to collector current and increases the voltage drop across its emitter resistor, R4. This positive signal controls output transistor Q3. The positive signal on the base of Q3 increases the emitter to collector current. The control field of the exciter generator is in the collector circuit. Increasing the output of the exciter generator increases the field strength of the ac generator, which increases the generator output. An under speed switch, located near the F+ terminal, prevents generator excitation when the frequency is at a low value. The switch closes and allows the generator excitation when the generator reaches a suitable operating frequency. Resistors R27, R28, and R29 connect in series with the normally closed contacts of relay K1. The coil of relay K1 connects across the power supply (CR4) for the transistor amplifier. Electricity from the 28-volt dc bus goes to the exciter generator field to flash the field for initial excitation when the generator starts turning. When the field of the exciter generator energizes and the ac generator output voltage increases, relay K1 energizes, opening the field flash circuit. 3-27

Figure 3-24 — Solid-state voltage regulator schematic. Another type of solid-state voltage regulator (Figure 3-24) operates by sensing the voltage existing on the lines. It amplifies the changes in this signal, and varies the average current supplied to the field winding of the integral exciter. The voltage regulator consists of a sensing circuit with input rectifiers, a temperature compensated Zener diode, reference and error- detecting bridge, and a three-stage transistor amplifier. The output of the bridge circuit is a voltage inversely proportional to the difference between generator voltage and regulator set voltage. This output is referred to as the error signal. Transformer T1 in the regulator supplies three-phase, ac generator output. It provides isolation from the generator and delivers correct utilization voltages. The transformer o utput passes through the full-wave bridge rectifier (CR1) to obtain a dc voltage to supply the comparison circuit. The rectifier output is proportional to the average of the three line voltages. The voltage reference and error-detecting bridge uses this voltage for comparison with the constant voltage across the Zener diode (CR5). This achieves a means of telling whether the generator output is too high or too low. Potentiometer R7 permits adjustment to the desired voltage. The glow tube (V1) serves to increase the sensitivity of the voltage reference and error-detecting bridge. Thermistor RT1 provides temperature compensation in the comparison circuit. It offsets the effects of changes in other elements of the circuit that result from temperature variations to maintain a nearly constant voltage. 3-28

Figure 3-25 — Pulse width modulation diagram. The error-detecting bridge output voltage sawtooth wave shape is because of the ripple resulting from the semifiltered, three-phase rectifier supply. This sawtooth voltage goes to the input of the first stage of the three-stage transistor amplifier. Overdriving the second and third stages obtains an essentially square-wave output. The effect of the error detecting bridge output is to modulate the width of the pulses passing through the amplifier. Varying the output current to the exciter field varies the width of the square- wave impulses. Figure 3-25 shows a pulse width modulation diagram. As the voltage rises (shown by the dotted back-to-back sawtooth), the square-wave pulse to the exciter field is off longer than it is on. This causes the output of the ac generator to decrease. The decrease in voltage causes the back-to-back sawtooth to drop to its normal value (shown by the solid waveform). This causes the on and off times of the square-wave pulse to the exciter field to be about equal. Varying the on and off excitation to the exciter field controls the ac generator output. Refer to Figure 3-25 as you read this paragraph. The power for operating the three- stage transistor amplifier comes through the full-wave bridge rectifier (CR4) from transformer T1. Obtaining amplifier power this way requires special consideration. There are conditions that require excitation when no voltage is available to supply the amplifier. Such conditions exist during initial buildup of system voltage and during three- phase short circuit on the generator. Control relay (K1) connects across the full-wave bridge rectifier (CR4) overcoming these obstacles. When the relay is de -energized, its contacts provide PMG voltage to the exciter field. When generator voltage is 90 volts line to line, voltage across CR4 energizes control relay (K1), removing the self-excited field circuit. The voltage regulator then supplies the exciter field. The absence of phase shift and fast response characteristics of transistor-type amplifiers eliminates feedback networks and stabilizing transformers in this voltage regulator.

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Figure 3-26 — Simplified CSD functional diagram. Frequency Control Because of the fixed number of poles, the only means of fine tuning the output frequency is controlling rotor RPM. CSD units are often used for controlling generator roto r RPM. CSDs receive drive power from hydraulic power, pneumatic power, or the accessory drive section of an engine. A CSD unit is located between the aircraft engine and the ac generator in most aircraft for this purpose. The purpose of the CSD is to transfer and convert aircraft engine variable-speed rotation to a constant-speed rotation, which drives the generator. The CSD consists of a variable­displacement hydraulic pump, constant-displacement hydraulic motor, and a governing system. The governing system controls the rate of flow from the pump, thereby controlling the speed of the motor. There are several other components in the CSD that are necessary for self-regulating constant-speed operation. Among these components are three output-driven gear pumps: the charge pump, replenishing pump, and scavenge pump. A gear on the CSD output shaft drives these pumps, the limit governor, and basic governor. The pump wobbler and the pump section of the cylinder block assembly form the variable displacement pump in the CSD. Figure 3-26 shows a simplified CSD functional diagram. The pump wobbler consists of an outer stationary shell and an inner race. The inner race is separated from the wobbler shell by bearing rollers. It is free to turn with the pump pistons, which are always in contact with the race during operation. Two control pistons in the CSD housing move the wobbler sideways to vary the output of the pump. The CSD functions in three different phases of operation. They are the overdrive phase, straight-through phase, and the underdrive phase.

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The CSD makes up the difference in RPM when the engine input RPM is less than the RPM required for the generator. The CSD does this by causing the pump wobbler system to respond to governor signals. This response causes the pump to supply more oil to the motor. The difference between the input and output RPM depends on the quantity of oil pumped by regulating the wobbler pump. The CSD is in overdrive anytime the motor wobbler (output) is rotating faster than the cylinder block assembly (input). When the input RPM equals the required output RPM, the rotary motion transmits through the CSD without hydraulic action. The pump wobbler would, theoretically, be positioned through the action of the governor to be concentric with the cylinder block assembly. The pump neither pumps oil to the motor nor accepts oil from the motor in this condition. The motor pistons lock in position against the motor wobbler, forcing the wobbler to rotate at the same speed as the cylinder block assembly. Because the drive starts in underdrive and operates normally in overdrive, this straight-through condition is only temporary. The CSD acts to subtract from the input rotation when engine input RPM exceeds the output RPM requirements for the generator. The wobbler pump accomplishes this in response to the governor signal. The pump-motor action for underdrive is the reverse of the action required for overdrive. The pump performs in a negative pumping action in the underdrive phase. The generator load opposes the driving force of the CSD, so it always tries to slow the wobbler. The cylinder block assembly then rotates faster than the motor wobbler. Excess input torque is dissipated in the reverse pumping action to the charged oil system. The CSD is in underdrive whenever the motor wobbler is rotating more slowly than the cylinder block assembly. The CSD goes into underdrive to protect the generator from overspeed when the engine overspeeds or if the basic governor fails. The underspeed pressure switch in the governor oil line functions to break the electrical circuit of the ac control system. The system is protected during an underspeed condition. In some CSDs, aircraft engine oil from the engine lubricating system is the hydraulic medium. In this case, the CSD also functions as a pump for supplying the generator with engine oil for cooling. Oil-cooled generators are of smaller construction than air-cooled generators having a similar rating because of cooling capabilities. CIRCUIT PROTECTION The generator and equipment and systems the generator powers need protection if a malfunction occurs. Circuits designed to sense malfunctions and energize relays provide protection by either warning the pilot of the malfunction or disconnecting the generator. The circuit protection needed and the methods used to control the malfunctions depend on aircraft and equipment design. For example, in a single-pilot aircraft, all malfunction detection and correction might be automatic. In a multi-piloted aircraft, the generating system may only warn the flight crew of a problem. This leaves corrective action to the discretion of the pilot in command. A supervisory panel provides regulation and circuit protection for both the operating generator and equipment it powers in newer generating systems. This single component provides the same functions as several components in older power generating systems. The supervisory panel provides voltage regulation at 120/208 volts ac, while some types of CSDs provide frequency control at 400 Hz. The supervisory panel further has relays and other associated circuitry to disconnect the generator from the load if any of the following conditions occur: 3-31

 Underfrequency  Overfrequency  Undervoltage  Overvoltage  Feeder fault (A condition where the current leaving the generator does not pass through the load. System design has cut out the need for feeder fault protection in systems where it isn't likely to occur.) UNDERFREQUENCY AND OVERFREQUENCY CONTROL You should refer to Figure 3-27 as you read this section. The PMG output is 39 volts at 600 Hz when the generator is on speed. The voltage reference bridge and the frequency sensitive bridge sample output voltage and frequency. The band-pass filter is tuned to 600 Hz (called its resonant frequency). Its minimum resistance and maximum current flow occur at 600 Hz. The output of the bridge networks are equal and opposite at this frequency. The underfrequency/overfrequency sensor senses an on-frequency condition energizing the underfrequency/overfrequency relay (K1). Current flows through contacts 4 and 6 of energized relay K1. This allows generator control relay (K2) to energize if the frequency remains within tolerance for at least 3 seconds. 3-32

Figure 3-27 — Generator control system voltage regulator/supervisory panel.

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The band-pass resistance increases and output of the circuits is unbalanced if the PMG frequency changes from the desired 600 Hz. The underfrequency/overfrequency sensor senses the unbalance and causes K1 to de-energize and immediately cuts off SCR-1. K2 de-energizes and disconnects any input to the exciter stator coils and reduces the generator output voltage to zero. Contacts 1 and 2 of K1 change frequency tolerance from 600 Hz ± 42 to 600 Hz ±53 by adding resistance to the voltage reference bridge circuit when K1 energizes. This prevents the relay from chattering when the generator is operating at or very near its tolerance limit. VOLTAGE CONTROL A voltage regulating circuit changes PMG ac voltage to dc voltage and controls its amplitude. The voltage regulator senses all three phases of the generator output. If the average of these voltages is low, dc voltage to the exciter stator coils increases until output voltage is at the desired level. If output voltage is high, the voltage regulator decreases its output to the exciter stator coils until voltage is within tolerance. The generator system maintains three-phase output voltage to 120 volts ± 2 through a wide range of loads from 1 to 120 kVA. One-phase load may be one-third more than the other two-phase loads, and voltage will not vary more than 5 volts between phases. It takes 1.7 amperes of current through the exciter stator coils to produce the desired magnetic field to generate a 120-/208-volt, 60-kVA load. Undervoltage Refer to Figure 3-27 as you read this section. The undervoltage sensing and control circuit allows generator output to power the distribution system when voltage rises to 105 volts during initial generator buildup. However, it does not de-energize the generator output until one or more phases fall below 90 volts. The undervoltage sensor monitors generator output. In conjunction with K1, it also energizes auxiliary control relay (K3), connecting generator output to the power distribution system. When K3 energizes, its contacts arm a timing circuit that acts automatically when one or more phases are 90 volts or less. The timing cycle duration is electronically divided into a 3-second period and a 1-second period, in that sequence. The two periods are additive. The total time involved before an undervoltage trip occurs is about 4 seconds. The delay circuitry allows time for corrective measures (circuit breakers or current limiters to open) to remove the cause of the undervoltage. The generator stays on line if the cause of the undervoltage is removed and voltage rises to 105 volts before the initial time delay lapses. This cancels the lapsed increment and the full 4-second delay is reinstated. However, after a 4-second delay, differential protection latch-out relay (K6) energizes, energizing lockout relay (K4), and removes power from K3 and K2. An excessive load on the generator can cause an undervoltage (a short circuit in a system that has a defective circuit breaker or fuse). This condition, if allowed to continue, could cause a fire or destroy the generator. Therefore, both K6 and K4 have holding circuits to keep them energized even when the undervoltage condition is corrected. To check for correction of the undervoltage, use the following procedure:

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1. Place the GEN CONTROL SWITCH to the OFF position 2. Pull and reset the GEN CONT circuit breaker 3. Return the GEN CONTROL SWITCH to ON Overvoltage As you read this section, refer to Figure 3-27. An overvoltage sensor senses line voltage above 129 volts and starts a time delay. When started, the delay times out for a time inversely proportional to the overvoltage. A voltage of 130 volts on a single phase may have a delay of 3 to 4 seconds. A large overvoltage on all three phases may have a delay of a few milliseconds. When the delay completes timing, it triggers SCR2 into conduction and allows K4 to energize. An overvoltage occurs if a voltage regulator malfunctions or if a large load (several loads) is removed from the generator at once. The voltage regulator is not fast enough to react when the generator loses several loads quickly. That is, it is possible for an overvoltage to occur during normal operation of the generating system. K4, supplying its own holding circuit, prevents the generator from powering the load again. If you place the GEN CONTROL SWITCH to either the RESET or the OFF position and back to ON, the generator stays on line. This prevents a generating system with a malfunctioning voltage regulator from cycling on and off. Feeder Fault System A short occurring between the generator and distribution system would cause a fire because there aren't any protective devices (such as circuit breakers and fuses). To protect against this possibility, a feeder fault circuit (Figure 3-27) was designed. The generator armature winding (output) has current transformers on each side of each winding. One set of current transformers (on the grounded side) is as close to the armature windings as possible. The other set is as close to the distribution system (and its protective devices) as possible. Then, the transformer's connections are made so the voltages produced cancel each other out. The input to the feeder fault sensor would then be nearly zero. A short to ground or phase to phase would place a voltage across R2, causing the feeder fault sensor to energize differential protection relay (K7). K7 then acts to energize K6, K6 energizes K4, and K4 de-energizes K2 and K3. Because K7 remains energized by its own contacts using PMG voltage, the system cannot be reset until removal of PMG voltage by stopping the generator. Power in AC Circuits In a dc circuit, the equation P = EI is used to compute power. For example, watts (P) equal volts (E) time’s amperes (I). If 1 ampere flows in a circuit at a pressure of 200 volts, the power is 200 watts. The product of the volts and the amperes is the true power in a dc circuit. In ac circuit, a voltmeter shows the effective voltage and an ammeter shows the effective current. The product of these two readings is apparent power. Only when the ac circuit is of pure resistance is the apparent power equal to the true power. When the impedance of the circuit is either inductive or capacitive, current and voltage are not in phase, and true power is less than apparent power. You obtain true power by using a wattmeter to read the system. The power factor is the ratio of true power to apparent power, and is equal to true power divided by apparent power.

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Figure 3-28 — Three-phase ac bus.

Equipment using ac power should have as near a unity power-factor load as practicable. This improves the efficiency of power distribution by reducing the line current and 12 R (power) losses. Most ac loads in an aircraft are somewhat inductive, resulting in a lagging power factor. Power-factor corrections are made by connecting a capacitor of the proper capacitance in parallel with the circuit. Make the connection as close to the inductive load as possible. The non-energy component of the current in the inductive branch is 180˚ out of phase with the capacitive current. These currents circulate between the capacitor and inductive loa d and do not enter the line. The vector sum of capacitor current and total inductive load current is equal to line current. The line current is now in phase with the applied voltage to the parallel combination of the inductive load and the capacitor. This reduction in line current reduces line loss and increases the efficiency of transmission. Additional information on power factor and power-factor correction is found in NEETS, Module 2, Introduction to Alternating Current and Transformers, NAVEDTRA 14174. Refer to it, if you need to review this topic. Power Distribution You have learned that various sources are used to provide electrical power to operate aircraft electrical equipment and systems. In this section, the system that connects the electrical power source to the equipment is discussed. Each manufacturer develops a system that meets the needs of their particular aircraft design. A system of priorities ensures certain critical equipment is operable if there is a malfunction. For example, if a power lead used to start an engine shorted out during flight, it is inappropriate to sacrifice all electrical power, especially power to lighting, navigation equipment, flight instruments, and other essential equipment. Therefore, systems of like priority are on a common line called a bus. Each type of aircraft has a group of buses identified by the priority of the equipment it powers. For e xample, a flight-essential bus may power emergency lighting, critical f light and engine instruments, and/or an emergency radio. Less important critical equipment rece ives power from an essential bus. Normal systems used to complete the assigned mission or provide crew comfort are on the main bus. The input to a bus may be either dc or ac. The output from the bus has a protective device such as a circuit breaker, fuse, or current limiter. A three-phase ac bus has three separate common lines, one for each phase (Figure 3-28). Sometimes, schematics show three phases drawn as one line. 3-36

Figure 3-29 — Simplified P-3C electrical power distribution.

The P-3C aircraft is an example of versatility and flexibility in electric systems. Figure 3- 29 shows a portion of the P-3C power distribution system.

Operation of the P-3C electrical power distribution system during normal flight conditions is entirely automatic. The crew only monitors the control panel for any indication of a malfunction. Control of the system is also automatic during ground operation, except switching to and from external power.

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POWER SOURCES This section contains a discussion of a representative power distribution system used in F/A-18 Super Hornet aircraft. Electrical power is provided to the electrical power distribution system buses from these sources: 1. Left generator 2. Right generator 3. External power Two engine-driven Airframe Mounted Accessory Drives (AMAD) supply the mechanical driving force to turn two 50/65 kVa,115/200vac, three-phase, four-wire wye connected, 400Hz Variable Speed Constant Frequency (VSCF) generating systems. The generators work to supply electrical power both in engine start mode and ground maintenance mode. The two generators, connected to independent buses, supply electrical power to the essential, non-essential and maintenance buses. Either of the generators can supply the entire electrical demand of the aircraft in the event one generator fails. Two power supplies supply the dc power. The power supplies convert 115vac to 28vdc for power distribution used as the primary dc system. The generators send direct dc voltage to the Power Distribution Panels (PDP) as an Essential Bus Backup (EBB). Additionally, a battery system also provides the maintenance bus dc power as a secondary backup if either power supply fails. Either power supply is capable of supplying the entire dc requirements of the aircraft. AC Bus Distribution System The distribution system consists of sixteen PDP, two essential circuit breaker panels and a multitude of Relay Module Assemblies (RMA), which comprise the electrical system: 1. Left main 115vac bus 2. Right main 115vac bus 3. 26vac bus The left and right 115vac buses and the 26vac bus, through a s plit bus distribution network, distribute internal/ external power through the ground power distribution system and ground power switching system. A set of four magnetically held, ground power (GND PWR) switches, which minimizes operating time on selected equipment, control external input power to components and systems. Internal power is provided by the two generators and external power provided by a deck edge source or power cart. The power distribution system provides the essential buses and PDPs, the required ac, and dc power to aircraft systems. DC Bus Distribution System The main sources of dc power are the left and right power supplies. The left and right power supplies provide primary dc power by converting 115vac, 3-phase, and 400 Hz power from their respective generators to power the left and right 28vdc bus in PDPs 1 and 6. The left and right 28vdc buses provide power to the dc essential bus. A 28vdc output is provided as EBB power when both generators are operating. Both power supplies are powered through bus tie circuitry with a single generator failure. The maintenance battery, a secondary dc power source, supplies power to the essential 3-38

NOTE Pins E and F are jumpered in the external power cable plug. and maintenance buses. The battery is controlled by a magnetically held battery (BATT) switch and has voltmeter readout of maintenance battery when the ac buses are not powered. The voltmeter indicates battery charger converter voltage when ac buses are powered. Operation As you have already learned, two generators and two power supplies or an external ac power source provide electrical power. During ground operations, operating in the Ground Maintenance Mode (GMM) or external electrical power can supply electrical power to the aircraft electrical systems by using ground power switching where power is optimized by selecting the systems needed for testing and troubleshooting. If engines are operating and hydraulic pressure available, the generators are a useable source of power and ground power switching functionality is disabled. External Power The external electrical power system permits application of three-phase, ac power to the aircraft electrical power distribution system (Figure 3-30). Three-phase external power goes to the external power monitor and contacts A1, B1, and C1 of the external power contactor. The power monitor prevents application of external power not within tolerances. If an undervoltage, an overvoltage, an underfrequency, an overfrequency, and phasing condition exist, the power monitor disconnects external power from the power distribution system. When all the power parameters are within tolerance, the external power monitor switch control relay energizes, supplying 28vdc from the external power monitor rectifier, through the external power (EXT PWR) switch lower contacts to pin F of the external power receptacle.

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Figure 3-30 — External ac power and ground power switching functional block diagram.

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The power then runs through pin E and energizes the coil in the external power contactor. Then, three-phase power from the external power cable is provided to contacts A2, B2, and C2 through the energized contacts of the external power contactor, which powers the left 115vac bus and right 115vac bus through their respective contactors. Note the right bus tie contactor shown energized, powers the right 115vac bus. The right 115vac bus will not be powered when the battery switch is set to ON and the PARK BRAKE is released. Power from the left 115vac PDP No. 5 bus goes to power the left power supply and the right 115vac bus PDP No. 2 goes to power the right power supply. PDP No. 2 also powers the 26vac transformer assembly to provide step down phase C voltage for other various aircraft components. After engine start and the left generator comes on line (L GEN light extinguishes), the left power contactor automatically disconnects external power. The indications of external power application cease when the respective generator light goes out and the ground power switches reset to the auto position because of operating on internal power. At that time, the only control the pilot has over external power being applied or removed is the hand signals between the pilot and the plane captain. Ground Power Switching Operation During external power application (Figure 3-30), the EXT PWR control switch is set to RESET, then back to NORM and the external power contactor energizes. L 28vdc bus power from PDP No. 1 is routed to the GND PWR switches through the energized contacts of external power contactors and the de-energized contacts of left and right power contactors. The energized external power contactor connects negative side of ground power relays with respect to ground. When GND PWR switches in AUTO position (shown), 28vdc is routed to the ground power switches contacts. The holding coils energize whenever any ground power switch is set to A ON or B ON. 28vdc is then routed from the ground power switch contacts to the positive side of coils of the ground power relays and the ground power fault sensing relays. The energized ground power relays de-energize the controlled equipment. With any ground power switch (except for 4 to A ON) set to A ON or B ON for 3 seconds and no overheat condition exists (avionics cooling supplied), 28vdc is removed from respective ground power relays. The de-energized relays apply power to specific equipment which is controlled by ground power switch relays. A ground is routed to the respective switch holding coil from the external power contactor through energized undercool warning relay. The under cool warning relay is energized when no overheat condition exists. Depending which ground power switch is actuated, the holding coil will energize, holding the switch in selected position either A or B ON. If ground power is interrupted for any reason, the external power contactor will de-energize and the holding coils de-energize, thus returning all GND PWR switches to the AUTO position. If an avionics overheat condition exists, the under cool warning relay de-energizes and all grounds are removed from the holding coils. The holding coils will de-energize and the GND PWR switches return to the AUTO position. With ground power switch 4 placed to A ON position, the holding coil receives ground directly from external power contactor. The holding coil will not de-energize if avionics overheat condition exists. Systems controlled by ground power switch 4 to A ON do not need cooling air.

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Figure 3-31 — Simplified ac power distribution schematic. Main Generators Refer to Figure 3-31. With external power connected and the left engine start is initiated, the left AMAD mechanically turns the left generator, and the left generator comes on line when all parameters are within tolerance. Momentarily, the L GEN light on the caution light indicator panel extinguishes. The left power contactor energizes with the left generator online. The left generator now supplies power through the left power contactor, contacts A1 and A2 to the left 115vac buses at PDPs No. 5 and 9. Additionally, through the energized contacts of the right bus tie contactor, contacts A2 to A1 and through the de-energized contacts of the right power contactor, contacts A2 to A3, the right 115vac bus in PDPs No. 2 and 8 are energized. The left and right 115vac buses, No. 2 and 5 in turn, provide 115vac power to the left and right power supplies and the phase C, 26vac bus. The left power contactor removes external power from the aircraft bus system when it energizes. The external power unit is now shut down, and the cable removal may be done safely. After right engine start is initiated and the right generator is on line, the right power contactor energizes, allowing the right generator to power the right 115vac bus at PDP No. 2. The energized contacts of the right power contactor, contacts A1 to A2, prevent the left generator from powering the right 115vac bus. Likewise, the energized left power contactor, contacts A1 to A2 prevents the right generator from applying power to the left 115vac bus. 3-42

Figure 3-32 — Simplified dc power distribution schematic. For primary flight control operation, the generators provide the required 28vdc power to the flight control computers from the two isolated PMGs and converter regulators th rough switching relays. The generator will shut off the respective 28vdc output if the 28vdc output voltage does not meet designated output criteria. In the event 28vdc shuts down, the flight control system circuits receive power from the essential buses, powered by the left or right power supplies or maintenance battery. The left power supply (Figure 3-32) receives power from the left 115vac bus, PDP No. 5 for primary dc power operation. The left power supply provides primary dc power to the left 28vdc bus and essential and maintenance buses. One generator supplies both power supplies through bus tie circuitry. The right primary dc bus is connected to the left primary dc bus through two dc bus tie current limiters. The left or right primary dc buses power the essential bus through the de-energized contacts of the battery contactor contacts A3 to A2. The functioning power supply supplies all dc loads through the dc bus tie current limiter holders during a single-power supply failure. The maintenance battery supplies power to essential dc loads during a dual-power supply failure.

The generators provide the 28vdc essential bus backup for essential bus back up operation. They are supplied by isolated PMG and converter regulators contained inside the generator. The right generator is primary EBB power supply source for the 28vdc system and supplies the dc power during dual power supply failure. The left generator supplies EBB power during a dual power supply and right generator failure through the 3-43

de-energized right EBB contactor, contacts A2 to A3. The secondary source of dc EBB power is either the left or right generator, where either generator supplies all the essential dc bus loads. The left or right generator energizes the respective EBB contactor. If no faults are present, 28vdc EBB power is routed through energized contacts of the battery contactor to the essential dc buses. If faults are present, the signal data computer sends the data to the mission computer when a 28vdc power source can no longer provide flight control or EBB power and the mission computer commands the left digital display indicator to display L or R DC FAIL caution. For secondary dc power (Figure 3-32) to the essential buses, the secondary dc power source is the maintenance battery. The maintenance battery supplies power to essential loads when the right 28vdc bus is not powered. With BATT switch set to ON, aircraft weight-on-wheels, APU control switch set to OFF and the right dc bus not powered, one minute time delay battery cutoff relay applies 28vdc to the holding coil of the battery switch. The coil of the battery cutoff relay energizes one minute after power is applied. The battery contactor energizes and powers the essential dc buses after relay switching operation is complete. This operation, which is used for extracting maintenance monitor panel codes from the nose wheel digital display indicator, is used mainly by maintenance personnel. The right generator provides power to the right 115vac bus. The right 115vac bus provides power to the right power supply. The right 28vdc bus receives power from the right power supply and is connected to the left 28vdc bus through two bus tie current limiter holders and powers the essential buses. The bus is not affected by a loss of either generator or either power supply because the bus receives power from both power supplies. Component Failures The power distribution system design ensures power is available to operate all aircraft equipment. This includes all equipment essential to accomplish the assigned mission and ensure safety of flight. Power must be available for continued safe operation if a component or engine should fail. The distribution system design provides a continuous power source under all adverse conditions. As you know, either of the generators is capable of supplying the entire load of the aircraft. A master reset for any failed generator is provided by the electrical power control panel. Likewise, either of the power supplies is able to supply the entire dc load. Adverse conditions that could occur in the aircraft's electrical systems include the following:  Left generator failure  Right generator failure  Left power supply failure  Right power supply failure The following paragraphs discuss each adverse condition. Refer to Figure 3-31 and Figure 3-32 as you read about each condition. LEFT GENERATOR FAILURE – When the left generator drops off line, the left power contactor de-energizes. The right generator powers the left 115vac buses through the energized left bus tie contactor, the de-energized external power contactor and de- energized left power contactor. Left and right ac buses receive 115vac power. RIGHT GENERATOR FAILURE – When the right generator drops off line, the right power contactor de-energizes. The left generator powers the right 115vac bus through 3-44

Figure 3-33 — Grounded system. the energized left power contactor and energized right bus tie contactor and de- energized right power contactor to the right 115vac buses. Left and right ac buses receive 115vac power. LEFT POWER SUPPLY FAILURE – When the left power supply fails, the left 28vdc bus PDP No. 1 is powered from the right power supply through both the left and right dc bus tie current limiter holders. The functioning power supply supplies all the dc loads, through the de-energized battery contactor, thus powering the dc essential buses. There is no crew station warning of a single power supply failure. RIGHT POWER SUPPLY FAILURE – When the right power supply fails, the right 28vdc bus PDP No. 6 is powered from the left power supply through both the left and right dc bus tie current limiter holders. The functioning power supply supplies all the dc loads through the de-energized battery contactor powering the dc essential buses. In case of the loss of both power supplies, the 28vdc maintenance battery provides power to the essential buses through the energized battery contactor. As shown, the power distribution system maintains its integrity with a loss of either generator or power supply. This ensures all systems are available for safe flight and mission accomplishment. Grounded Systems The term grounded system means that one leg of the system connects to a common conductor. This common conductor can be the earth, the skin of the aircraft, or to a structural member of the aircraft. This conductor may serve as one leg of the circuit when the grounded leg of the circuit connects to a good electrical conductor. This cuts out the need for a separate conductor for this leg of the circuit. Figure 3-33 shows a simple grounded system. Even though the grounds are at different points, the potential at these points is the same because they connect to a common conductor. The letter N designates any wire that completes the equipment circuit to the ground network. A ny wire designated as N may come in contact with ground at any point without causing the equipment to malfunction. Grounding three-wire systems can be done by grounding one of the phases, usually the B- phase in aircraft. Make sure you ground the same phase in all equipment. Figure 3-34 shows the grounded three-phase systems. In four -wire systems, the neutral is ground. 3-45

Figure 3-34 — Grounded three-phased systems. The grounded circuit is better than the ungrounded one because it reduces overall weight by using fewer conductors. This results in a reduction in cost and space requirements. Other advantages are that troubleshooting is simplified and the impedance of the ground return path is lower than that of a run conductor. A disadvantage of a grounded system is that short circuits result when a bare spot on any ungrounded conductor touches ground. Another disadvantage is having circuits of different potentials and frequencies use a common ground. There is a possibility of one circuit feeding into another. This problem often happens in electronic circuits. Ungrounded Systems The term ungrounded system means that the circuit in no way connects to ground. All conductors run from the power source to the loads. Circuits of this type are often referred to as being above ground. The ungrounded system has one advantage — it prevents one circuit from feeding into another because the circuits are completely insulated from each other. The system has the disadvantage of adding more weight because it requires more conductors than the grounded system. This results in added cost and space requirements. Both the grounded and ungrounded systems are used for specific purposes in modern aircraft. Single-Phase and Polyphase Systems Single-phase systems are of simple design and construction. They are used when there are relatively low power requirements. Polyphase systems are more complicated in construction and design. These systems are used when high power is required. These systems provide a smoother source of power. Single-phase power is available from polyphase systems. The load on the polyphase system must be kept balanced when doing this.

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End of Chapter 3 POWER GENERATION AND CONTROL SYSTEMS Review Questions 3-1. What unit converts mechanical energy into electrical energy?

A. Engine B. Power transformer C. Generator D. Converter

3-2. The use of what system resulted in better avionics systems design and use?

A. AC power B. DC power C. Hertz D. Volt

3-3. The generator output frequency varies directly in proportion to ________.

A. inch-pounds of torque B. inertia of engine C. engine drive speed D. overspeed

3-4. What are the two types of ac generators?

A. Stationary and brushless B. Permanent magnet and oil-cooled C. Magnetic and three-phase D. Brush and brushless

3-5. To accomplish voltage regulation in ac generators, you control the strength of ________.

A. the RPM B. the magnetic field C. the output winding D. the two-separate phase

3-6. A three-phase generator contains how many separate power sources?

A. One B. Two C. Three D. Four

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3-7. What is the phase difference between voltages in a three-phase generator?

A. 110 degree B. 115 degree C. 120 degree D. 210 degree

3-8. What unit provides control voltages and power for voltage regulation?

A. Rectifier B. Exciter C. Exciter armature D. Permanent magnet generator

3-9. Feeder fault protection is provided by ________.

A. two three-phase differential transformers B. generator mechanical failure warning device C. control the rotor speed D. constant speed output

3-10. The device that provides the driving force for a generator is known as ________.

A. transformer B. prime mover C. mechanical governor D. inverter

3-11. What unit uses a dc motor to drive an ac generator and is used during emergency situations?

A. Permanent magnet generator B. Permanent magnet rotor C. Inverter D. Transformer

3-12. What is the only moving part in a transformer rectifier?

A. Inverter B. Armature C. Brush assembly D. Cooling fan

3-13. An autotransformer offers its greatest savings at a turn ratio of ________.

A. 1 to 1 B. 2 to 1 C. 3 to 1 D. 4 to 1

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3-14. What unit is used to connect meters to high voltage/current systems?

A. Autotransformers B. Brush C. Full wave rectifiers D. Instrument transformers

3-15. What are the three most common aircraft batteries in use today?

A. Wet-cell, vent-sump, and lead-acid B. Vented, cadmium, and silver­zinc C. Lead-acid, nickel-cadmium, and silver­zinc D. Storage, lead-acid, and cadmium

3-16. The ram-air generator is powered by ________.

A. air caused by aircraft moving through atmosphere B. air-pressure surrounding aircraft C. variable pitch blade D. three-phase emergency generator

3-17. The GTCP-95 centrifugal switch shuts down the unit at what percent RPM?

A. 35 B. 75 C. 95 D. 106

3-18. What are the two common methods used for voltage regulation?

A. Maintaining the voltage to the generator exciter winding and decreasing the load on the generator B. Varying the current to the generator exciter winding and maintaining a constant load on the generator C. Varying the current to the generator and varying the load on the generator. D. Increasing the voltage and the load on the generator exciter windings

3-19. During what conditions will a supervisory panel disconnect the generator from the load?

A. Voltage reaches 120/208 volts ac B. Underfrequency C. Engine lubrication system is on D. Engine overspeeds

3-20. What is the permanent magnet generator (PMG) output?

A. 39 volts B. 60 volts C. 90 volts D. 120 volts

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3-21. How much current is used to produce the desired magnetic field to generate a 120-/208-volt, 60-kVA load?

A. .5 amperes B. 1.0 amperes C. 1.2 amperes D. 1.7 amperes

3-22. With one phase load one-third more than the other two phases, the voltage between phases should not vary more than?

A. .5 volts B. 5 volts C. 120 volts D. 208 volts

3-23. The undervoltage protection circuit allows the generator to assume the load at what voltage?

A. 105 B. 120 C. 210 D. 300

3-24. What is the length of the undervoltage timing cycle?

A. 2-second period and a 1-second period B. 3-second period and a 1-second period C. 3-second period and a 4-second period D. 4-second period and a 6-second period

3-25. What voltage will the overvoltage protection circuit start a time delay?

A. Below 105 B. Between 110 and 120 C. Between 120 and 128 D. Above 129

3-26. What is a feeder fault?

A. A pump for supplying the generator with engine oil for cooling B. Circuit protection for both the operating generator and equipment it powers C. A condition where the current leaving the generator does not pass through the load. D. A voltage regulating circuit and control

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3-27. What is the power factor in ac circuits?

A. Resistance is the apparent power equal to the true power B. Currents circulate between the power capacitor and inductive load C. The ratio of true power to apparent power, and is equal to true power divided by apparent power D. The vector sum of capacitor current and total inductive load current is equal to line current

3-28. How are power-factor corrections accomplished?

A. By connecting a capacitor of the proper capacitance in parallel with the circuit B. By using a wattmeter to read the system C. By making connection as far as possible from the inductive load D. By ensuring the capacitor inductive load does not enter the line

3-29. What unit prevents external power that is not in tolerance from being applied to the F/A-18 aircraft?

A. Right power contactor B. Left bus tie contactor C. External power monitor D. External power contact or

3-30. When the left power supply fails in the F/A-18 aircraft, which components allow power to the left side of the dc bus distribution system?

A. Utility power receptacle and left 115vac bus B. Left and right bus tie current limiter holders C. Left 115vac bus and left primary dc bus D. Right power supply and the battery charger

3-31. During ground power switching operation, what is the condition of the left and right power contactors when external power is applied to the aircraft?

A. Energized B. On line C. Negative with respect to ground D. De-energized

3-32. What does the term "grounded system" mean?

A. When the grounded leg of the circuit does not connect to a good electrical conductor B. The ac essential and dc essential buses receive power C. One leg of the system connects to a common conductor D. When all the wires complete the equipment circuit to the ground network

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3-33. What does the letter "N" designates in wire numbering?

A. Unfiltered unregulated wire B. Reference wire for the primary circuit network C. Wire for the auxiliary network circuits D. Wire that completes the equipment circuit to the ground network

3-34. In the four-wire generator system, which lead connects to ground?

A. B phase B. Neutral C. De-energized D. Energized

3-35. Why is the grounded circuit more advantageous than the ungrounded circuit?

A. Circuits of different potentials and frequencies use a common ground B. It runs the power source directly to the loads C. Short circuits result when a bare spot on any ungrounded conductor touches ground D. Because it reduces overall weight by using fewer conductors

3-36. What is an advantage of the ungrounded circuit?

A. Troubleshooting is simplified and the impedance of the ground return path is lower than that of a run conductor B. There is a possibility that one circuit feeds another C. It prevents one circuit from feeding into another D. It requires more conductors than the grounded system

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CHAPTER 4 AIRCRAFT ELECTRICAL SYSTEMS

As an Aviation Electrician’s Mate (AE), you will work with many electrical systems and may work directly or indirectly with all other aviation maintenance ratings. For example, you may work with Aviation Machinist Mates (ADs) on power plant discrepancies or with Aviation Structural Mechanics (AMs) on electrohydraulic malfunctions. Consequently, you must be a well-rounded aviation technician. In this chapter, you will learn about various systems that AEs regularly maintain. Maintaining naval aircraft systems is the number one priority; teamwork and maximum effort will pay dividends in flight safety and mission accomplishment. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Recognize operating principles and construction features of aircraft lighting. 2. Describe the types of lamps used in aircraft lighting and their circuits. 3. Identify the types, purposes, and uses of aircraft internal lighting. 4. Recognize operating principles and characteristics of aircraft electrohydraulic and pneumatic systems. 5. Define terms used in aircraft environmental control systems. 6. Identify the components used in aircraft environmental control systems. 7. Explain the operating principles and features of aircraft temperature and pressure control systems. 8. Describe the operating principles and features of aircraft anti-icing system. AIRCRAFT LIGHTING SYSTEMS The lighting system in an aircraft serves two primary purposes; it provides specialized light sources outside the aircraft and illuminates the interior. Exterior lights provide illumination at night for navigation and formation flying. Other exterior light operations include signaling, landing, and anti-collision. The interior lighting provides illumination for instruments, equipment, crew stations, and cabins. In addition, the electrical, electronic, and mechanical systems use lights to indicate normal operation or a possible malfunction. Lights also show the position of the landing gear, bomb bay doors, etc. Various types and sizes of light assemblies are used on present-day naval aircraft. The lighting requirement governs the selection of a particular light assembly. Most light assemblies consists of a housing (fixture), a lamp, and a lens. Description and Types of Lamps Aircraft lamps are devices that provide sources of artificial light. The incandescent light is the most common type. It uses an electrical source to heat a filament until it is white hot. Normally, the source voltage is 26 to 28 volts alternating current (ac) or direct c urrent (dc). However, some lighting systems use lower voltage lamps and step-down 4-1

Figure 4-1 — Lamp bases. transformers to supply 3 to 6 volts to the lamps. The lower operating voltage allows the lamp filament to be larger, and it helps to reduce lamp failure due to vibrations. The parts of a lamp are the bulb, filament, and base. Incandescent lamps vary chiefly in electrical rating, base type, bulb shape, and bulb finish. The electrical rating of lamps is expressed as a combination of volts, watts, amperes, and candlepower. (Candlepower is the luminous intensity expressed in candles and used to specify the strength of a light source.) The lamp rating is found on either the base or the bulb of the lamp. On small lamps, an identifying number represents the electrical rating. This number is the same as the Military Specification (MS) dash number and can be found on the lamp base. The base types vary as to size, number of electrical contacts, and the method of securing in a socket. The most common types of bases are the single- or double- contact bayonet (push in and turn). These bases are used on aircraft since they lock in the socket and do not loosen because of vibration. Single contact bases are used in single-wire systems. In this system, one side of the lamp filament connects to the base and the other to the contact. Dual filaments use a common contact to the base. Single and double-contact lamps are not interchangeable. Some bases are of the screw type. They aren’t used often because they loosen easily. Other lamps have an indexing-type base that has offset index pins. (Figure 4-1) The index-type base is used to make sure that the lamp sits in the socket and the light shines in the proper direction.

Some lamps do not have a base; they solder directly onto a circuit board or permanently mount in control box panels. These lamps go through careful testing and selection so they last the life of the aircraft. Figure 4-2 shows one such lamp, the grain of wheat lamp. The bulb shape is shown by a combination of letters and numerals. The letter shows the bulb shape, and the number shows the approximate maximum diameter of the bulb, in eighths of an inch. The codes and the shapes they indicate for the more common glass bulbs are as follows: 4-2

Figure 4-2 — Grain of wheat lamp. Figure 4-3 — Common bulb shapes. G Globular GG Grimes globular S Straight T Tubular PAR Parabolic aluminum reflector R Reflector

By looking at the letter designations, you know that a bulb designated as T-6 is a tubular bulb with a ⁄ -inch diameter. A variety of sizes and shapes of bulbs are listed in the Defense Logistics Agency (DLA) Identification List. This list is available through your local supply support center. Figure 4-3 shows some common bulb shapes. Most aviation lamps are either clear glass or frosted on the inside. For a particular application, however, a bulb may be partially frosted; for example, to cut down emission of light in a particular direction. Another bulb may be partially silvered to prevent emission in a specified direction and/or to concentrate the light in other directions. Some applications call for colored bulbs; for example, in instrument illumination and safety lights. The letters just before the MS dash number shows bulb finish-R for red, SB for silvered bowl. If no letter is present, the lamp is clear glass or frosted. You can also provide colored lighting by using clear lamps with a colored lens cover. There are many special-purpose lamps in use on naval aircraft. Three of the most common types are listed below: 1. The parabolic light is a sealed-beam light used in the landing and taxi light systems. Signal lights also are used in sealed-beam lamps. 2. The midget-flange type of light is for use in instrument panels and control boxes. 3. Fuselage and signal lights use lamps having two filaments in parallel to provide fast signaling (smaller filaments heat and cool faster).

4-3

When a lamp fails, the replacement must be the same as the original lamp or an approved alternate. For those aircraft that carry onboard spare lamps it must be ensured that they are shock mounted. If they were not, they would probably fail earlier than the original lamps because cold filaments are subject to fatigue failure sooner than hot ones. You need to be sure that the glass bulb of the lamp is clear and free from grease and dirt. To help keep bulbs clean, do not touch the glass bulb with your bare hands, if possible. Exterior Lighting Many types of lights are used to meet the exterior lighting requirements of naval aircraft. The principal types of exterior lights are the navigation or position lights, anti-collision lights, landing lights, and formation lights. Figure 4-4 shows the components used in the exterior lighting system of a carrier-type aircraft. This figure shows the lights common to naval aircraft, but does not show every type of light in use on different aircraft. The lighting requirements vary from one aircraft to another, depending upon the aircraft type.

Figure 5-8 — Philadelphia rod set for target reading of less than 7000 feet. 4-4

Figure 4-4 — Exterior lighting.

4-5

NOTE The position for these navigation lights is on a stationary surface to the extreme left, right, and aft on helicopters. Navigation Lights (Position Lights) Navigation lights on the aircraft attract visual attention to its position and heading at night. A standard minimum set of navigation lights, meeting Federal Aviation Administration (FAA) requirements for light distribution and intensity, is on all military heavier-than-air aircraft. The standard minimum set of navigation lights for night operations consists of the following:  One red light on the tip of the left wing  One green light on the tip of the right wing  One white light on the tail, so it is visible over a wide angle from the rear

In some aircraft configurations, navigation lights burn steadily; in other configurations they burn steadily or flash about 80 flashes per minute. Fuselage Lights Fuselage signal lights are part of the aircraft navigation lighting system. They provide a method of visual signaling. When installed, two or more fuselage lights are necessary, one on the top and one on the bottom of the aircraft. If it is not practical to install the light on the bottom of the aircraft, such as on a seaplane or with a radome obstruction, you install two lights. Position one on each side as near the bottom of the aircraft as possible. Fuselage lights may burn steadily or flash at a constant rate. On some aircraft, manual keying of lights for signaling is available. Anti-collision Lights Anti-collision beacon lights are an FAA requirement for all aircraft. Their primary purpose is flight safety during daylight hours as well as at night. One type of anti-collision light consists of two 40-watt reflector-type lights and a red lens assembly. An alternating current (ac) or direct current (dc) motor rotates the bulb assembly, causing it to flash 80 to 90 flashes per minute. Slip rings provide electrical power to the bulbs. On another type of anti-collision light, the bulb is stationary. The flashing is caused by motor-driven reflectors in the light assembly. Most aircraft are now using strobe lights to provide anti-collision warning to other aircraft. This system has a 3,000 to 3,500 candlepower white light for day and 150 to 200 candlepower red light for night. Both flash at 60 flashes per minute. Aircraft equipped with in-flight fueling tanker capabilities can turn off the lower anti- collision light during delivery of fuel to another aircraft. Tanker aircraft use a bluish- green lens over their anti-collision lights to identify fueling capability to other aircraft in need of fuel. 4-6

Figure 4-5 — Retractable landing lights. Many aircraft have been retrofitted with Infrared (IR) anti-collision light assemblies which provide anti-collision warning to aircraft operating with night vision goggles or other infrared detection systems. IR anti-collision light assemblies consist of a ring housing with six IR diodes located around the outer edge of the ring. Landing Lights Modern naval aircraft have high candlepower landing lights to illuminate the landing strip. Multi-engine aircraft usually have a landing light mounted on each wing. Single- engine aircraft use only one landing light, which normally mounts on the port wing. Landing lights are usually retractable (Figure 4-5). Sealed-beam lights with a rating of 28 volts, 600 watts are used. They use split-field series dc motors or ac induction motors to provide drive operation. The 28-volt ac is provided by an autotransformer located within each light assembly. In the retracted position, movable lights are flush with the undersurface of the wing. In some aircraft, you can install a landing light in the nose wheel fairing door. Other aircraft have a fixed landing light, mounted in the leading edge of the wing. As you read this section, refer to Figure 4-5 view A. You are going to learn about the operating principles of a typical retractable landing light.

4-7

The landing light switch on the pilot’s lighting panel controls the landing light motor. When the switch is in either the EXTEND or RETRACT position, power goes simultaneously to the magnetic brake (releasing the brake) and the drive motor. Limit switches open the motor circuit when the light has reached its limit of travel in either direction (extended or retracted position). This light can be put in any position between its travel limits by turning the switch off. When the power to the motor is off, the gear train holds the light in that position. The lamp illuminates by a sliding contact after the light assembly travels downward about 10 degrees. Illumination continues until the lamp again reaches the 10-degree position when traveling in the upward direction. The lamp will remain lighted in any extended position past the 10-degree position, regardless of power application to the control motor. The pilot can adjust the angle of the light beam to fit the operation. A switch lets the pilot turn the lamp off while the light is in any position, as shown in Figure 4-5, view B. The landing light assembly design lets you adjust the maximum extended position for each particular aircraft installation. The light opens to an extended position of 73 degrees, ±3 degrees, from the retracted position. The light assembly is capable of being extended to positions ranging from 50 degrees to 85 degrees from the retracted position. Technicians must not let landing lights overheat or be damaged by extending them against ground support equipment when they are ground tested. You should NEVER look directly at an illuminated landing light because it can cause permanent eye damage. Approach Lights The carrier aircraft approach light systems give the pilot and Landing Signal Officer (LSO) positive indication of safe or unsafe landing configurations. All shipboard naval aircraft have approach lights mounted so they are clearly visible to the LSO. Installations of the approach lights vary with aircraft. Older aircraft have them in the port wing leading edge. Modern aircraft, such as the EA-6B and F/A-18 (Figure 4-6, frame 1), have the approach lights on the nose landing gear or landing gear door. Most jet aircraft have a three-lamp Angle of Attack (AOA) approach light assembly and AOA indexer light assembly used simultaneously. These type aircraft are carrier-based. The one described herein is for the F/A-18. The AOA Approach Light Assembly is three lights (red, amber and green) which correspond to the aircraft’s AOA indication and incorporates a test switch to allowing testing the lights. Additionally, an AOA Indexer Light assembly provides a visual indication to the pilot during landing. These lights are varied in brightness by an AOA control knob on the Head-up Display ( HUD) control panel and a lights test is provided by a lights test switch on the Interior Lights Control panel. When the red light is on, it shows that the angle of attack of the aircraft is low. When the green light is on, it shows that the angle of attack of the aircraft is high. When the amber light is on, the angle of attack is optimum for the landing approach.

4-8

Figure 4-6 — Angle-of-Attack components.

The AOA Approach Lights work in conjunction with the crew station Angle of Attack Indexer lights which give the pilot angle-of-attack information. These lights are mounted directly to the left side of the Heads-Up Display (HUD) Combiner Assembly (Figure 4-6, frame 2). They are clearly visible to the pilot, as not to obstruct vision. Power for the Approach lights system is provided by three separate power distribution busses, two 28 volts direct current (vdc) and one 115 volts alternating current (vac) and p rotected by the Approach lights circuit breakers. Both the approach lights and the indexer lights are operated by three discrete signals (High, Low and On Speed) from the Flight Control Computer B (Channel 4) through a Relay Module Assembly (RMA). The Flight Control Computer provides a ground to energize the corresponding AOA relay inside the RMA, depending on the Angle of Attack Transmitter’s output. At very low angles of attack, a red-colored inverted V illuminates in the range of 0-6.9 degrees or lower (Figure 4-6, frame 2). This warns the pilot to increase the aircraft angle of attack. At slightly low angles of attack, both the inverted V and circular symbol (doughnut) illuminate in the range of 6.9 to 7.4 degrees (Figure 4-6, frame 3). At optimum desired a ngles of attack, the amber-colored doughnut illuminates in the range of 7.4 to 8.8 degrees (Figure 4-6, frame 4). At slightly high angles of attack, both the V and doughnut illuminate in the range of 8.8 to 9.3 degrees (Figure 4-6, frame 5). At very high angles of a ttack, a green-colored V symbol illuminates in the range of 9.3 degrees and up (Figure 4-6, frame 6), warning the pilot to decrease the aircraft angle of attack. Additionally, a movable Angle-of-Attack scale will move vertically as the aircraft pitches. The center of 4-9

NOTE Manufacturers sometimes identify lights having the same purpose as formation lights by another name such as join- up lights or strip lights. NOTE Numbers (Degrees) in HUD Combiner indicate local AOA as seen at Angle of Attack Transmitter by maintenance personnel. the scale indicates the best approach angle with respect to the velocity vector. Scale is limited to the HUD field of view.

The indexer information corresponds to the approach lights assembly information displayed to the LSO. Approach lights show the LSO three conditions;  Approach Angle- of-Attack  Landing Gear is down and locked and Weight- Off-Wheels (W-OFF-W)  Arresting Gear is extended An Approach Light flasher and Approach Light control relay cycles power to the approach light assembly causing the approach lights to flash when the Hook Bypass Switch is set to the carrier position and the arresting gear hook is not down. An Annunciator Dimming relay provides circuit switching to provide dimming of the Indexer lights when operating at night for night or night vision modes of operation. A test switch mounted to the AOA Approach lights, when actuated, tests the AOA approach light assembly and turns on the three AOA Approach lights. Formation Lights Naval aircraft have formation lights for night formation flying. In a typical formation light installation, wingtip formation lights mount within the upper and lower surfaces of the aft section of each wingtip. Translucent diffusing windows mount flush with the wingtip surface above and below each light. The right-hand (starboard) light covers are green and the left-hand (port) light covers are red. The lamp assembly is accessible through the lower wingtip cover plate. Fuselage formation lights installations (Figure 4-4, callouts 5, 6, 7, 8) are in box assemblies mounted on each side of the fuselage and outer wings. Each has a plastic window for light emission and the lamp is accessed by removing the box cover. These lights connect in parallel with the wingtip formation lights; therefore, they illuminate simultaneously.

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Figure 4-7 — Typical helicopter lighting. Taxi Lights Taxi lights help the pilot maneuver the aircraft before and after flight. On aircraft having a nose wheel, the taxi light assembly mounts on the movable strut so the light turns with the wheel. Light installation provides maximum visibility for the pilot and co-pilot. (Figure 4-4, callout 10). In-Flight Refueling Probe Light Most modern carrier aircraft have limited range due to their small fuel capacity. To increase range and flight time, day and night in-flight refueling is necessary. Most naval shipboard aircraft have provisions for in-flight refueling. For night refueling, in-flight refueling (IFR) probe lights (Figure 4-4, callout 12) mount on the fuselage forward of the probe or on the probe itself. The light lens is usually red or white in color and illuminates the refueling probe and drogue at night. Hover Lights and Spotlights In helicopter installations, hover lights illuminate the area directly beneath the aircraft (Figure 4-7, callout 2). These lights serve several purposes, including landing and search and rescue. A spotlight can be mounted in the nose of a helicopter (Figure 4-7, callout 3), and it can extend or retract. An ON/ OFF/ RETRACT switch on the pilot’s collective stick and a spring loaded, four-way thumb switch marked EXTEND/ RETRACT/ LEFT/ RIGHT control the spotlight. It can rotate through 360 degrees of azimuth.

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Interior Lighting Various types of lights and lighting systems are used for interior illumination of naval aircraft. Almost all lights and lighting systems fall under one of the following types:  Instrument lighting  Crew station lighting  Cabin and passageway lighting  Indicator lights An important consideration in interior lighting of aircraft is the prevention of undue eyestrain. Eyes adjust slowly to changing light intensities, which can cause fatigue or eyestrain. Aircraft lighting designs produce as little discomfort as possible. As an AE, you maintain the aircraft lighting systems. You should follow the specifications when replacing fixtures and lamps. The following general considerations are taken into account when working with the interior lighting of aircraft:  Use lenses to diffuse light that lies within the pilot’s field of vision.  Eliminate all bright spots of light, direct sources of light, and reflections.  Use sparingly any surface that reflects light, such as chromium or nickel.  Use quick -change lighting fixtures so that lamps may be changed rapidly. For specific aircraft lighting and fixture specifications refer to the applicable Maintenance Instruction Manual. Instrument Lights The first use of artificial light in aircraft was for the instrument illumination. Operating modern aircraft depends on instruments; therefore, instrument lighting is important. There are different methods of instrument lighting, and selecting the best is a difficult decision. No matter what system of lighting is in the aircraft, the light must not be visible outside the aircraft. Indirect (mask) lighting is desirable because it doesn’t produce objectionable reflections. The lamps mount in the instrument panel. The panel has a reflector cover (mask), which has openings in it for observing the instruments. The light reflects until it becomes diffused and floods the entire panel. Even though this system produces satisfactory lighting, it is not in common use today because of space and weight requirements. Another method of instrument lighting, commonly called post lighting, uses installation of specially adapted shields in front of the instruments. Small lamps in red-filtered sockets mount in the front surface of the cover shields, which cast light down and onto the instruments. These shields mount around the instrument for vision and are flanged to direct the lights properly for dial illumination. Rheostats control the intensity of light in most aircraft (Figure 4-8). The rheostat is a variable resistor used to limit current through the circuit. When the pivoting arm reaches the high-resistance end of the rheostat, it slips off the contact surface and breaks the circuit. This arrangement is a rheostat switch.

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Figure 4-8 — Rheostat switch.

A typical aircraft instrument lighting system is shown in Figure 4-9. The lighting equipment consists of edge-lighted control panels, individual lights for instruments, and red floodlights for overall lighting. The edge-lighted panels provide diffused lighting for the control and indicator panels. Some instruments currently have an integral light source, and all instrument designs will soon have integral lighting. Since individual lights for instruments are compatible with the integrally lighted instruments, the two types of light sources may be intermixed. The crew station interior lighting system provides primary and secondary lighting for the instruments, instrument panels and consoles. Displays and indicators come on to show information in both day and night environment. Edge-lighted control and indicator panels (Figure 4-9) have a non-gloss, black background with white lettering for maximum contrast. The black edge face is mounted so light is diffused through the plastic panel lettering and knob openings. The same light is also infused into the knobs to show position in night time environment. Bulb or edge panel replacement is relatively simple.

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Figure 4-9 — Interior lighting. The MODE switch, located on the INTR LT panel, is used to select one of three crew station lighting modes; allowing the pilot to optimize interior lighting for ambient light conditions (Figure 4-9, frame 1).  NVG –Selecting the Night Vision Goggle NVG position reduces the brightness range for the warning, caution, and advisory lights, the Up Front Control Display (UFCD), Multi-Purpose Color Displays (MPCD), and Electronic Fuel Display (EFD). This disables the integral console lights and the white floodlights and enables six NVG compatible floodlights to illuminate the consoles (Figure 4-9, frames 11, 12, 13, 14, 15 and 16).  NITE –Selecting the NITE position reduces the brightness range for the warning, caution, and advisory lights, the UFCD, MPCD, and EFD (Figure 4-9, frame 2). 4-14

 DAY - Provides the maximum brightness range for all interior lighting.

The INST PNL lighting knob (Figure 4-9, frames 3 and 4), located on the INTR LT panel, is used to control the brightness of the integral lighting varying 0-5 VAC to the Cockpit Electric Light Control for the main instrument panel and the standby magnetic compass. Clockwise rotation of the knob increases main instrument panel lighting intensity from the OFF to BRT positions. The strobe function of the SHOOT light is disabled when the instrument lights are on. The FLOOD knob (Figure 4-9, frames 5 and 6), located on the INTR LT panel, is used to control the brightness of the white cockpit floodlights. Eight floodlights are provided for secondary lighting; three above each console and one on either side of the main instrument panel. Clockwise rotation of the knob increases floodlight intensity from the OFF to BRT positions. The FLOOD knob and all white floodlights are disabled in the NVG mode. There is no brightness control for the six NVG floodlights. The CONSOLES lighting knob (Figure 4-9, frames 7 and 8), located on the INTR LT panel, is used to control the brightness of the integral lighting varying 0-5 VAC to the Cockpit Electric Light Control for the left and right consoles, the hydraulic pressure gauge, and both circuit breaker panels. Clockwise rotation of the knob increases console lighting intensity from the OFF to BRT positions. The CONSOLES knob and integral console lighting are disabled in the NVG mode. The LT TEST switch, located on the INTR LT panel, is spring loaded to the OFF position. The switch is used to test important crew station lighting to verify bulb integrity prior to flight. The switch requires AC electrical power to operate.  TEST- Powers all operating warning, caution, and advisory lights, the AOA indexer lights, the integral background lighting on the EFD (BINGO, MODE, and BRT), changes MENU to ENG on the DDIs, provides a CHECK SEAT caution in the F/A-18E, and annunciates the landing gear warning tone.  OFF – Lights test off. The WARN/CAUT knob, located on the INTR LT panel, is used to control the brightness of the warning, caution, and advisory lights in the reduced brightness range. Clockwise rotation of the knob increases warning, caution, and advisory light intensity from the OFF to BRT positions. The brightness is full bright in the DAY mode and in the reduced brightness range in the NITE and NVG modes. Following a power interruption in either DAY or NITE mode, the warning, caution, and advisory lights default to the maximum brightness range. Following a power interruption in the NVG mode, the warning, caution, and advisory lights remain in the reduced brightness range. The CHART light knob (Figure 4-9, frames 9 and 10), located on the INTR LT panel, is used to control the brightness of the NVG compatible chart light. The chart light is located on the canopy bow at the 11:00 o’clock position and rotates on two axes. Clockwise rotation of the knob increases chart light intensity from the OFF to BRT positions. In some aircraft, grain-of-wheat lamps have permanent mounts in the control-box panels. Having been aged and carefully selected for output and reliability, the lamps should last for the lifetime of the aircraft. 4-15

Figure 4-10 — Crew station utility light. Cavities and shallow trenches in the panel back accommodate the panel connector, lamps, and route lamp leads to the panel connector. The connector, lamps, and leads are then potted into the back of the panel with a clear plastic potting material. Normally, the lamps are in parallel, and failure of one lamp will not appreciably degrade the panel lighting. The main advantages of embedded lamps are long life, ruggedness, resistance to aircraft vibrations, and better illumination. Crew Station Lights The term crew station lighting is rather broad. Its meaning varies, depending on the type aircraft being described. In fighter-type aircraft, it may be interior lighting that consists of individually lighted instruments and switches, lighted control panels, and necessary floodlights. It includes the interior lighting just mentioned as well as many other special lighting assemblies. A much used lighting device is the small incandescent spotlight known as a utility light assembly. These assemblies, installed at crew stations, are in a position where they provide illumination of equipment that the crew member uses. The light from the lamp assembly can focus in either a small spot or in a wide beam. It also has a red filter for night use. Figure 4-10 shows one type of light in which an ON-OFF switch and intensity control rheostat control light operation. Crew station extension light assemblies provide crew members with an extension light for reading maps or illuminating small areas. These assemblies consist of a connecting cord, switch, and lamp housing assembly. By adjusting the assembly, you can change the size of the light beam. Indicator Lights Crew station personnel get aircraft operating status from various indicator (warning, c aution, and advisory) lights in the crew station. These lights have many purposes, such as showing the position of the landing gear, arresting hook, wings, and bomb bay doors. They can also show low oil pressure, equipment over temperature, generator failure, and other data. The construction of indicator lights varies, depending upon the particular job they perform. They mount in the aircraft where they are easily noticeable when glowing. It is important that bulb replacement be quick and easy since some lights relay vital information concerning safety of flight. Whenever practical, assembly design allows bulb replacement in flight without the use of tools. Most warning light designs have a push- 4-16

Figure 4-11 — Legend-type lights. to-test feature or a test switch to determine if the bulb is good. The test switch energizes various test relays, which, in turn, either provide power or a ground circuit for energizing the lights. This allows the bulb to be checked without operating the equipment. Legend-type lights (Figure 4-11) show specific functions on the lens surface. They are prominent on late model aircraft.  The warning lights are red. They warn the crew of an emergency or unsafe operating condition, which requires immediate corrective action.  The caution lights are yellow. They alert the crew to a minor malfunction or impending dangerous condition requiring attention, but not necessarily immediate corrective action.  The advisory lights are green. They show the crew a safe or normal configuration, or a performance condition.

Indicator lights must be bright enough to see during daylight operation but not too bright to cause eyestrain at night. Brilliance control is obtained by connecting resistors in the lighting circuit, and by placing dimmer caps on the lights. Other ways of control are special adaption of edge lighting, and special types of lenses that dim by twisting the lens. An example of the use of indicator lights is the system used to show high oil temperature. A thermoswitch in the oil return line closes when the temperature reaches above normal. This usually provides a path to ground illuminating the OIL HOT light. The light goes out only when the oil cools to normal temperature. Another application of the use of an indicator light is the landing gear unlocked warning light. In some aircraft, a red light glows in the translucent landing gear control lever when the gear is not in either the up or down position. 4-17

AIRCRAFT ELECTROHYDRAULIC AND PNEUMATIC SYSTEMS The word hydraulics is from the Greek word for water. Hydraulics originally meant the study of physical behavior of water at rest and in motion. Today the meaning includes the physical behavior of all liquids, including hydraulic fluid. Hydraulics is the science of liquid pressure and flow. In its application to aircraft, hydraulics is the action of liquids, forced under pressure through tubing and orifices, to operate various mechanisms. The primary concern in working with hydraulics is to control the flow of fluid. You do this by using solenoids that simply turn on, shut off a flow of fluid, or change flow direction. In some cases, electrical devices may schedule a precise amount of fluid flow. In any case, you must understand the characteristics of liquids. You should recall from chapter 1 that a liquid has no definite shape and conforms to the shape of its container. Also, a liquid can be only slightly compressed. Another fact you should recall is the ability of a liquid to transmit pressure. The physics of fluids and basic hydraulic principles are contained in Fluid Power, NAVEDTRA 14105. Study the first two chapters of this publication with this chapter if you need to review basic hydraulic principles. Although some aircraft manufacturers make greater use of hydraulics than others, the hydraulic system of the average modern aircraft performs many functions. Among the units commonly operated by hydraulics are the landing gear, wing flaps, speed brakes, wing folding mechanisms, flight control surfaces, canopy, bomb bay doors, wheel brakes, and arresting gear. Basic Hydraulic System All hydraulic systems are essentially the same, regardless of their function. Hydraulics are used on the farm, in industry, aboard ship, and many other places as well as in aircraft. Regardless of its application, each hydraulic system has a minimum of four components— reservoir, pump, selector valve, and actuating unit—plus lines through which the fluid flows. Figure 4-12 shows a basic hydraulic system, with the four essential components and their relationship within the system.

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Figure 4-12 — Basic hydraulic system, hand-operated pump.

The reservoir provides storage for a supply of fluid for operation of the system. It replenishes the fluid of the system when needed, provides room for thermal expansion, and normally provides a means for bleeding air from the system. The pump creates a flow of fluid. The pump in the pictured system is hand operated; however, aircraft systems have engine-driven or electric-motor driven pumps. There are two types used in naval aircraft—the piston type and the gear type. One aircraft may incorporate both types due to the requirements of the various hydraulic systems. The selector valve directs the flow of fluid. These valves actuate either manually or by solenoid; they may operate directly, or they may operate indirectly with the use of mechanical linkage. The actuating unit converts fluid pressure into useful work. The actuating unit may be an actuating cylinder or a hydraulic motor. An actuating cylinder converts fluid pressure into useful work by linear/reciprocating-mechanical motion. A hydraulic motor converts fluid pressure into useful work by rotary-mechanical motion. In Figure 4-12 you can trace the flow of hydraulic fluid from the reservoir through the pump to the selector valve. With the selector valve in the #1 position, fluid flow created by the pump is through the valve to the right-hand end of the actuating cylinder. Fluid pressure then forces the piston to the left. As the piston moves left, fluid on that side of the piston flows out, up through the selector valve, and back to the reservoir. With the selector valve in the #2 position, fluid from the pump flows to the left side of the actuating cylinder, reversing the process. You can stop piston movement at any time by moving the selector valve to neutral. In this position, all four ports close, and pressure is equal in both working lines. Additions may be made to the basic system to provide additional sources of power, operate additional cylinders, make operation more automatic, or increase reliability. These additions are all made on the framework of the basic hydraulic system diagram shown in Figure 4-12. 4-19

Electrohydraulic Systems There are two types of electrically controlled, hydraulically operated components: selector valves that start, stop, or change the direction of a fluid flow (similar to a switch in an electrical circuit); and control valves that schedule fluid flow (much like a potentiometer controls current flow). Each of these components is manufactured in varying degrees of complexity, depending on its use. Each manufacturer uses its own identification for its components, such as transfer valve, engagement valve, servo valve, etc. Aircraft hydraulic systems normally operate at a pressure of 3,000 PSI; therefore, each component must be capable of operating at this pressure. You should observe all applicable safety precautions when working with high-pressure hydraulic or pneumatic systems. Automatic Flight Control System (AFCS) provides automatic operation of flight control surfaces through electrical input and hydraulic control to maintain aircraft heading, altitude, and attitudes as selected by the pilot. There are many varieties of AFCS systems including both digital and analog systems such as the AN/ASW-31 analog and AN/ASW-60 Digital Computer System. For our purposes we will be referring to the analog AN/ASW-31 system. Hydraulic Surface Control Booster System The hydraulic surface control booster system contains both selector and control components. The AFCS engagement valve is a selector valve. It either applies pressure to or removes pressure from the AFCS portion of the booster. The hydraulic transfer valve controls the direction and amount of fluid to the booster system. AFCS ENGAGEMENT VALVE – The AFCS engagement valve (Figure 4-13) is shown in disengage or relaxed position. No voltage is at the coil, and 3,000 PSI hydraulic pressure pushes the solenoid piston to the left. This action allows the hydraulic fluid to flow to the right side of the spring-loaded piston. With the same pressure on both sides, the spring-loaded piston moves to the left, preventing any further flow in the system.

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Figure 4-13 — Hydraulic surface control booster with no mechanical input and AFCS disengaged.

When AFCS engages (Figure 4-14), a 28-volt dc signal goes to the solenoid coil to drive its piston to the right. This action relieves hydraulic pressure from the right side of the spring-loaded piston. This allows pressure on the left side to move the piston to the right, compressing the spring. Hydraulic pressure then flows to the other AFCS components of the boost system. 4-21

Figure 4-14 — Hydraulic surface control booster with no mechanical input and AFCS engaged.

Hydraulic Transfer Valve With AFCS engagement, a dc control signal goes from the AFCS computer to the transfer valve. In the static state (zero signal), dc flow through the transfer valve coil is equal, and the plunger mechanism remains in the centered position (with help from the centering springs). Hydraulic pressure from the engagement valve flows to the top land on the hydraulic transfer valve piston. Fluid also flows through a small orifice to the bottom land of the same piston. Pressure to the bottom of the piston is regulated by controlling the amount of fluid leaving the modulated pressure chamber. If chamber pressure is 1,500 PSI and one drop of fluid leaves the chamber for every drop that enters, the pressure remains constant. To increase pressure, an unbalanced dc voltage on the coil causes the 4-22

plunger to block the return orifice for an instant so more fluid enters the chamber than leaves. Conversely, an unbalanced dc signal from the AFCS computer that causes the plunger to rise for an instant lowers chamber pressure. With a zero input signal from the AFCS computer, the transfer valve piston is in the centered position (Figure 4-13). The pressure acting on the smaller surface area of the piston’s upper land and the modulated pressure acting on the larger area of the piston’s lower surface are in balance, keeping the piston centered. With an electrical signal input, as shown in Figure 4-14, a higher pressure in the modulated chamber causes the piston to rise. This action allows fluid to flow to another component of the boost system. The larger the electrical signal input, the more pressure difference on the piston; therefore, the more fluid flow. Booster System Operation The booster operates in two separate modes— manual and AFCS. In the manual mode, control surface deflection desired by the pilot starts by a lateral movement of the control wheel (Figure 4-13). This small movement of the control wheel transmits through the feel lever and feel rod to the control lever. The control lever is free to move about its pivot, and a centering spring holds the modulating piston in place. Movement of the control lever displaces the main control valve, porting fluid from two hydraulic systems to the hydraulic actuator piston. (The main control valve and hydraulic actuator work identically, whether in manual or AFCS modes.) When the hydraulic actuator piston moves, both the feel lever and the power arm posi tion the control surface and the control wheel to the position the pilot selects. The control surface moves faster as the pilot applies more pressure to the control wheel. The direction of the pressure determines the direction of control surface movement. If there is no pressure at the control wheel, the main control valve centers (Figure 4-13). Trapped fluid holds the hydraulic actuator piston in position. The manual operation of the dual shutoff and bypass valves is from the flight station. If there is a loss of hydraulic pressure or a malfunction in the boost system, the flight crew pulls the handle. Operation of the valves shuts off hydraulic pressure to the hydraulic actuating piston and opens both working lines to each other. This prevents hydraulic lock in the actuator piston and allows control surface movement without the aid of the boost system. Trapped hydraulic pressure in the modulator piston aids the mod piston centering spring in holding the piston centered. This allows positive control of the main control valve by the control lever. Upon AFCS engagement, the engagement valve ports hydraulic fluid to the hydraulic transfer valve and to the engagement piston. Movement of the engagement piston locks the control lever in its centered position and prevents mechanical inputs from the control wheels. This piston can, however, be overpowered by about 15 pounds of force on the control wheel in an emergency. Electrical inputs from the AFCS computer change to hydraulic fluid pressures in the hydraulic transfer valve, and the pressure controls the modulator piston. The modulator piston changes the hydraulic pressure into linear mechanical motion, which repositions the main control valve. Movement of the main control valve ports hydraulic fluid to the actuator piston to position the control surface and control wheel as previously described. 4-23

NOTE The cause of asymmetry must be determined and the trouble corrected before the reset operation can be accomplished. Electrical devices on the booster provide information inputs to the AFCS. The modulator piston linear transducer tells the AFCS the rate of control surface movement, and the surface position synchro transmitter provides position information. Read chapter 8 for more information about automatic flight control system operation. Wing Flap Systems The Wing Flap Systems are made up of the Wing Flap Asymmetry System and the Wing Flap Position Indicator System. They are installed for the detection and automatic arrest of asymmetry (unequal degree of extension) of the wing flaps and to indicate the amount of wing flap extension from 0 to 100 percent. Wing Flap Asymmetry System The Wing Flap system includes: two Wing Flap Asymmetry Detectors, one on each wing rear beam; a chain and cable system attached to each wing flap to drive the camshaft within the Asymmetry Detector; a Flap Asymmetry Dual Hydraulic Solenoid Valve Assembly (Wing Flap Asymmetry Shutoff Valve) in the hydraulic service center upstream of the main drive assembly; a Wing Flap Brake Assembly outboard of the actuator; a WING FLAP BRAKE RELAY in the hydraulic service center; a WING FLAP ASYMMETRY RELAY in the hydraulic service center; and a FLAP ASYM caution indicator on the center instrument panel in the flight station. On aircraft incorporating Airframe Change (AFC) 475, the WING FLAP ASYMMETRY LIGHT RELAY, the WING FLAP ASYMMETRY TEST/ RESET SWITCH, and the SHUTOFF VALVE TEST LIGHT are also located in the hydraulic service center.

Wing Flap Asymmetry Operation The cable system in the wing rotates the independent switches at identical speed as long as the wing flaps move symmetrically. In normal operation, the cams in the Asymmetry Detectors are timed to operate the six switches in each Asymmetry Detector so that closure of a switch on one side never occurs while the corresponding switch in the other Asymmetry Detector is closed. In the event of a mechanical failure of a type which would allow only one flap to stop when the mechanism is in motion or only one to start when motion is initiated, corresponding switches in each Asymmetry Detector would be closed at the same time, completing a circuit through them to the WING FLAP ASYMMETRY RELAY which controls the WING FLAP BRAKE RELAY and the Wing Flap Asymmetry Shutoff Valve. The Wing Flap Brake Relay completes the circuit to the Wing Flap Brake assemblies at the outboard actuators. This would immediately cutoff the hydraulic fluid supply to the main drive and the flaps would stop at the point of failure. Wing Flap Brake assembly engagement would prevent aerodynamic forces from retracting the flap with the broken line of drive. The Wing Flap Asymmetry Shutoff valve is self-latching and must be pressed to rearm the system.

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Wing Flap Asymmetry Test/Reset Operation The Wing Flap Asymmetry Test/Reset Switch will allow a confidence check of the Wing Flap Asymmetry Shutoff valve and Shutoff Valve Test light without tripping the Wing Flap Asymmetry system. At the hydraulic service center, the Wing Flap Asymmetry Test/Reset switch is set and held to TEST SHUTOFF VALVE. The circuit is completed to the Wing Flap Asymmetry Shutoff valve which shuts off hydraulic oil pressure to the Wing Flap Control valve. The circuit is also completed to the Shutoff Valve Test light and the FLAP ASYM caution indicator on the center instrument panel. The RESET position of the Wing Flap Asymmetry Test/Reset switch electrically will reset the Wing Flap Asymmetry relay after a tripped condition. Wing Flap Position Indicator System The system includes: one Wing Flap Position Transmitter located in the hydraulic service center and mounted on the Flap Drive Control Mechanism; and one Wing Flap Position Indicator located on the copilot instrument panel in the flight station. Wing Flap Position Indicator System Operation The Position Indicator, located on the copilot instrument panel, provides indications in percent of flap travel from 0 to 100 percent. With the Position Indicator pointer at T.O. & APPROACH (takeoff position), 77 percent of flap extension is indicated, and at LAND position, 100 percent of flap extension is indicated. The Position Indicator receives an electrical signal from a synchro transmitter located in the hydraulic service center on top of the Rudder Boost Shutoff Gearbox mounted atop the Flap Drive Control Mechanism. A splined shaft that extends from the Position Transmitter senses any movement of the wing flaps. As the Flap Drive Mechanism rotates, the Rudder Boost Shutoff Gearbox rotates, and the splined shaft of the Position Transmitter causes the synchro transmitter to send a signal to the Position Indicator, causing the Position Indicator pointer to maintain the exact wing flap position on the dial. Wing Flap Asymmetry System Functional Signal Flow Description The following paragraphs provide signal flow description shown in Figure 4-15.

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Figure 4-15 — Wing flap asymmetry system functional signal flow description.

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The LATCH RELEASE & SOV circuit breaker on the MAIN DC BUS circuit breaker panel supplies the two coils (set and reset) of the WING FLAP ASYMMETRY RELAY with 28 vdc. Whenever the right and left wing flaps are extending or retracting asymmetrically with respect to each other, the WING FLAP ASYMMETRY RELAY is energized from switch contacts in the left and right wing Asymmetry Detectors. The closed contacts of the WING FLAP ASYMMETRY RELAY provide a path for 28 vdc from the LATCH RELEASE & SOV circuit breaker energizing the WING FLAP BRAKE RELAY, the WING FLAP ASYMMETRY LIGHT RELAY, and the Wing Flap Asymmetry Shutoff Valve, through the WING FLAP ASYMMETRY TEST/RESET SWITCH. The Wing Flap Asymmetry Shutoff Valve completes the ground return circuit to turn on the SHUTOFF VALVE TEST LIGHT and to the Indicator Lights Control to turn on the FLAP ASYM caution indicator on the center instrument panel. With the WING FLAP BRAKE RELAY energized, 28 vdc from the L & R BK FLAP circuit breaker on the MAIN DC BUS circuit breaker panel is sent to the Wing Flap Brake Assemblies as a wing flap brake signal to actuate the solenoid. Each solenoid releases a locking mechanism, allowing a spring-loaded brake to set on their respective wing flap torque tube. The energized Wing Flap Asymmetry Shutoff Valve shuts off the hydraulic oil pressure and return lines to the wing flap drive motors. The WING FLAP ASYMMETRY LIGHT RELAY provides a parallel ground return path to turn on the FLAP ASYM caution indicator in case of a Wing Flap Asymmetry Shutoff Valve failure. The SHUTOFF VALVE TEST LIGHT remains off. When set and held to TEST SHUTOFF VALVE, the WING FLAP ASYMMETRY TEST/RESET SWITCH routes 28 vdc from the LATCH RELEASE & SOV circuit breaker on the MAIN DC BUS circuit breaker panel to energize the Wing Flap Asymmetry Shutoff Valve. From the ground return inputs, both the FLAP ASYM caution indicator on the center instrument panel and SHUTOFF VALVE TEST LIGHT are turned on from the Wing Flap Asymmetry Shutoff Valve. The Wing Flap Brake Assemblies are not actuated. The reset coil of the WING FLAP ASYMMETRY RELAY is energized by setting the WING FLAP ASYMMETRY TEST/RESET SWITCH to RESET. The Wing Flap Brake Assemblies and the Wing Flap Asymmetry Shutoff Valve must be reset manually. Landing Gear System Most naval aircraft have hydraulically actuated, electrically controlled, retractable landing gear. Normally, locking the landing gear in the retracted or extended position is automatic. Figure 4-16 is an electrical schematic of the landing gear control circuit of a patrol type aircraft.

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Figure 4-16 — Landing gear control circuit. Circ uit Operation The landing gear circuit includes an electrical, solenoid-operated selector valve to control hydraulic actuation of the landing gear. Current goes to the landing gear selector valve through two single-pole, double-throw (SPDT) switches. These switches operate b y a cam on the landing gear control lever, permitting current to flow to either the up or down coil of the valve. Coil selection depends on control lever position. The landing gear control circuit also includes electrical control for emergency extension of the nose gear with emergency hydraulic system power. A center-off switch (SPDT type) provides control of a double solenoid-actuated hydraulic selector valve. The center-off position of the switch is the normal position during operation of the landing gear with the main hydraulic system. The down position selects emergency hydraulic system power for nose gear extension. The bypass position is for use only during retraction of the nose 4-28

gear with main hydraulic power after extension by emergency hydraulic power. The bypass also allows for release of emergency system hydraulic pressure any time you desire. Additional Circuits In some aircraft, the landing gear control circuit also controls the supply of power to the propeller-reversing circuit (through left and right main torque-link switches) and to the stores release circuit (through left main torque-link switch). The left main gear torque- link switch supplies power to energize the landing gear lever locking solenoid. The weight of the aircraft must be off the landing gear shock strut before you move the landing gear control lever from the wheels down position. The solenoid will be de- energized by the up sense switch of the left main gear after the gear retracts. This switch energizes a relay, whose normally closed contacts are in series with the solenoid circuit. The left and right main gear torque-link switches (parallel connected) are in series with the power to the throttle lever-locking solenoid to prevent the throttles from being placed in reverse propeller range. When one of the torque-link switches actuates (by aircraft weight compressing one or both main landing gear strut oleos), the throttle levers come into the reverse pitch range. The left main landing gear torque-link switches (series connected) also open the stores release circuit. This prevents accidental release of wing station external stores when the weight of the aircraft has compressed the shock strut oleo. A solenoid mechanically prevents movement of the landing gear control lever from the wheels down position when the aircraft weight is on the gear. When the weight of the aircraft is on the landing gear, this solenoid de-energizes, allowing the solenoid armature pin to protrude outboard. This solenoid armature pin position mechanically prevents movement of the landing gear control lever from the wheels down position. Depressing the solenoid armature pin allows control lever movement for emergency or test procedures. Landing Gear Warning System There are two types of visual warnings for the landing gear. One indicator is located in the pilot and copilot landing gear c ontrol levers and the other uses a flashing WHEELS indication on the pilot and copilot instrument panel. Additionally, on aircraft incorporating AFC 460, an aural indication is provided in the flight station (Figure 4-17). Landing Gear Control Lever Warning Indicators The pilot and copilot Landing Gear Control Lever warning indicators come on by movement of the landing gear control lever. Selecting either UP or DN on the landing gear control lever closes a set of contacts on the Landing Gear Control Lever control switch. A ground supplied by either the uplock or downlock s witches completes the circuit to the Indicator Lights Control, turning on the indicators.

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Figure 4-17 — Schematic for landing gear warning system

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Figure 4-18 — Schematic for landing gear warning system. Landing Gear Wheels Warning Indicators These flashing WHEELS indicators are controlled by the downlock switches and engine p ower lever switches. If the landing gear is up, or in transit and the engine power levers are set below cruise, a ground is applied to the Indicator Lights Control from the landing gear downlock switches, which start the indicators flashing. The WHEELS indicators may be cancelled by pressing the LANDING GEAR WARNING OVERRIDE pushbutton switch. Landing Gear WHEELS Warning Indicators and Horn On aircraft incorporating AFC 460, the flashing WHEELS warning indicators and Landing Gear Warning Horn are controlled by the downlock switches, Power Lever S witch Assembly, Power Lever Actuator and Switch Assembly, 142 Knot Airspeed Switch, 153 Knot Airspeed Switch, and Flap Handle Position Switch.

Landing Flap Position Warning On aircraft incorporating AFC 460, when the wing flap control lever is set to FLAPS DOWN, the flap handle position switch contacts are closed. A ground is supplied to so und the Landing Gear Warning Horn and start the WHEELS warning indicators flashing if the landing gear is not down and locked. This warning cannot be canceled using the LANDING GEAR WARNING OVERRIDE pushbutton switch. 4-31

Airspeed Switch and Power Lever Switch Warnings On aircraft incorporating AFC 460, two conditions can initiate the landing gear warnings to sound the Landing Gear Warning Horn and flash the WHEELS warning indicators. The 153 Knot Airspeed Switch is in series with the 48 degree microswitches in the Power Lever Actuator and Switch Assembly and downlock switches to eliminate nuisance warnings during normal flight evolutions above 153 knots. Landing gear warnings below 153 knots but above 142 knots can be canceled by use of the LANDING GEAR WARNING OVERRIDE pushbutton switch. A second 142 Knot Airspeed Switch is in series with the 60 degree microswitches in the Power Lever Switch Assembly, which provides a landing gear warning that cannot be cancelled if the engine power levers are at or below FLT START and airspeed is below 142 knots with the landing gear up or in transit. Landing Gear Position Indicators Three indicators on the copilot instrument panel display UP when gear is retracted, striped (barber pole) when gear is in transit or power is not applied; or wheel symbol when gear is down. LANDING GEAR POSITION LEFT, NOSE, and RIGHT indicators

are driven by grounds supplied by their respective uplock or downlock switches. Their 28 vdc power is supplied by the EXTENSION MAIN DC BUS circuit breaker panel LANDING GEAR POS IND circuit breaker. Arresting Gear System The arresting gear system stops the aircraft during carrier landings and emergency field arrestments. The primary component of the system is the arresting hook on the underside of the aft fuselage. The hook is pivoted at the forward end allowing up, down, and sideways motion. The hook engages a runway or cross deck pendant when in the down position. A lever in the crew station controls raising and lowering of the hook (Figure 4-19). The lever electrically energizes a hydraulically actuated vertical damper cylinder for hook retraction and trips an uplatch mechanism through mechanical linkage for lowering the hook. Two horizontal dampers and one vertical damper dampen hook motion from deck impact forces. Two centering spring assemblies maintain the hook in the center position.

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Figure 4-19 — Arresting gear control and indicating system.

E xtension The arresting hook is extended by moving the control handle in the crew station from the up to the down position. This removes tension from the cable and allows the uplatch mechanism to deflect to the opposite extreme of travel (aft). The arresting hook is then free to extend by action of the vertical damper cylinder and its own weight. At the same time, a switch in the control handle actuates, de-energizing both the arresting hook relay and the selector valve solenoid. This action permits fluid from the vertical damper cylinder to return. The surge damper prevents return line high pressure surges, caused by hook extension, from damaging other subsystems by reducing the surge.

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Retraction Moving the crew station lever from the down to the up position puts tension back onto the cable. This action moves the uplatch mechanism forward to receive the arresting hook and latch it in the retracted position. As the lever is brought up, a switch in the lever mechanism actuates, sending current through the de-energized arresting hook relay to energize the selector valve solenoid. Hydraulic pressure then goes to the vertical damper cylinder to raise the hook. When the hook reaches the retracted position, it actuates the up limit switch, which completes the circuit, energizing the arresting hook relay. This, in turn, breaks the circuit to the selector valve solenoid and stops the flow of hydraulic fluid to the vertical damper. The arresting hook relay has a 1.1 second time delay to assure the hook is up and locked before removal of hydraulic pressure. Nosewheel Steering System The nosewheel steering system (Figure 4-20) is an electrically controlled, hydraulically operated system. It provides a nonlinear relationship between the rudder pedals and the angular position of the nosewheel. The system provides the pilot with adequate directional control of the aircraft during ground operation. It consists of a hydraulic steer- damp unit, solenoid-operated shutoff valve, command potentiometer, steering feedback potentiometer, steering amplifier, and an electrical control system. Control of the electrical power to the system is through either of the ground safety switches on the left or right main landing gear. The landing gear handle switch, a push-button switch on the pilot’s stick grip, and the rudder pedals also control electrical power. When the steering system electrical circuits energize, the steering system aligns for steering operation. At this point, the solenoid-operated hydraulic shutoff valve opens to supply fluid to the steering actuator. Simultaneously, all related circuitry for controlling the steering servo valve activates. The electrical section of the steering system is essentially a bridge circuit. One side of the bridge circuit runs through the command potentiometer; the opposite side runs through the feedback potentiometer. Output of each potentiometer goes to the steering system amplifier. Amplifier output currents flow to the hydro-mechanical servo valve coils, and the net signal actuates the servo valve, causing nosewheel steering action. During operation, the nosewheel steering system attempts to maintain symmetry of the electrical bridge circuit. The circuit is symmetrical when both potentiometers supply equal voltages to the amplifier and the servo valve is at a null position. The bridge circuit becomes unbalanced when initiating a turn request by the repositioning of the command potentiometer wiper. Current differential in the servo valve windings reflects an unbalanced bridge circuit. Movement of the servo valve ports hydraulic pressure to the appropriate side of the actuator piston to cause the nosewheel to turn.

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Figure 4-20 — Nosewheel steering system.

S teer-Damper Unit The steer-damper unit is an electrically controlled, hydraulically operated package on the nose gear strut assembly. The unit provides both nosewheel steering and the required shimmy damping effect. The package consists of a check valve, servo valve, bypass valve, two unidirectional restrictors, the steering actuator, and fluid compensator. The check valve prevents reverse flow from the unit to the shutoff valve. The servo valve controls the actuator position by controlling fluid flow to and from the actuator in response to signal variations from the amplifier. The bypass valve closes off the 4-35

interconnecting passages between both ends of the actuator whenever hydraulic pressure is available to the unit. This permits the actuator to act as a steering unit instead of a damping unit. The unidirectional restrictors provide a restricted reverse flow to dampen nosewheel shimmy. The steering actuator is a balanced piston-type hydraulic actuator, which provides the force to turn the nosewheel. Also, with the restrictors, the steering actuator provides the shimmy damper action. The fluid compensator is in the return passage in the unit and traps a quantity of fluid at 40 to 100 PSI. The compensator supplies fluid to the actuator through the bypass valve and the restrictor when the unit is being used as a shimmy damper, and extra fluid is necessary to prevent cavitation of the actuator. Since the compensator traps fluid in the actuator, it includes thermal relief provisions to prevent excessive pressure buildup within the steer- damper unit. Shutoff Valve The steering shutoff valve is a three-way, two-position, normally closed, solenoid- operated valve. The valve controls hydraulic system pressure to the steer-damper unit. When the valve de-energizes, fluid flow is cut off from the steerdamper unit. When the valve energizes (during normal steering or arrested landings), pressure flows to the check valve, bypass valve, and servo valve in the steer-damper unit. Command Potentiometer The steering system has a rotary-type, pedal-position (command) potentiometer. This potentiometer mechanically links to, and is driven by, the rudder pedals. It provides a nonlinear steering response. The potentiometer sends a signal to the nose gear steering amplifier showing the degree and direction of turn commanded by the pilot. Moving the potentiometer, with the steering system operating, unbalances a bridge circuit causing the steering amplifier to signal the servo valve to turn the nosewheel. Feedback Potentiometer The steering feedback potentiometer assembly consists of a potentiometer that attaches to and is driven by the drive arm on the nose gear spindle. As the nose gear moves in response to the pilot’s command, the feedback potentiometer feeds back a signal to the steering amplifier. When the feedback potentiometer signal matches the command potentiometer signal, the amplifier output causes the servo valve to neutralize and stop movement of the nosewheel. The feedback potentiometer assembly contains a swivel disconnect switch, which opens when the nosewheel turns 750 to 800 either side of straight ahead. This action electrically de-energizes the circuit to prevent reverse s teering or damage to the aircraft. Amplifier The steering amplifier is a small transistorized differential amplifier. The amplifier detects the differential positions of the command potentiometer and the steering feedback potentiometer. Any difference in signals received results in a signal going to the servo valve to port hydraulic pressure to the steering actuator. This causes the nosewheel to turn and the steering feedback potentiometer to move. When the feedback potentiometer signal matches the command potentiometer signal, the nosewheel stops turning. In addition, the amplifier contains a circuit that provides a centering signal, which holds the nosewheel in the straight-ahead position during arrested landings.

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Operation When the nosewheel steering switch button is depressed, power goes to the electrical control system and the solenoid-operated shutoff valve. As the valve opens, hydraulic pressure flows to the servo valve on the steer-damper unit. Signals go to the amplifier from the command potentiometer and from the feedback potentiometer on the steering linkage. If these signals are equal, the amplifier signals the servo valve to stay in the neutral position. When the nosewheel (and consequently the feedback potentiometer) does not correspond to rudder pedal position (and command potentiometer), the signals going to the amplifier are different. This causes the amplifier to send a signal to the servo valve. The signal sent to the servo valve causes the valve to port pressure to the steering actuator in the steer-damper unit. The hydraulic pressure causes the actuator to move the nosewheel (and the feedback potentiometer) to a position corresponding to rudder pedal position (command potentiometer). When the nosewheel reaches a position corresponding to the pedal position, the signals to the amplifier are equal. The amplifier now signals the servo valve to a neutral position. When the servo valve goes to a neutral position (with the steering system energized and hydraulic pressure available), it blocks off both pressure and return passages to the steering actuator. Thus, the actuator is hydraulically locked in position. The steering actuator will remain locked in position until the servo valve receives a signal to turn the wheel or hydraulic and/or electrical power is removed from the system. When the steering system is not in use, the steer-damper unit performs the functions of a shimmy damper. The system accomplishes shimmy damping by trapping hydraulic fluid on both sides of the steering actuator piston and forcing this fluid from one side of the actuator to the other side through the restrictor. Catapulting System The catapulting system (Figure 4-21) provides catapult handling and attachment capabilities for carrier operations. The system consists of a catapult launch bar, a launch bar actuating cylinder and gimbal, swivel joints, a crew station controlled selector valve, leaf centering spring, leaf retracting springs, and a catapult tension bar socket. The launch bar is swivel mounted on the nose gear outer cylinder and can extend and retract during taxi operations. The launch bar automatically retracts after catapulting. A launch bar warning light comes on during any of the following conditions:  The launch bar control switch is in EXTEND.  The selector valve is in bar extend position (solenoid A energized).  The launch bar is not up and locked with weight off the landing gear.  The launch bar control switch is in RETRACT and the launch bar actuator is not up and locked. Accessories for the catapulting system include a tension bar and a catapult holdback bar. The catapult tension bar socket mounts on the nose gear axle beam and provides for attachment of the tension bar for tensioning the aircraft before catapulting. 4-37

Figure 4-21 — Catapulting system control and indicating system. Placing the launch bar control switch in EXTEND completes a power circuit through the weight-on-gear switch, applying 28-volt dc to the launch bar selector valve extend solenoid (solenoid A). The launch bar valve position switch also completes a circuit through its closed contacts to apply 28-volt dc to the launch bar warning light. The warning light comes on and remains on as long as the switch is in EXTEND. When the selector valve actuates, hydraulic pressure extends the launch bar actuator. A plunger on the end of the valve mechanically actuates the launch bar valve position switch. Switch actuation completes a parallel circuit to the warning light. When the control switch is in the OFF position, with weight on the landing gear, the warning light should go off. If the control switch is in RETRACT, the warning light should come on until the launch bar is up and locked. If the light remains on, the selector valve did not cycle from the bar extend position, and pressure is still on the launch bar actuator extend side. The hydraulic pressure to the launch bar actuator unlocks locking fingers inside the actuator and extends the actuator to lower the launch bar. As the locking fingers unlock, they close the contacts of the launch bar up-lock switch inside the actuator. After catapulting, the weight-on-gear switch moves to the weight-off position. This applies 28- volt dc to the launch bar selector valve retract solenoid (solenoid B) and to the launch bar warning light through the energized contacts of relay K1. K1 remains on until the launch bar is up and locked. Power to the launch bar selector valve retract solenoid provides automatic hydraulic retraction and locking of the launch bar after catapulting. The launch bar actuating cylinder, locking in the fully retracted position, de-energizes K1, turning off the launch bar warning light. The approach light circuit goes through de- energized relay K1, giving an additional indication that the launch bar is up and locked. The launch bar control switch goes to RETRACT and is held thereto retract the launch bar hydraulically. This completes a power circuit to apply 28-volt dc to the launch bar selector valve retract solenoid (solenoid B). The energized selector valve directs hydraulic pressure to retract the launch bar. The control switch returns to OFF when released, de-energizing the selector valve. 4-38

Figure 4-22 — Speed brake system. NOTE Figure 4-22 shows the speed brake control switch being held in the OUT position. This allows hydraulic pressure to port to the extend side of the speed brake actuators, forcing the speed brakes to the extend position. Speed Brake System Speed brakes are moveable control surfaces used for reducing the speed of aircraft. Some manufacturers refer to them as dive brakes, others call them dive flaps. On some aircraft, they are on the sides or bottom of the fuselage; on others they attach to the wings. Regardless of their location, speed brakes serve the same purpose on all aircraft, they keep the speed from building up too high in dives. They can also slow down the speed of aircraft preparing to land. Speed brakes have electrical control and hydraulic operation. Figure 4-22 shows a typical speed brake system, and you should refer to it while reading this section.

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Extension Extension of the speed brakes is done by placing the speed brake control switch (located on the throttle lever grip) to the OUT position. The OUT position is a momentary contact position, the switch being spring loaded to return to STOP when released. Holding the control switch in the OUT position energizes both selector valve solenoids (solenoids A and B). This action connects hydraulic pressure to the actuator extend side and connects the return to the actuator retract side. Any desired brake position may be attained and will be held by a hydraulic lock within the selector valve. The speed brake out warning light illuminates when either the left or right speed brake is not fully retracted. The warning light circuit completion is through either the left or right speed brake retract position switch. Retraction In normal operation, the speed brakes retract by moving the speed brake control switch to the IN position. This de-energizes solenoid A by removing the power to the solenoid and energizes the speed brake relay. The action of the speed brake relay removes the power from solenoid B. With both selector valve solenoids de-energized, the selector valve permits pressure flow to the retract side of the speed brake. It also allows return flow from the extend side to the hydraulic return. The speed brake warning light goes out when both speed brakes are in the fully retracted position. Emergency Retraction To accomplish emergency speed brake retraction, place the emergency speed brake switch in the EMER RETRACT position. This switch de-energizes solenoids A and B of the selector valve (normal solenoid positions for retracting the speed brakes), which connects both the extend and retract sides of the speed brake actuators to the system return. With the removal of hydraulic pressure, the speed brakes are shut by the airstream. This action is necessary only if the speed brake relay does not energize when the speed brake control switch is in the IN position. If an electrical failure (such as a popped circuit breaker) occurs with the speed brakes extended, retraction is the same as actuation of the emergency retract switch. Canopy System Solenoid valves have many applications in naval aircraft. The canopy hydraulic system shown in Figure 4-23 is an electrohydraulic system using a solenoid selector valve. The canopy selector valve is a four-way, two-position spool valve, which actuates either electrically or manually. The canopy system consists of a sliding canopy mounted over the crew station area and the components required for normal operation and emergency jettison of the canopy. The entire system is hydraulically operated except for the electrical canopy jettison device. Hydraulic power for operation of the canopy is from either the combined hydraulic system or the hand pump system.

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When a canopy switch is in the CLOSE position, it completes a circuit from the 28-volt dc bus to a terminal on the isolation switch. Current flows from the terminal, through the closed contacts of the canopy switch, to solenoid 2 of the canopy selector valve. The selector valve then energizes to the close position. Now, hydraulic pressure from either the combined hydraulic system or the hand pump system flows through the selector valve into the canopy close line. Pressure through a flow regulator to the rod end of the canopy actuating cylinder causes the piston and rod to retract closing the canopy. When the canopy switch is in the OPEN position, power flows from the 28-volt dc bus through the control circuit breaker, the isolation switch terminal, the opposite contacts of the canopy switch, to solenoid 1 of the canopy selector valve. The selector valve energizes to the open position, reversing the sequence of pressure and return flow. Hydraulic pressure flows through the canopy open line to the canopy seal valve and hydraulically trips the seal valve. Thus, the air pressure in the canopy seal is dumped, allowing the seal to deflate. Pressure in the canopy open line continues its flow through a flow regulator to the back head end of the cylinder. This action extends the cylinder, opening the canopy. Hydraulic fluid in the opposite end of the cylinder returns through the canopy close line to the selector valve, across the valve, and into the combined hydraulic system. Fluid then returns to the main reservoir of the system. Figure 4-23 — Canopy system. 4-41

Figure 4-24 — Canopy jettison. For canopy jettison (Figure 4-24), the emergency canopy jettison switch, or either of the two emergency outside jettison switches initiates a firing circuit. Closing any of the three switches completes the circuit to the electrically fired canopy jettison cartridge on the canopy cylinder back head end. The cartridge discharges through the canopy cylinder, causing the canopy to jettison.

Pneumatic Power System The pneumatic power system supplies compressed air for various normal and emergency pneumatically operated systems. The compressed air is held in storage cylinders in the actuating systems until required by actuation of the system. These cylinders and power system manifold receive an initial charge of compressed air or nitrogen from an external source. In flight, the air compressor replaces the pressure and volume lost through leakage, thermal contraction, and system operation. The air compressor receives supercharged air from the engine bleed air system. This ensures an adequate air supply to the compressor at all altitudes. The air compressor is driven by an electric or a hydraulic motor. The system under discussion in this chapter is hydraulically driven. (Figure 4-25)

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Figure 4-25 — Pneumatic system. T he aircraft hydraulic system provides power to operate the hydraulic-motor-driven air compressor. The air compressor hydraulic actuating system consists of a solenoid- operated selector valve, flow regulator, hydraulic motor, and motor bypass line check valve. When energized, the selector valve allows the system to pressurize and run the hydraulic motor. When de-energized, the valve blocks off hydraulic pressure, stopping the motor. The flow regulator compensates for the varying hydraulic system flow and pressures by metering the fluid flow to the hydraulic motor. Thus, excessive speed variation and/or overspending of the compressor is prevented. A check valve in the motor bypass line prevents system return line pressure from entering the motor and stalling it. 4-43

The air compressor is the pneumatic system’s pressurizing air source. The compressor activation or deactivation is by the manifold pressure sensing switch, an integral part of the moisture separator assembly. The moisture separator assembly is the pneumatic system’s pressure sensor-regulator and relief valve. The manifold pressure switch governs air compressor operation. When manifold pressure drops below 2,750 PSI, the pressure sensing switch closes, energizing the separator’s moisture dump valve and hydraulic selector valve, activating the air compressor. When manifold pressure reaches 3,150 PSI, the pressure sensing switch opens, de-energizing the hydraulic selector valve, deactivating the air compressor and dump valve. The deactivated dump valve vents overboard any moisture in the separator. The separator includes a thermostat and heating element. The thermostatically controlled wraparound blanket heating element prevents moisture from freezing within the reservoir in low-temperature atmospheric conditions. The safety fitting at the moisture separator inlet port protects the separator from internal explosions due to hot carbon particles or flames that the air compressor may emit. A chemical drier further reduces the moisture content of the air emerging from the moisture separator. An air charge valve provides the entire pneumatic system with a single external ground servicing point. An air pressure gauge, near the air charge valve, is for servicing the pneumatic system. This gauge shows manifold pressure. The ground air charge line air filter prevents entry of particle impurities into the system from the ground servicing power source. AIRCRAFT ENVIRONMENTAL SYSTEMS The proper operation of today’s aircraft requires maintaining a proper environment not only for personnel but also for equipment on board. Similarities exist in the electronic equipment controlling these systems and the components within these systems. The environmental systems on most aircraft include crew station/cabin air conditioning and pressurization, equipment cooling, windshield anti-icing and defogging, and equipment pressurization systems. Some aircraft accomplish equipment cooling by routing additional ducting from the crew station/cabin system; other aircraft use a separate cooling system. These systems are not the exclusive responsibility of the AE, but rather are the responsibility of other ratings with the AE assisting. Terms and Definitions The Air Conditioning System conditions ambient air to provide crew environmental control and electronic equipment cooling both in flight and on the ground. You must become familiar with some terms and definitions to understand the operating principles of air conditioning systems. The following terms are self-explanatory: engine heat, solar heat, electrical heat, and body heat of personnel. These sources of heat make cabin air conditioning necessary, as does ram air temperature. Ram air temperature is the frictional temperature increase created by ram compression on the skin surface of an aircraft. This factor becomes serious only at extreme airspeeds. For example, for an aircraft flying at 45,000 feet, at 1,200 MPH, the ram air temperature would be about 2,000°F on some parts of the aircraft. This extreme temperature, plus the heat from other sources, would cause cabin temperature to rise to about 190°F. The maximum temperature that a crew member can endure and still maintain top physical and mental efficiency is about 80°F. Prolonged 4-44

NOTE Some of the terms were discussed in chapter 1, “Basic Physics.” Although they are briefly defined in this chapter, a review of chapter 1 will be helpful. exposure to a temperature greater than 80°F will seriously impair mental and physical abilities. Furthermore, under low-speed operating conditions at low temperature, cabin heating may be necessary.

Definitions for some of the terms relative to temperature control are as follows: Absolute temperature – Temperature measured along a scale that has zero value at that point where there is no molecular motion (–273.1°C or –459.6°F). Adiabatic – A word meaning no transfer of heat. The adiabatic process is one in which no heat transfers between the working substance and any outside source. Ambient temperature – The temperature measured in the area immediately surrounding the object under discussion. Ram air temperature rise – The increase in temperature created by the ram compression on the surface of an object traveling at high speed through the atmosphere. The rate of increase is proportional to the square of the speed of the object. Temperature scales.  Celsius (C) – a scale on which 0° represents the freezing point of water, and 100° is the boiling point of water at sea level.  Fahrenheit (F) – a scale on which 32° represents the freezing point of water, and 212° is the boiling point of water at sea level. Cabin System The primary function of the cabin air conditioning and pressurization system is to maintain crew station temperature and pressure within parameters for crew safety and comfort. To do this, the system forces a mixture of dehumidified refrigerated air and hot engine bleed air through crew station louvers. The temperature of the mixture is automatically maintained through a continuously selective range by a temperature control system, consisting of temperature sensors with associated flow control valves and an electronic controller. A pressure regulator and safety valve control cabin pressure. The manual dump control is available if the safety valve malfunctions. A block diagram for the airflow of a cabin air conditioning and pressurization system is shown in Figure 4-26. 4-45

Figure 4-26 — Cabin air-conditioning and pressurization diagram.

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Figure 4-27 — Environmental Control System (ECS) control panel. NOTE Many ACACS components are part of the ECS Pack Assembly which is removed/installed as unit. System Controls The Environmental Control System (ECS) control panel is shown in Figure 4-27. Switches and knobs on the ECS control panel assembly control the air- conditioning operation. The Air Cycle Air Conditioning System (ACACS) cools and conditions engine bleed air for use in the ECS and other systems from the crew station. The air-conditioning system control consists solely of the ECS Temperature/Flow Controller. From here out, it will be referred to as the ECS Temp/Flow Controller. The ECS Temp/Flow Controller is a microprocessor-based, fully integrated digital control unit with an upload capability that allows upgrades when system changes are necessary. The ECS Temp/Flow Controller provides automatic electrical control and monitoring of many systems in the ECS group and it provides fault detection and fault isolation of system components. It uses continuous and initiated Built-In-Test (BIT) to monitor system degradation status and communicates with the aircraft’s Mission Computer (MC) to produce Maintenance Status Panel (MSP) codes as required. To understand the operation, the aircraft is in flight and bleed air is being sent through the ECS system.

The ECS panel assembly in the crew station contains the switchology to operate the aircraft air-conditioning system. The MODE switch provides system mode selection data to the ECS Temp/Flow Controller. In response, the ECS Temp/Flow Controller will provide command signals to several components in the avionics cooling and defog and air cycle air-conditioning systems. The MODE switch is a three-position toggle switch with the positions AUTO, MAN, or OFF/RAM. When set to the forward AUTO position, the ECS Temp/Flow Controller directs the System Flow Modulating Pressure Regulator to satisfy the total airflow requirements of the cabin or crew station and the avionics cooling system as dictated by flight conditions and pilot input. When the MODE switch is set to the middle position, MAN, the ECS Temp/Flow Controller removes the position command to the System Flow Modulating Pressure Regulator; thus, the valve regulates maximum airflow to the crew station and provides sufficient airflow for crew comfort and 4-47

avionics cooling. When the ECS MODE switch is set to aft position, OFF/RAM, the ECS Temp/Flow Controller shuts down the Air Cycle Air Conditioning System (ACACS). The System Flow Modulating Pressure Regulator is driven fully closed. This stops all airflow to the ACS Turbine/Compressor. At the same time, the ECS-off status is reported to the MC. The Servo Air Moisture Separation mode is warm, dry air provided by the Warm Air system for use as muscle air for the Avionics Flow Control Valve and the Cabin Flow Control Valve. The cabin TEMP control knob, located on the ECS panel, provides aircrew temperature selection data to the ECS Temp/Flow Controller. In response, the controller provides command signals to the Cabin Flow Control Valve and the Cabin Add-Heat Valve which are part of the cabin cooling and defog system and provides signals to determine the proportional amounts of engine bleed air and refrigerated air mixture from the LCS No. 3 Heat Exchanger. The cabin TEMP control knob is a rotary type knob and the selections are COLD to HOT. Automatic (AUTO) Mode Operation As you read this section, refer to Figure 4-28. With engine bleed air supplied to the aircraft and the Secondary Bleed Air Pressure Regulating Shut-off Valve opened, engine-supplied bleed air is ready to be distributed to the components of the air- conditioning system. Placing the crew station ECS MODE switch to AUTO (Figure 4-28, frames 1, 2 and 3), the ECS Temp/Flow Controller receives ECS MODE switch position signals. Simultaneously, the Power Distribution Panel (PDP) No. 6, 28 VDC is routed through the CBN RAM AIR VLV circuit breaker to the normally open contact of the ECS control panel. ECS processor signals— internal to the ECS Temp/Flow Controller—are sent from the controller through the ECS MODE switch closed contacts and back to the ECS Temp/Flow Controller (Figure 4-28, frame 4). ECS command signals are sent to the Compressor Inlet Pressure Sensor, Condenser Icing Pressure Sensor, Condenser Delta Pressure Sensor, Compressor Protective Temperature Sensor, Condenser Outlet Temperature Sensor, and the Secondary Heat Exchanger Bypass Valve (Figure 4-28, frames 5 through 12). Additionally, an ECS Temp/Flow Controller output signal is sent to the System Flow Modulating Pressure Regulator closed contacts. The ECS Temp/Flow Controller receives and processes these input signals from the ECS control panel switchology, ECS sensors, and control valves (Figure 4-28, frames 10 and 11). Hot bleed air, precooled by the Primary Heat Exchanger (Figure 4-28, frame 13), enters the compressor end of the Turbine/Compressor Assembly mounted to the top of the ECS Pack. The Compressor Inlet Pressure sensor monitors the inlet air of the Turbine/Compressor temperature and senses if temperature reached is the set value as determined by the ECS controller. If predetermined values are too high, the ECS controller will send a signal to the MC displays as appropriate. When bleed air from the System Flow Modulating Pressure Regulator enters the Turbine/Compressor Assembly, it is compressed to approximately twice its inlet pressure. A portion of the air leaving the compressor is directed downstream to the Compressor Protective Temperature Sensor. Temperature sensor output signals are sent back to the ECS Temp/Flow Controller for processing. Cooled compressed bleed air is also sent to the Secondary Heat Exchanger Bypass Valve and around the Secondary Heat Exchanger, Liquid Cooling System (LCS) No. 1 Heat Exchanger, and the Intermediate Water Extractor. This bypass air is sent directly to the inlet air side of the Reheater Heat Exchanger. The Secondary Heat Exchanger Bypass Valve modulates the amount of bypass air to maintain a set temperature. This also prevents icing of the condenser when the ACACS is producing more air than is required by the system. Air not 4-48

bypassed, flows through the Secondary Heat Exchanger and is cooled by ram air before passing to the LCS No. 1 Heat Exchanger. Under most flight conditions, air in the LCS No. 1 Heat Exchanger is further cooled by passing heat to the LCS fluid. This cooling allows the delivery temperature of the ECS pack to be further reduced (Figure 4-28, frames 14, 15, and 16). From the LCS No. 1 Heat Exchanger, air flows through the Intermediate Water Extractor. Water condensed in the Secondary Heat Exchanger and the LCS No. 1 Heat Exchanger is removed by the Intermediate Water Extractor and routed to the Secondary Heat Exchanger Water Spray Bar. Cooling in the Secondary Heat Exchangers is augmented with a water mist from the Water Spray Nozzles and Water Spray Bar. Both Intermediate Water Extractor outlet air and bypass air enter the Reheater Heat Exchanger and the Condenser/LCS No. 3 Heat Exchanger. Here, the air is further cooled. The Reheater Heat Exchanger uses Water Extractor outlet air to absorb heat while the Condenser/LCS No. 3 Heat Exchanger uses cold turbine discharge air to absorb the heat. Water condensed in the Heat Exchangers is removed by the Water Extractor. Following water extraction, the cool dry air is warmed (reheated) in the Reheater Heat Exchanger to improve turbine operation. The dry conditioned air flows to the turbine end of the Turbine/Compressor Assembly where it is converted to mechanical energy driving the turbine and reducing the turbine discharge air to subfreezing temperatures. The turbine and the compressor are protected from overheat damage by protective temperature sensors. Electrical signals sent to the Condenser Icing Pressure sensor monitor temperature to the LCS No.3 Heat Exchanger. If too cold, signals are sent from the ECS Temp/Flow Controller to the Turbine Anti-ice Add Heat valve (Figure 4-28, frames 17 and 18) energizing the valve coil allowing heated airflow. Cold, dry air leaving the turbine is mixed with warm air from the Turbine Anti-ice Add Heat Valve to control temperature and prevent condenser icing. The mixed air passes through the Condenser/No. 3 LCS Heat Exchanger where it provides final temperature control of the LCS cooling fluid and airflow. After leaving the Condenser/No. 3 LCS Heat Exchanger, the air is transported to the crew station Cabin Flow Control Valve and the Avionics Flow Control Valve for avionics cooling. The Cabin Flow Control Valve directs air to the crew station and the Avionics Flow Control Valve directs cooled air to the avionics bays. To increase heat exchanger cooling capacity during low speed flight or ground operations, Auxiliary Air Inlet doors and Bleed Air Ejectors increase ram airflow across the Primary and Secondary Heat Exchangers. Airflow to the ejectors is applied or removed through an ejector valve controlled by the ECS Temp/Flow Controller but are forced closed on the ground when the BLEED AIR switch is OFF, independent of the ECS Temp/Flow Controller. Manual (MAN) Mode Operation As you read this section, refer to Figure 4-28, frames 19, and 20. When MAN mode is selected, the ECS Temp/Flow Controller receives an ECS MODE switch position signals from the ECS control panel. In return, the ECS Temp/Flow Controller processes and removes the command signal to the Turbine Anti-ice Add Heat Valve. The valve de- energizes and warm air is completely isolated from the LCS No. 3 Heat Exchanger. Simultaneously, the ECS Temp/Flow Controller sends signal s to the de-energized contacts of the ECS OFF/RAM relay, turning the System Flow Modulating Pressure Regulator motor. When the motor turns, it controls the spring-loaded valve, thus controlling the amount of air to the turbine. The ECS Temp/Flow Controller removes the position signal to the System Flow Modulating Pressure Regulator’s normally closed 4-49

contacts as the contacts are mechanically moved by the turning of the motor. ECS processor signals—internal to the ECS Temp/Flow Controller—are sent from the controller to the various sensors controlling the inputs to the controller. ECS command signals are sent to the Compressor Inlet Pressure Sensor, Condenser Icing Pressure Sensor, Condenser Delta Pressure Sensor, Compressor Protective Temperature Sensor, Condenser Outlet Temperature Sensor, and the Secondary Heat Exchanger Bypass Valve. These signals maintain the operation in the manual operation as they performed in the automatic mode. Once the ECS Temp/Flow Controller processes these sensor signals, manual operation occurs. Hot bleed air from the Bleed Air system enters the Turbine/Compressor Assembly; the compressor sends the precooled air back downstream to the Compressor Protective Temperature Sensor where a temperature sensor output signal is sent back to the ECS Temp/Flow Controller. Cooled compressed bleed air is also sent to the Secondary Heat Exchanger Bypass Valve around the Secondary Heat Exchanger, LCS No. 1 Heat Exchanger, and the Intermediate Water Extractor. This bypass air is sent directly to the Reheater Heat Exchanger inlet. The Secondary Heat Exchanger Bypass Valve modulates the amount of bypass air to maintain a set temperature at the Reheater side of the cold air inlet. This also prevents icing of the condenser when the ACACS is producing more air than is required by the system. Air not bypassed, flows through the Secondary Heat Exchanger and is cooled by ram air before passing to the LCS No. 1 Heat Exchanger. Under most flight conditions, air in the LCS No. 1 is further cooled by passing heat to the LCS fluid. This cooling allows the delivery temperature of the ECS pack to be further reduced. From the LCS No. 1 Heat Exchanger, air flows through the Intermediate Water Extractor. Water condensed in the Secondary Heat Exchanger and the LCS No. 1 Heat Exchanger is removed by the Intermediate Water Extractor and routed to the Secondary Heat Exchanger Water Spray Bar. Cooling in both the Primary and Secondary Heat Exchangers is augmented with a water mist from the Water Spray Nozzles and Water Spray Bar. Both Intermediate Water Extractor outlet air and bypass air enter the Reheater Heat Exchanger and the Condenser/LCS No. 3 Heat Exchanger. Here, the air is further cooled. The Reheater Heat Exchanger uses Water Extractor outlet air to absorb heat while the Condenser/LCS No. 3 Heat Exchanger uses cold turbine discharge air to absorb the heat. Water condensed in the Heat Exchangers is removed by the Water Extractor. Following water extraction, the cool dry air is warmed (reheated) in the Reheater Heat Exchanger. The dry conditioned air flows to the turbine end of the Turbine/Compressor Assembly where it is converted to mechanical energy driving the turbine and reducing the turbine discharge air to subfreezing temperatures. The turbine and the compressor are protected from overheat damage by protective temperature sensors. Electrical signals sent to the Condenser Icing Pressure Sensor monitor temperature to the LCS No.3 Heat Exchanger. Cold, dry air leaving the turbine passes through the Condenser/No. 3 LCS Heat Exchanger where it provides final temperature control of the LCS cooling fluid. After leaving the Condenser/No. 3 LCS Heat Exchanger, the air is transported to the crew station Cabin Flow Control Valve and the Avionics Flow Control Valve for avionics cooling and crew comfort. To increase heat exchanger cooling capacity during low speed flight or ground operations, Auxiliary Air Inlet Doors and BLEED Air Ejectors increase ram airflow across the Primary and Secondary Heat Exchangers. Airflow to the ejectors is applied or removed through an ejector valve controlled by the ECS Temp/Flow Controller but are forced closed on the ground when the BLEED AIR switch is OFF, independent of the ECS Temp/Flow Controller. 4-50

OFF/RAM Mode Operation This mode is also referred to as the Emergency mode. When OFF/RAM mode is selected (Figure 4-28, frame 22), the ECS Temp/Flow Controller receives ECS MODE switch position signal from the ECS Control Panel. In return, the ECS Temp/Flow Controller processes and sends signals through the ECS OFF/RAM relay’s energized coil to the System Flow Modulating Pressure Regulator Motor and the regulator is driven to the full-closed position. This shuts off bleed air to the Turbine/Compressor Assembly from the Secondary Pressure Regulating Shutoff Valve. Airflow to the crew station is now provided by ram air entering the system and the airflow for the avionics bays for avionics cooling is provided by the Emergency Aft Avionics Cooling Fan and ram air provided by the Auxiliary Air Inlet Doors. Throttle position Rotary Variable Differential Transformers (RVDTs) and aircraft Mach number determines when the Auxiliary Air Doors are opened and closed. The ECS Temp/Flow Controller sends ECS OFF status signal to MC for processing. When the system is shut down, the ECS Temp/Flow Controller commands the Ejector Shutoff Valve open. The Primary and Secondary Heat Exchanger Ejectors pull extra ram air across Primary and Secondary Heat Exchangers to allow more cooling. Built-in-Test (BIT), provided by the ECS controller evaluates performance of Air Cycle Air Conditioning System components and provides any related maintenance codes to display in the Maintenance Status Panel.

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Figure 4-28 — Air Cycle Air-Conditioning System schematic diagram.

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Equipment Cooling System Ram air is the primary means of ventilation for the forward and aft equipment compartments. The Ram Air Thermal Switch, Equipment Cooling Valve, Forward Compartment Ram Air Valve, and the Aft Compartment Ram Air Valve control ventilation of these compartments. The equipment cooling system flow diagram is shown in Figure 4-29. When the aircraft is in flight, the equipment cooling valve closes and the forward and aft compartment ram air valves open. Under these conditions, the right forward and aft equipment compartments are ram air ventilated. When temperature in the right wing ram air duct reaches 46.1°C (115°F), the equipment cooling valve opens, and the forward and aft compartment ram air valves close. This permits moist, cooled bleed air into the right forward and aft equipment compartments, ensuring adequate cooling when ambient air temperatures are excessive.

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Figure 4-29 — Equipment cooling flow diagram.

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With an engine running, the equipment cooling system (Figure 4-30) automatically engages when the AIR COND MASTER switch is at NORM. Under these conditions, the circuit from the 28-volt DC Essential Bus to the solenoid of the Engine Bleed Air Shutoff Valve is complete. Bleed air now flows to the Air Cycle Refrigeration Unit. The discharge from the cooling turbine supplies the cooled bleed air used by the Equipment Cooling System (ECS).

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Figure 4-30 — Equipment cooling schematic diagram.

W hen the aircraft is airborne and the ram air temperature is below 46.1°C (115°F), voltage from the 120-volt AC Primary Bus runs through the Overpressurization Relay and the unoperated Ram Air Thermal Switch to the Equipment Cooling Valve, the Forward Compartment Ram Air Valve, and the Aft Compartment Ram Air Valve. The Equipment Cooling Valve runs to the fully closed position, and the Forward Compartment and Aft Compartment Ram Air Valves run to the open position. Ram air flows to the right forward and aft equipment compartments and the cooled bleed airflow is shut off. When the ram air temperature exceeds 46.1°C, the Ram Air Temperature Thermal Switch operates. Voltage from the 120-volt AC Primary Bus goes through the Overpressurization Relay, and the Operated Ram Air Thermal Switch to the Equipment 4-56

Cooling Valve, the Forward Compartment Ram Air Valve, and the Aft Compartment Ram Air Valve. The Equipment Cooling Valve runs to the open position, and the Forward Compartment and Aft Compartment Ram Air Valves run to closed position. Undried, cooled bleed air flows to the forward and aft equipment compartments. The ram air to these compartments is shut off. When an overpressurization condition occurs, the pressure switch operates. This closes the Overpressurization Relay. Voltage from the 120-volt AC Primary Bus runs through the Overpressurization Relay to the Equipment Cooling Valve, the Forward Compartment Ram Air Valve, and the Aft Compartment Ram Air Valve. The Equipment Cooling Valve goes to the open position and the Forward and Aft Compartment Ram Air Valves run to the closed position. Under these conditions, the overpressure dumps into the forward and aft equipment compartments. The Computer Pressure Regulator maintains a pressure of 2 to 3 PSI in the equipment cooling line. The pressure-regulated, partially-dried cooled bleed air from the regulator automatically mixes with controlled quantities of hot bleed air. The servo-controlled Computer Temperature Control Valve controls the mixture process. This valve modulates in response to temperature signals from the computer temperature sensor in the computer inlet duct to maintain a computer duct temperature of 4.4° ± 2.8°C (40° ± 5°F). The temperature controlled air flows through the Ballistics Computer Set. Sonic venturis in the Ballistics Computer Outlet Duct and the Transmitter Modulator Inlet Duct limit the flow of cooling air. A Computer Cooling Shutoff Valve upstream of the Computer Temperature Control Valve provides a safety override for the Computer Temperature Control Valve. If the computer duct temperature exceeds 65.6°C (150°F), the computer thermal switch opens. This de-energizes the computer overheat indicator relay and completes the circuit to the COMPUTER OVERHEAT caution light. Voltage from the 28-volt essential DC bus runs to the Computer Overheat Indicator Relay. This energizes the Computer Cooling Control Relay, thereby closing the Computer Cooling Shutoff Valve. If the computer duct temperature drops below 65.6°C, the contacts of the Computer Thermal Switch close, energizing the Computer Overheat Indicator Relay. This action interrupts the circuit to the COMPUTER OVERHEAT caution light. The Computer Cooling Control Relay remains energized through the Computer Emergency Cooling Switch and its own contacts. Placing the CMPTR EMER COOL switch momentarily to RESET de-energizes the Computer Cooling Control Relay, completing the circuit to open the Computer Cooling Shutoff Valve. If the COMPUTER OVERHEAT caution light should cycle on and off indicating a constant overheat condition, place the CMPTR EMER COOL switch to ON, permanently closing the Computer Cooling Shutoff Valve. This allows uncontrolled, cooled bleed air to duct to the Ballistics Computer. Air exhausted from the Ballistics Computer Set and from the crew station circulates through the less critical electronic equipment compartments. This air supplies direct cooling requirements for Communication, Navigation, and Identification (CNI) equipment in the aft equipment compartment. Cabin and Vent Suit Temperature Control System The Cabin and Vent Suit Temperature Controller is a transistorized electronic device. It operates on 120 volts, 400 hertz. Maximum power consumption is 28 watts. Electrically, the controller consists of two channels— the cabin temperature channel and the ventilated suit channel. Both channels operate in basically the same manner. Bridge circuits compare temperature selector and temperature sensor resistances, which 4-57

Figure 4-31 — Cabin temperature channel. represent selected and actual temperature conditions and generate a resultant DC error voltage. The DC error voltage is then modulated and amplified. The resulting AC signal is the controller output, and it has phase and voltage characteristics proportional to the magnitude of the DC error signal. It goes to the two-phase servomotor of the appropriate control valve, where it modulates the valve to maintain the selected temperature. The Cabin Temperature Channel (Figure 4-31) of the temperature controller uses three bridge circuits to maintain cabin temperature. Selection of the desired temperature with the Cabin Temperature Selector varies resistance of one leg of the cabin temperature bridge. The varying cabin temperature changes the resistance of the cabin sensor in the second leg of the bridge circuit. Thus, a selected temperature must have a change of cabin temperature to provide bridge balance between the cabin sensor and selector. Whenever bridge imbalance exists, the resulting DC voltage is AC modulated by the modulator circuit, amplified by the Cabin Amplifier, and fed to the Control Valve Servomotor to either increase or decrease cabin temperature.

The function of the Cabin Duct Limit Bridge is to limit the temperature of the cabin inlet air. A diode-biasing network permits the Cabin Duct Limit Bridge to override the Cabin Temperature Bridge upon reaching the duct temperature limit. The resulting DC voltage 4-58

Figure 4-32 — Vent suit channel. goes to the Modulator Circuit, the Cabin Amplifier, and to the Control Valve Servomotor to decrease cabin temperature. The Cabin Duct Anticipator Bridge functions with sudden changes of air temperature in the cabin air inlet duct. To do this, the error voltage is capacitor coupled to the modulator. When the cabin temperature is being held at the selected temperature with constant cabin inlet air temperatures, the Anticipator Bridge is in balance. Should duct temperature suddenly change, with all other conditions remaining the same, the Anticipator Bridge is unbalanced. This unbalance causes an error voltage to go to the modulator circuit. The resulting amplified signal regulates the Cabin Control Valve to return the duct air to its original temperature. Error voltages caused by Cabin Temperature Bridge or Cabin Duct Limit Bridge imbalance override the Cabin Duct Anticipator Bridge, provided the duct air temperature error is gradual or small. One additional error voltage is capacitor coupled to the modulator circuit. The voltage change is proportional to the rate of change of the feedback potentiometer of the Cabin Dual Temperature Control Valve. This voltage change is applicable only when the valve is being regulated (the feedback potentiometer is rotating). Therefore, the error feedback voltage seen at the control input is proportional to the feedback potentiometer rate of change. Once regulation of the valve starts, the potentiometer rate-of-change voltage reduces initial starting voltage to the control valve actuator motor, slowing valve actuator rotation. The Vent Suit Channel (Figure 4-32) of the Temperature Controller uses one bridge circuit to maintain suit temperature. Operation of this bridge is similar to that of the Cabin Temperature Bridge. The Suit Temperature Selector resistance in one leg balances against the suit duct temperature sensor resistance in the opposite leg. Imbalance between the legs of the Suit Temperature Bridge results in a DC error voltage. This voltage is AC modulated by the modulator circuit, amplified, and goes to the Suit Temperature Control Valve, altering suit air temperatures.

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Rate-of-change voltage from the feedback potentiometer in the ventilated suit temperature control valve is capacitance-coupled to the modulator circuit. As in the cabin temperature channel, this voltage is present only when the valve is operating. It reduces the initial starting voltage to the valve once actuator rotation has started, thus slowing actuator rotation. Anti-Icing and Deicing Equipment The anti-ice and defrost system on some aircraft having the air-conditioning and pressurization system described in this chapter use the air cycle systems as an air source. The anti-ice and defrost equipment consists of the Windshield Anti-ice System (Figure 4-33) and the Windshield Defrost System. Each receives its hot air supply from the same manifold. The Anti-icing Switch is on the pilot’s Temperature Control Panel. In a typical system, a windshield overheat thermostat and a shutoff valve in the windshield defrosting duct operate together to prevent windshield overheating. The thermostat opens the valve automatically when the temperature becomes too high. This action diverts the hot defrosting air to the air-conditioning outlet at the floor of the crew station until the temperature drops. Windshield overheating occurs only if the cabin air temperature high limit pickup fails. Windshield Anti-Icing and Defogging System An electrical anti-icing and defogging system is now in the windshield panels of current naval aircraft. The panels are constructed of two pieces of semi-tempered plate glass laminated with a vinyl plastic core. The core acts as a safety device to prevent shattering in a collision with birds when it is in the heated condition. The resistance heating element for anti-icing and defogging consists of a transparent, electrically conductive film, evenly distributed over the inner surface of the outer pane of glass. The system includes the following additional components:  A windshield wire terminal box, located between the windshield panels  A temperature-sensing element embedded in each panel  Two windshield autotransformers and a heat control relay  A dual windshield control unit  A windshield heat control toggle switch located in the crew station The system receives power from the AC buses through the windshield heat control circuit breakers. When the Windshield Heat Control Switch is in HIGH, 115 volts, 400 hertz goes to the left and right amplifiers in the Dual Windshield Control Unit. The Windshield Heat Control Relay then energizes, applying two phases of AC power at 200 volts, 400 hertz to the Windshield Heat Autotransformers. These transformers provide 218-volt AC power to the Windshield Heating Current Bus Bars through the Dual Windshield Control Unit Relays. The sensing element in each windshield has a POSITIVE temperature coefficient of resistance and forms one leg of a bridge circuit. (Some systems use a thermistor, which has a negative temperature coefficient, as a sensing element to control windshield temperature.)

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Figure 4-33 Windshield anti-icing system.

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When windshield temperature is above calibrated value, the sensing element has a higher resistance value than needed to balance the bridge. This decreases the flow of current through the amplifiers, and the relays of the control unit are de-energized. As the temperature of the windshield drops, the resistance value of the sensing element also drops. Now, the current through the amplifiers again reaches sufficient magnitude to operate the relays in the control unit, thus energizing the windshield heaters. When the windshield heat control s witch is in LOW, autotransformers provide 115 volts, 400 hertz ac to the left and right amplifiers in the dual windshield control unit. In this condition, the transformers provide 121-volt ac power to windshield heating current bus bars through dual windshield power relays. The sensing elements in the windshield operate in the same manner as described for high heat operation. The calibrated units maintain windshield temperature between 40°C and 49°C (105°F to 120°F). Wing and Tail Anti-Icing Some aircraft have a Thermal Anti-icing System that prevents formation of ice on the leading edge of the wing panels and tail surfaces. This system uses hot air to heat the leading edges. The hot air comes from a combustion heater in the wings and tail section, or from the compressor section of a jet engine. An electrically operated pump supplies fuel for the anti-ice heaters. Heater demand determines the amount of fuel the pump delivers. In some aircraft, thermostats automatically determine heater demands. Various types and combinations of solenoid-operated valves control fuel flow and airflow in the system. Wing and Tail Deicing Some aircraft may have air-inflated, rubber deicer boots on the leading edges of wing and tail surfaces. Air pressure or vacuum is alternately applied to these boots and cracks off any ice that has formed. Once the ice has cracked, the force of the airstream peels it back and carries it away. Pressure for inflating the air cells in the deicer boots is normally from engine-driven pumps. Air pressure or vacuum is alternately applied either by the use of motor-driven rotary distribution valves or by the combination of an electronic timer and solenoid distributor valves. Empennage Deicing The P-3 empennage deicing system combines both anti-icing and deicing for the vertical and horizontal stabilizers. The system uses electrical power to prevent or remove the accumulation of ice. Electrical heating elements are in the leading edges and surfaces of the empennage. The system has ac power and dc control. Parting strips in the leading edge of each stabilizer accomplish anti-icing requirements. The parting strips remain on once the system actuates. Twenty cyclic heat areas accomplish deicing requirements. During normal operation, each of the cyclic areas heat for 8 seconds and are off for 168 seconds. Overheat sensors and a thermal-sensitive relay protects the system. Any time the system detects an over heat, power is shut off automatically to prevent damage to the metal structure. Pitot Tube Anti-Icing To prevent the formation of ice over the opening in the pitot tube, a built-in electric heating element (Figure 4-34) is used. A switch on the pilot’s console controls power to the heaters. System power is from either the ac or the dc bus. Exercise caution when ground checking the pitot tube since the heater cannot operate for long periods unless 4-62

Figure 4-34 — Pitot tube anti- icing circuit. the aircraft is in flight. Also, the danger exists for ground personnel to be accidentally burned.

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End of Chapter 4 AIRCRAFT ELECTRICAL SYSTEMS Review Questions 4-1. What are the two primary purposes of aircraft lighting?

A. Provide illumination for night navigation and identification B. Provide specialized exterior lighting and illuminating the interior C. For manual lighting D. To indicate normal operation and control of electrical systems

4-2. Why should spare light bulbs be shock mounted?

A. To prevent filament fatigue failure. B. To keep bulbs clean C. To make sure bulbs lock in the socket D. To prevent bulbs from contacting wires

4-3. What type of light is used to attract visual attention to the aircraft’s position and heading at night?

A. Reflector B. Anti-collision C. Globular D. Navigation

4-4. What type of aircraft is identified by a bluish-green anti-collision light?

A. Sear ch and Rescue B. Tanker C. Reconnaissance D. Emergency

4-5. When the angle of attack is perfect for the landing approach, what color AOA light illuminates?

A. Green B. Red C. Amber

D. Flashing green

4-6. As an AE your primary concern with hydraulics will be to ________.

A. convert fluid pressure to work B. schedule precise amount of fluid flow C. increase reliability D. control the flow of the fluid

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4-7. In a hydraulic system, what unit creates fluid flow?

A. Check valves B. Pump C. Actuating cylinder D. Selector valve

4-8. What does the hydraulic actuating unit convert fluid pressure into?

A. Thermal expansion B. Bleeding air C. Useful work by linear/reciprocating mechanical motion D. Electrical energy

4-9. Why does the arresting gear hook relay have a 1.1 second delay?

A. To ensure the hook is up and locked before removing hydraulic pressure B. To move the up latch mechanism forward to receive the hook latch C. To ensure current is sent through hook relay D. To break the circuit to selector relay

4-10. During emergency retraction, what is used to close the speed brake?

A. Speed brake relay B. Airstream pressure C. Selector valve D. Solenoids

4-11. The frictional temperature increase created by ram compression on the skin surface of the aircraft is known as ________.

A. extreme airspeed B. adiabatic temperature C. ram air temperature D. low-speed temperature

4-12. What is the primary purpose of an air-conditioning and pressurization system?

A. Crew safety and comfort B. Equipment reliability C. System control and operation D. Electrical systems optimal operation

4-13. The minimum requirements for navigation lights on all military heavier-than-air aircraft are established by what agency?

A. Department of Defense B. Department of Transportation C. Federal Transportation Administration D. Federal Aviation Administration

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4-14. Flight safety is the primary purpose of which of the following lights?

A. Position B. Formation C. Fuselage D. Anti-collision

4-15. What lights provide the pilot with angle- of-attack information?

A. Approach lights B. Indexer lights C. Instrument lights D. Position lights

4-16. With reference to the indexer lights, an inverted V indicates to the pilot that the angle of attack is in which of the following positions?

A. Slightly low B. Slightly high C. Very high D. Very low

4-17. What is the function of the arresting gear override switch?

A. It activates a light showing the LSO that the arresting hook is extended for landing B. It allows the approach lights to signal that the aircraft is unprepared to land C. It allows the approach lights to function properly while the arresting hook is up D. It shows the pilot the position of the arresting hook

4-18. What is the purpose of the in-flight refueling probe light?

A. To illuminate the drogue of the refueling aircraft only B. To illuminate the probe of the aircraft being refueled only C. To illuminate the probe of the receiver aircraft and the drogue of the refueling aircraft D. To indicate fuel flow and stoppage of fuel flow into the aircraft being refueled

4-19. When a controllable light is mounted in the nose of a helicopter, it has what total number of degrees of azimuth travel?

A. 90° B. 180° C. 270° D. 360°

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4-20. When replacing aircraft interior lamps or light covers, you should make sure that replacements meet the same specifications as the originally installed units for which of the following reasons?

A. Lighting fixture will require aircraft wiring alterations B. Aircraft’s power circuitry will be overloaded C. Original specifications were based on scientific considerations of necessity and crew comfort D. Equipment replacement supply inventory will increase

4-21. What are the advantages of the grain- of-wheat instrument lamps over other types of lamps used in instrument systems?

A. Longer life only B. More rugged only C. Better illumination only D. Longer life, more rugged, and better illumination

4-22. Which of the following statements describes the push- to-test feature on warning lights?

A. It provides a means for checking the condition of the warning light bulb only B. It provides a means for checking the system’s circuits only C. It provides a means for checking the system’s circuits and the condition of the warning light bulb D. It provides a means for momentarily activating all equipment and circuits in the respective systems

4-23. All hydraulic systems contain a minimum of which of the following basic components?

A. Pump, selector valve, actuator, and reservoir B. Pump, pressure regulator, switch, and reservoir C. Selector valve, filter, pump, and actuator D. Selector valve, actuator, pressure lines, and return lines

4-24. The Wing Flap system includes two Wing Flap Asymmetry Detectors that are located on ________.

A. each wing flap B. each wing rear beam C. the main wing fold D. leading wing tip

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4-25. What provides electrical signals to the Position Indicator of the Wing Flap Position Indicator System?

A. Rudder Boost Shutoff Gearbox B. Position transmitter C. Synchro transmitter D. Flap Drive Control

4-26. Which of the following functions is common to both the left and right main gear torque-link switches?

A. Preventing the throttles from being placed in the reverse propeller range while airborne B. Furnishing power for the landing gear control lever locking solenoid C. Disabling wing station external stores circuits D. Energizing the bomb bay door control circuit

4-27. The flashing WHEELS indicators are controlled by the downlock switches and ________ lever switches.

A. engine power B. engine control C. gear downlock D. wheels warning indicator

4-28. Retraction of the arresting hook is electrically controlled and hydraulically actuated; however, extension of the hook is accomplished by the use of what kind of power?

A. Hydraulic only B. Electrical only C. Electrical and hydraulic D. Mechanical

4-29. What is the purpose of the time delay relay in the relay panel?

A. To dampen hydraulic pressure surges when the system is first engaged B. To ensure that the arresting hook is fully up and locked before hydraulic pressure is removed C. To prevent the arresting hook from dropping if the handle is inadvertently moved to the down position D. To ensure that the arresting hook is completely down before hydraulic pressure is applied

4-30. The launch bar control switch is placed to retract, and the warning light remains on. What component failure is indicated?

A. Valve position switch B. Weight on gear switch C. Selector valve D. Control switch 4-68

4-31. Which of the following is a description of the action of a speed brake control switch?

A. Speed brake extension will stop when the switch is released from the OUT position B. Speed brake retraction will stop when the switch is momentarily held in the IN position C. The speed brake moves fully to the indicated position when the switch is momentarily held in either the IN or OUT position D. The degree of extension or retraction is controlled by the degree of movement of the switch lever to or from the STOP position

4-32. All of the following conditions create a demand for cabin air conditioning in aircraft flying at extreme airspeeds. Which one is the principal cause of cabin temperatures rising above the level at which the crew can maintain top physical and mental efficiency?

A. Engine heat B. Solar heat C. Body temperature D. Ram air friction

4-33. Temperature that is measured from a point at which there is no molecular motion is known as the ________ temperature.

A. standard B. absolute C. critical

D. ambient

4-34. The Air cycle air -conditioning system maintains the crew station air temperature at a comfortable and safe level by forcing which of the following kinds of air through crew station ECS Louvers?

A. Hot engine bleed air and ambient refrigerated air B. Hot engine bleed air and dry refrigerated air C. Refrigerated air and cabin pressure air D. Ambient refrigerated air and humidified hot engine bleed air

4-35. With the crew station ECS MODE switch in the AUTO position, the desired crew station temperature is maintained by which of the following means?

A. Variation in the opening of the cabin flow valve B. Proportional amounts of engine bleed air and refrigerated air being mixed by the condenser LCS no. 1 heat exchanger C. Proportional amounts of ram air and hot engine bleed air being mixed by the avionics flow valve D. Proportional amounts of ram air and refrigerated air being mixed by the secondary heat exchanger bypass valve

4-69

4-36. When the crew station MODE switch is in the OFF/RAM position, what are the conditions of the system flow modulating pressure regulator valve and the secondary bleed air pressure regulating shutoff valve?

A. The system flow modulating pressure regulator valve is closed and the secondary bleed air pressure regulating shutoff valve is open B. The system flow modulating pressure regulator valve is open and the secondary bleed air pressure regulating shutoff valve is closed C. Both system flow modulating pressure regulator valve and the secondary bleed air pressure regulating shutoff valve are closed D. Both system flow modulating pressure regulator valve and the secondary bleed air pressure regulating shutoff valve are open

4-37. Cabin temperature changes are anticipated by what component(s)?

A. Cabin temperature sensor only B. Cabin temperature and duct sensors C. Cabin duct dual temperature sensor D. Temperature control wheel

4-38. What is the function of the cabin duct limit bridge?

A. To limit the temperature of the cabin inlet air B. To anticipate sudden cabin temperature changes C. To select the desired cabin temperature D. To override the cabin duct anticipator bridge

4-39. What is the purpose of the voltage from the feedback potentiometer in the cabin dual temperature control valve?

A. To reduce the starting voltage once the valve actuator motor has started rotating B. To ensure the error feedback signal to the modulator circuit is the correct phase C. To prevent oscillations of the valve actuator motor D. To cause the valve actuator motor speed to increase, giving more positive control to the temperature regulation

4-40. What is the location of the heating element for windshield anti-icing and defogging?

A. On the outer surfaces of both glasses B. On the outer surface of the outer glass C. Between the inner surface of the outer glass and the vinyl plastic core D. Between the inner glass and the vinyl plastic core

4-70

4-41. What do aircraft equipped with wing deice systems use to prevent ice buildup on the leading edge?

A. Air pressure or vacuum alternately applied B. Hot air from combustion heaters or engine compressors C. Airstream air pressure D. Ram air temperature

4-42. What does the P-3 empennage anti/deicing system use to prevent ice buildup?

A. Airstream air pressure B. Ram air temperature C. Vacuum and air pressure D. Electrical heating elements

4-43. What method is used to deice the empennage of P-3 aircraft?

A. A powered, controlled system provides constant heat to the empennage B. Hot engine bleed air is circulated under the surfaces C. The leading edges and surfaces of the empennage are electrically heated D. Heat from the cabin routed through control valves heats the empennage

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4-72

CHAPTER 5 AIRCRAFT POWER PLANT ELECTRICAL SYSTEMS Every type of aircraft has unique propulsion requirements and each has a specific power plant system to meet that need. The performance of every power plant and propulsion system is dependent upon multiple electrical components and systems to include ignition, fire warning, fire extinguishing, anti-icing, fuel control, and indicating systems. It is the responsibility of the Aviation Electrician’s Mate (AE) to ensure these systems operate at optimum performance for maximum engine efficiency. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Identify the types of starting equipment used to start aircraft engines. 2. Recognize operating parameters and characteristics of aircraft engine ignition systems. 3. Describe operating conditions and characteristics of aircraft engine temperature control systems. 4. Recognize operating parameters and characteristics of aircraft engine starting systems. 5. Explain the operating principles and characteristics of aircraft power plant anti- icing and deicing systems. 6. Recognize operating conditions and characteristics of aircraft engine fire warning and extinguishing systems. 7. State the operating parameters and characteristics of aircraft fuel transfer systems. 8. Explain the operating principles and characteristics of aircraft engine oil temperature control systems. 9. State the operating conditions and features of aircraft engine variable exhaust nozzle control systems. 10. Explain the operating principles and features of propeller synchrophasing systems. 11. Explain the operating principles and features of aircraft propeller control systems. 12. State the operating parameters and features of aircraft approach power compensator systems. 13. Recognize operating principles and characteristics of helicopter bladefold systems. 14. Explain the operating principles and characteristics of aircraft engine inlet bleed air systems.

5-1

Figure 5-1 — Airframe Mounted Accessory Drive AMAD. STARTING EQUIPMENT Jet engine starters provide high starting torque initially to overcome the engine rotor weight and high speed to increase rotor revolution per minute (RPM) until the engine is self-sustaining. The following paragraphs describe the various starting systems used on turbojet, turboprop, and turbofan engines. Airframe Mounted Accessory Drive (AMAD) The starting system on the F/A-18 aircraft is the two AMAD interchangeable gearboxes (Figure 5-1), each mechanically connected to the engines, and pneumatically connected to the Auxiliary Power Unit (APU). The AMAD transmits power from the Air Turbine Starter (ATS) to the engine for starting or motoring. It has three separate modes of operation: 1. Main Engine Start (MES) Mode. 2. Crossbleed Engine Start. 3. Ground Maintenance Mode (GMM).

In GMM mode, the engine is decoupled from the AMAD by a decoupling mechanism mounted to the bottom of the AMAD and accessed by the operator through a panel on 5-2

the lower section of the aircraft, forward of the engines. The ATS converts pneumatic power from the APU to mechanical power which operates the AMAD and drives the AMAD accessories such as the generators, hydraulic pumps, ATS, and Motive Flow Boost Pumps (MFBP). Only one engine can be operated in GMM as the de-couple switches prevent both to operate simultaneously. Air for the starting operation of the aircraft is received from the other engine or from an external air source. The AMAD consists of two identical AMAD units and connects to each engine through a mechanical Power Turbine Shaft (PTS). The high speed, highly stressed, and dynamically balanced PTS shaft mounts to the Main Engine Gearbox coupling which is mounted directly to the engine. Main Engine Start (MES) Mode In the MES mode of operation, an external air source or operating APU provides airflow to the system. With battery power applied, when the ENG CRANK Switch is set to right R position, signals to the Frequency Sensing Relay (FSR) energize the right ENG CRANK Switch holding coil for approximately 8 to 12 seconds. The right Air Turbine Starter Control Valve (ATSCV), mounted to the right Air Turbine Starter (ATS) opens at a controlled rate limiting the pressure rise to 15 psi per second up to a regulated pressure of 45 to 51 psi at the right ATS. The right ATS converts the pneumatic power to mechanical power and transmits through a clutch in the right ATS and directs air from the air isolation valve to rotate the right AMAD, right PTS and right engine. A Monopole Speed Sensor in the right ATS transmits equivalent rpm signals to the FSR to continue output power for the right crank switch holding coil. A rotational signal must be transmitted to the FSR within the 8 to 12 seconds or the FSR releases the right ENG CRANK Switch holding coil and the crank operation completely shuts down. When the right AMAD accelerates and the right generator achieves online speed, the right Generator Power Contactor and right Generator ON Relay energize causing right crank operation to stop and the APU, ready for left engine start, returns to standby power. After left engine completes cranking operation and the left Generator Line Contactor or left Generator ON Relay fails to terminate the left crank operation, then the FSR releases the left ENG CRANK switch holding coil between 61.5 and 63.5 percent rpm compressor speed ( and operation terminates. Once the left and right engines are both started and both generators are online for 60 seconds, the APU will automatically shut down. For ATS protection, if the left or right ATS fails to disengage from the left or right engine during crank operation and turning speed exceeds 14.2 percent rpm the left or right ENG CRANK Switch returns to OFF, and L ATS or R ATS caution will display on the Digital Display Indicator (DDI). If L or R ATSCV fails in open position, a L or R ATSCV caution is displayed and maintenance code 818 (left) or 819 (right) sets in the Nose Wheelwell Digital Display Indicator (NWWDDI) after 25 seconds. Left and right AMAD oil temperature is monitored by the respective AMAD Oil Temperature Sensor. If AMAD oil temperatures exceed 190 ° F (88 °C), R OIL HOT or L OIL HOT cautions will appear on LDDI. The caution will remain until oil temperature drops below 175 ° F (79 °C). Both AMAD oil pressures are monitored by the AMAD Oil Pressure Switches. If either AMAD oil pressure drops, the oil pressure switch opens by 110 psi and sets maintenance code 816 (left) or 817 (right) in NWWDDI, and a L AMAD PR or R AMAD PR caution appears on the LDDI. 5-3

Engine Drive Mode With aircraft engines running, the PTS transmits power from each engine to its respective AMAD. Power is transmitted to each AMAD accessory by way of the respective gear train in each AMAD. This provides the aircraft with electrical power from the generator, hydraulic power from the hydraulic pump, and engine fuel flow from the MFBP. Crossbleed Engine Start The APU compressed air is the primary engine crank air source. With a single engine on line, a second engine can be started by utilizing the compressed air from the previously started engine. The operating engine should be advanced to a minimum of 80 percent rpm to make sure bleed air output is sufficient to crank the opposite engine. With the APU switch OFF, for a crossbleed start, the operator sets the L or R engine crank switch to the non-operating engine crank position where 28vdc energize the desired ATSCV and opens the air isolation valve, allowing compressor bleed air pressure to turn the non-operating ATS, AMAD and engine. Once a rise in engine rpm occurs, at a minimum of 10 percent , the throttle is advanced to IDLE for the non- operating engine and light-off begins to make faster engine rotation. Once the non- operating engine is started, the operator returns the previous operating engine back to IDLE position. Ground Maintenance Mode (GMM) In GMM, with the APU running, the engine is decoupled from the respective AMAD by a decoupling (pull down and turn) mechanism. The ATS converts pneumatic power from the APU to mechanical power which operates the AMAD and drives AMAD accessories. Either L or R AMAD can be operated in GMM. GMM can also be powered by an external air source. Operating the Decouple Handle enables GMM, which uses compressed air to drive the decoupled AMAD through the respective ATSCV and ATS. The AMAD Couple Switch opens when the AMAD is decoupled. This prevents the use of cross bleed air for GMM. Setting the ENG CRANK Switch to L or R positions opens the selected ATSCV allowing compressed air to drive the ATS. During GMM, the FSR controls Pneumatic Control Unit (PCU) torque motor power, PCU torque motor controls the ATSCV opening by venting pressure from the open side of the actuator piston, and PCU Pneumatic Valve operation regulates the ATSCV airflow, which maintains the ATS speed at 56.6 to 58.6 percent rpm. Operating R AMAD in GMM recharges the APU accumulator by providing Hydraulic System 2B pressure. Approximately 30 seconds after starting GMM operation, the operator will see a L ATS or R ATS c aution displayed on LDDI, which is a normal indication. When ATS speed exceeds 14.2 percent rpm, a caution is displayed. This again is a normal condition because during GMM the engine is decoupled from AMAD and engine rpm remain at zero. If the AMAD oil overheats when in GMM, the Signal Data Computer (SDC) automatically shuts down the system. Air Turbine Starter (ATS) The ATS is a lightweight unit designed to start turbojet, turboprop, and turbofan engines when supplied with compressed air. The unit consists primarily of a scroll assembly, rotating assembly, reduction gear system, overrunning clutch assembly, and output shaft. An overspeed switch mechanism limits maximum rotational speed. As you read this section, refer to Figures 5-2 and 5-3. Compressed air, supplied to the scroll inlet, goes to the turbine wheel through the nozzle in the scroll assembly. The reduction gear 5-4

Figure 5-2 — Air turbine starter. system transforms the high speed and low torque of the turbine wheel to low speed and high torque at the output shaft. When at the desired starter rotational speed, the flyweights in the governor assembly throw out and open the limit switch. This switch sends a signal that shuts off the supply air. At a higher, predetermined rotational speed, the overrunning clutch assembly releases the output shaft from the rotating assembly. A source of compressed air flows to the shutoff valve inlet duct to drive the starter. The starter is a turbine air motor equipped with a radial inward-flow turbine wheel assembly, reduction gearing, splined output shaft, and a quick-detaching coupling assembly. The complete assembly mounts within one scroll assembly and gear housing (Figure 5-2). The ATS converts energy from compressed air to shaft power. This power goes to a splined output shaft at speed and torque values for starting aircraft engines. 5-5

Figure 5-3 — Air turbine starting system diagram. Initial control of the air shutoff valve (Figu re 5-3) is by a normally open, momentarily closed, start switch and a relay box. After pressing the start switch, the sequence of operation of the valve and starter is automatic. A normally closed, momentarily open, stop switch provides a means of manually stopping the starter when motoring an engine without fuel or in emergencies. When the external start switch momentarily closes, a double-pole, single-throw, holding relay in the relay box actuates. This action completes electrical circuits to the air shutoff V alve and the starter. When the external start switch opens, the holding relay receives a positive potential through the normally closed external stop switch. The relay also receives a negative potential through the closed overspeed control provided within the starter. The relay continues to hold until the external stop switch removes the positive potential or until the closed overspeed control removes the negative potential. 5-6

NOTE If engine shaft speed fails to exceed starter driving mechanism speed before the starter reaches cutoff speed, the cutout switch actuates, shutting down the starter.

CAUTION During operation of the starter, you should stand clear of the plane of rotation of the high-speed rotating turbine wheel. Only qualified ADs should install and service the unit and its pressure regulating valve. When high-pressure air is at the closed regulating valve inlet and the start switch energizes the holding relay, the regulating valve opens, admitting compressed air to the starter. The control mechanism regulates the compressed air to specified conditions. The mechanism senses upstream and downstream conditions, and it positions the valve butterfly to supply the desired flow. The compressed air enters the inlet port of the starter and expands as it flows radially inward through the nozzle vanes. The air flows against the blades of the turbine wheel to rotate them. The reduction gear system (Figure 5-2), which transmits power to the drive shaft, converts the high-speed, low-torque output of the turbine wheel to low- speed, high-torque output. As compressed air enters the starter inlet port through the pressure regulating valve, the turbine wheel rotates, transmitting torque to the drive jaw through reduction gearing. This torque transmits to the drive shaft through the splined drive shaft. When aircraft engine speed exceeds the starter drive shaft speed, the speed of the starter output shaft (directly connected to the engine drive) also exceeds the drive jaw speed, and the pawls begin to ratchet. When the output shaft (driven by the aircraft engine) attains enough speed, centrifugal force releases the pawls completely from the drive jaw. This action releases the starter from the aircraft engine. When the starter reaches cutoff speed, the internal overspeed control actuates, breaking the electrical circuit to the holding relay. Then, the pressure regulating valve butterfly closes and prevents compressed air from entering the starter.

A starting operation begins by momentarily closing the start switch to energize the pressure regulating valve circuit. Sequencing of the regulating valve and start operation becomes automatic. The start continues until engine light-off occurs, and the engine overspeeds the starter driving mechanism or until output shaft speed reaches the calibrated cutoff point. In either condition, disengagement of the starter from the engine or interruption of the supply air to the starter is automatic. Starter operation also stops when the stop switch momentarily opens, which closes the pressure regulating valve and interrupts the supply air to the starter.

5-7

Figure 5-4 — Cross-sectional view of a jet igniter plug. WARNING

Due to the high voltage and amperage of ignition systems, you should use extreme caution around the equipment. IGNITION SYSTEMS Three things are necessary to cause a fire: a combustible material (such as aircraft fuel), oxygen, and heat. A fire will not start without all three, and removing any one of the three puts the fire out. All internal combustion engines use fire to produce mechanical energy, and the piston engine uses the higher degree of fire—an explosion. The gas turbine (jet) engine also produces its energy through the use of fire. However, its operation is considerably different from the piston engine. Rather than a series of independent explosions, a jet engine produces a continuous burning fire. Ignition is necessary only during the start cycle to ignite the fire. Electronic ignition systems provide internal combustion for turboprop, turbofan, and turbojet engines. Unlike reciprocating engine systems, timing is not a factor in turbine- power ignition systems. All that is needed is a series of sparks with enough intensity to cause combustion. The exciter develops voltage of sufficient amplitude to produce a spark. The exciter unit contains a capacitor or capacitors to develop the voltage and current necessary to supply a spark plug (called an igniter). The resultant spark is of high heat intensity, capable not only of igniting abnormal fuel mixtures but also of burning away any foreign deposits on the plug electrodes. The exciter is a dual unit and produces sparks at each of two igniter plugs. The igniter plugs are, in general, similar to the spark plugs on reciprocating engines. The main differences are features necessary to operate at higher energies, voltages, and temperatures of jet engines. In general, the igniter plug is larger, more open in construction, and the gap is much wider than spark plugs of familiar design. Figure 5-4 shows a typical jet igniter plug. J et ignition is controlled through relays or switches that operate automatically during the engine start cycle. Fuel or oil pressure switches or centrifugal speed switches energize a relay to begin ignition. Ignition stops by actuation of a centrifugal switch at a speed between 45 percent and 65 percent of rated engine speed.

5-8

Figure 5-5 — Jet engine electronic ignition system. Electronic Ignition System The development of more powerful jet engines demands a reliable, maintenance-free ignition system. This chapter does not cover all ignition systems; rather, the system described represents most modern systems. An electronic ignition system has an advantage over the capacitor discharge system; it has no moving parts and breaker points or contacts that can become pitted or burned. The engine ignition system (Figure 5-5) provides the necessary electrical energy and control to begin engine combustion during aircraft armament firing and starting, and for automatic re-ignition in case of engine flameout.

Engine Ignition Exciter The engine ignition exciter is a dual-circuit, dual-output unit that supplies a high-voltage, high-energy electrical current for ignition. The exciter consists of a radio frequency interference filter and two power, rectifier, storage, and output elements. The exciter mounts on the forward part of the compressor section of the engine. 5-9

Engine Control Amplifier The engine control amplifier is the electronic control center of the engine. It controls the function of the ignition system as well as other engine operational functions. The amplifier mounts on the compressor section aft of the engine front frame. Engine Ignition Leads and Igniter Plugs The ignition leads are high-tension cables, which transmit electrical current from the exciter to the igniter plugs. The igniter plugs mount in the combustion chamber housing. Engine Alternator Stator The engine alternator stator is an engine-driven, single-phase, alternating current (ac) electrical-output unit mounted on the engine accessory gearbox. It supplies electrical power to the engine, independent of the aircraft electrical system. It contains three sets of windings. Two windings supply electrical power to the ignition exciter, and the third supplies electrical power to the control amplifier. Ignition Operation As you read this section, refer to Figure 5-5. With the ignition switch ON, the engine cranking for starting, and throttle advanced to 10-degree Power Lever Angle (PLA) position, current flows from the alternator stator to power the control amplifier. At the same time, the PLA ignition switch in the fuel control closes. The gas generator ( speed logic circuit closes the ignition relay to provide ignition when is within the 10 to 48 percent range. With the relay closed, it completes a circuit from the alternator stator ignition windings, through the ignition exciter, to the igniter plugs. Current flows from the alternator, through the control amplifier, to the ignition exciter. At the ignition exciter, current is intensified and discharged as a high-voltage output, and conducts through the igniter cables to the igniters. Current crossing the gaps in the igniters produces a continuous high-intensity spark to ignite the fuel mixture in the combustion chamber. When engine speed reaches 8,500 RPM, and Inter-Turbine Temperature (ITT) reaches operating range, a signal from the T5 temperature detectors flows through the T5 circuit to the control amplifier ignition logic circuit. The control amplifier ignition relay opens and ignition ends. Combustion then continues as a self- sustaining process. Ignition automatically reactivates when either a flameout occurs or when aircraft armament fires. When T5 temperature drops more than 800°F (427°C) from T5 selected by PLA, the T5 detectors signal control amplifier T5 flameout logic to close the amplifier ignition relay. This activates ignition system operation. Ignition continues until engine operating temperature is again normal and the 800°F temperature error signal cancels, causing the control amplifier to end ignition operation. An armament-firing protection circuit prevents flameout from armament gas ingested by the engine during armament firing. When firing aircraft armament, a signal from the armament trigger switch activates the armament-firing logic circuit in the control amplifier. The amplifier logic circuit causes ignition operation to activate. Ignition operation ends after a 1 second time delay in the amplifier logic circuit following release of the armament firing trigger. 5-10

ENGINE TEMPERATURE CONTROL SYSTEMS On most reciprocating engines, the engine cowl flaps control cylinder head temperature (CHT). In turbine-powered engines, turbine temperature is controlled differently. Engine temperature is a measure of power, and temperature is a product of fuel consumption. In most turbine-powered aircraft, then, the pilot selects desired power through a mechanical linkage to the fuel control. As the power increases, so does the temperature; in turbojet aircraft, this also causes an increase in engine speed. The only electrical circuits required are those to show temperature and speed except in newer aircraft such as the F/A-18E/F/G which use an electrical throttle quadrant control. The engine temperature control system on turboprop engines lets the operator control turbine inlet temperature and torque through the use of power and condition levers. These levers connect to each engine coordinator through pushrods, sectors, cables, and pulleys. When the engine is operating in the flight range, engine speed is constant. Engine power is controlled by increasing or decreasing fuel flow, which results in a corresponding change in turbine inlet temperature. The main components of the engine temperature control system are:  Power levers  Condition levers  Engine coordinators  Temperature datum controls  Turbine inlet thermocouples, and temperature datum switches.

5-11

Figure 5-6 — Engine temperature control (turboprop) system block diagram. Figure 5-6 shows the block diagram of an engine temperature control system. You should refer to it while you study this section.

Power Levers The power levers (one for each engine) can move separately or together to control engine power. The range of power lever settings is from REVERSE (reverse thrust) to MAX POWER (takeoff). Power lever switches within the cockpit pedestal supply electrical power to other systems. A detent at the FLT IDLE position prevents inadvertent movement of the power levers below FLT IDLE while airborne. To move the power levers to the taxi range, the levers must be raised from the detent. During a catapult-assisted takeoff, a retractable catapult grip helps the pilot maintain the power levers at MAX POWER. Condition Levers The condition levers are located next to the power levers on the cockpit pedestal. They have four positions— FEATH, GRD STOP, RUN, and AIRSTART. Switches at each condition lever position complete electrical circuits for other systems. The pilot must raise the detent release handle of each condition lever to move the levers to different positions. A detent holds the lever at FEATH, GRD STOP, or RUN. When the condition 5-12

lever is in the AIRSTART position, the propeller unfeathers and the engine starting cycle begins. The lever is held in the AIR START position until the engine speed reaches 100 percent RPM. Then, the lever is released, springs back to RUN and remains there for normal operation. When set to RUN, the condition lever positions the mechanical linkage to open the fuel shutoff valve. A mechanical stop in the pedestal prevents both condition levers from being set to FEATH at the same time. When set to FEATH, the condition lever electrically and mechanically closes the corresponding fuel shutoff valve and feathers the propeller. At GRD STOP, the condition lever electrically closes the fuel shutoff valve to shut down the engine. Engine Coordinators The coordinators are mechanical devices that coordinate the power and condition levers, propeller, fuel control, and electronic fuel trimming circuit. One engine coordinator mounts on each fuel control. The main components of a coordinator are a variable potentiometer, a discriminating device, and a cam-operated switch. A scale calibrated from 0 to 90 degrees attaches to the outside case, and a pointer secures to the main coordinator shaft. Pushrods, connected from the coordinator to a cable sector, transmit power and condition lever movement to the coordinator. Power lever movement through the coordinator changes resistance of the Variable Potentiometer and changes the temperature datum control temperature reference signal. The Cam- Operated Switch changes the temperature datum control from temperature limiting to temperature controlling with power lever above 66-degree coordinator and engine speed above 94 percent RPM. Power lever movement transmits to the coordinator, propeller, and fuel control through a series of rods and levers. With the condition lever in FEATH, the Discriminating Device mechanically positions propeller linkage toward feather and closes the fuel shutoff valve, regardless of the power lever setting. The temperature datum control consists of electronic units that automatically compensate for changes in fuel density, manufacturing tolerances in fuel controls, and variations in engine fuel requirements between engines. With the power lever above 66- degree coordinator (temperature controlling range) and the TEMP DATUM switch in AUTO, the temperature datum control compares the actual turbine inlet temperature signal and desired temperature reference signal. If there is a difference greater than 1.9 °C (4.5 °F), the control electrically signals the temperature datum valve to reduce or increase fuel flow to the engine. This action brings the turbine inlet temperature to the desired value. A damping voltage goes back to the control from a generator within the temperature valve motor, preventing overcorrection and stabilizing the system. When engine speed is above 94 percent RPM and the power lever is below 66-degree coordinator (temperature limiting range), the normal limiting temperature automatically becomes 978 °C (1,792 °F). However, when engine speed is below 94 percent RPM, regardless of power lever position, the limiting temperature is 830 °C (1,524 °F). This prevents high turbine inlet temperature during starting and acceleration when the compressor bleed valves are open. Dual-unit thermocouples are mounted radially in the turbine inlet case of each engine. The junction portion of the thermocouples protrudes through the case to sense the gas temperature before the gas enters the turbine section. Four leads, two of Chromel and two of Alumel, connect to each thermocouple to form two independent parallel circuits. One circuit connects to the cockpit turbine inlet temperature indicator. The second circuit supplies the temperature datum control with temperature signals for the electronic fuel trimming circuit. As the gases heat the thermocouples, they generate an electromotive force that goes to the cockpit indicator and the temperature datum control. 5-13

Because the thermocouple connections are in parallel, the signal they send is the average temperature of the thermocouples. If one parallel circuit fails, the other circuit continues to operate normally. Temperature Datum Switches The left and right engine temperature datum (TEMP DATUM) switches are on the engine control panel in the cockpit. Each switch has AUTO and NULL positions. When the switch is in AUTO, the engine RPM is above 94 percent and the engine coordinator is above 66 degrees, the temperature datum control compares the turbine inlet temperature to a reference temperature. If the temperatures differ, the temperature datum control electrically signals the temperature datum valve to bypass more or less fuel from the engine to bring turbine inlet temperature to the selected value. If the electronic fuel trimming circuit malfunctions, position the TEMP DATUM switch to NULL. The circuit de-energizes and the fuel control, through movement of the power lever, controls turbine inlet temperature. Over temperature protection is not available.

5-14

Figure 5-7 — Engine start control system. ENGINE START CONTROL SYSTEM In this section, you will learn about engine starting, engine ignition, and engine fuel temperature starting systems, in order to understand how they interrelate. The engine start control system covered in this section is found in the P-3C aircraft. Major Components Major components of the engine start control system include the air turbine starter, the speed-sensitive control, the ignition exciter, the engine fuel pump and filter, the fuel control, the fuel nozzles, the fuel control relay, the starting fuel enrichment valve, the temperature datum valve, the drain valves, and the compressor bleed air valves. You should refer to Figure 5-7 and Table 5-1 as you read about these components.

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Table 5-1 — Engine Starting Sequence 5-16

Air Turbine Starter You have learned that the ATS is pneumatically driven and mechanically connected to the engine through a gearbox. The air turbine starter operates on compressed air from an external Gas Turbine Compressor (GTC), internal APU, or bleed air from an operating engine. The compressed air goes through a manifold, an engine isolation bleed air valve, and a starter control valve into the starter’s turbine. The engine start switch, located in the cockpit, controls the opening of the starter control valve, which allows compressed air to enter the ATS. The speed sensitive control, through a holding solenoid, holds the engine start switch on until engine speed reaches 65 percent RPM. Speed-Sensitive Control The speed-sensitive control, located on the engine, contains internal switches that activate at three predetermined intervals. These intervals are relative to the engine’s normal speed— 16 percent, 65 percent, and 94 percent of engine RPM. Activation of these switches controls many operations in the engine start cycle. Ignition Exciter The ignition exciter (discussed earlier) is a dual electronic ignition unit that uses 28-volt dc from the ignition relay. The exciter steps up the voltage to a proper level for firing the igniter plugs. The exciter unit contains two identical circuits, each one independently capable of firing its own igniter plug. The speed-sensitive control energizes the ignition relay so the exciter is in operation between 16 and 65 percent of engine RPM. Engine Fuel Pump and Filter The fuel pump and high-pressure filter assembly mount on the rear of the accessories case. This assembly consists of a centrifugal boost pump, two gear-type pressure elements, and a high-pressure filter. Fuel, entering the pump assembly, passes through the centrifugal boost pump, which will raise the pressure to a minimum value and pass fuel through the low-pressure filters before going to the secondary element. There is a differential pressure switch connected across the inlet and outlet of the filters. If the pressure differential exceeds 7.5 PSI, the switch closes and completes a circuit to a filter light at the flight deck. Fuel then flows to the primary element and through the high-pressure filter assembly before entering the fuel control. Both the low and high pressure filters have bypass valves that open if the filters become clogged. The capacity of the pump’s primary element is 10 percent greater than that of the secondary element. If the primary element were to fail, the secondary element would provide enough flow to operate the engine. During engine starting, the elements operate in parallel to provide enough fuel flow at low RPM; above 65 percent, they operate in series. Parallel operation occurs during starting when engine speed is between 16 percent and 65 percent RPM. If both elements are operating properly, the paralleling light will be on between 16 percent and 65 percent RPM only. If the secondary element fails, the light never comes on; if the primary element fails, the light is on above 65 percent RPM. 5-17

Fuel Control The fuel control is on the accessories drive housing and mechanically links to the coordinator. The fuel control provides a starting fuel flow schedule that, in conjunction with the temperature datum valve, prevents over temperature and compressor surge. The fuel control schedule is 20 percent richer than the nominal engine requirements to accommodate the temperature datum valve. The valve bypasses 20 percent of the control output when in the null position. This excess flow gives the temperature datum valve the capacity to add as well as subtract fuel. The valve is then able to maintain the temperature scheduled by the coordinator and the temperature datum control. The fuel control includes a cutoff valve for stopping fuel flow to the engine. It actuates either manually or electrically. During engine starts, the cutoff valve remains closed until the engine reaches 16 percent RPM. The speed-sensitive control then opens the cutoff valve, permitting fuel to flow to the engine. Fuel Nozzles The fuel output from the temperature datum valve flows through the fuel manifold to the six fuel nozzles. Fuel flows through both the primary and secondary nozzle orifices during normal operation. At low fuel flow rates, fuel flows through the primary orifice only. Fuel Control Relay The fuel control relay is a fail-safe-type relay that energizes when the FUEL and IGNITION switch is off, or when the propeller is feathered. With this arrangement, the engine can still operate with an electrical power failure during flight. The pilot shuts the engine down by placing the FUEL and IGNITION switch in the OFF position, or by feathering the propeller. Power then goes to the fuel control shutoff valve, which closes and stops all fuel to the engine. Starting Fuel Enrichment Valve The primer switch (with two positions, ON and spring-loaded OFF) operat es the fuel enrichment valve, providing increased fuel flow during engine starting. The primer switch must be in the ON position and held there before the engine reaches 16 percent RPM. Further, it must remain on until the fuel control shutoff valve opens at 16 percent RPM or enrichment will not occur. The enrichment (primer) valve closes when fuel pressure in the fuel manifold reaches 50 PSI. Fuel enrichment is needed only in very cold climates. Temperature Datum Valve The temperature datum valve is located between the fuel control and the fuel nozzles. It is a motor-operated bypass valve that responds to signals from the temperature datum control. If the power lever positions are between 0 degrees and 66 degrees, the valve remains in null and the engine operates on the fuel flow scheduled by the fuel control. The valve remains in null unless the temperature datum control signals it to limit turbine inlet temperature. The valve then reduces the fuel flow (up to 50 percent during starting, 20 percent above 94 percent RPM) by returning the excess to the fuel pump. When turbine inlet temperature is at the desired level, the temperature datum control signals the valve to return to the null position. In power lever positions between 66 degrees and 90 degrees, the temperature datum valve acts to control turbine inlet temperature to a preselected schedule corresponding 5-18

to power lever position. This is the temperature controlling range. In this range the temperature datum control may signal the valve to allow more (higher temperature desired) or allow less (lower temperature desired) fuel to flow. Drain Valves A spring-loaded, solenoid-operated manifold drain valve is located at the bottom of the fuel manifold. It drains the fuel manifold when fuel pressure drops below 8 to 10 PSI. This action minimizes the amount of fuel dropping into the combustion liners while the engine unit is being stopped. Compressor Bleed Air Valves The fifth and tenth bleed air valves release air from the compressor to reduce the compressor load during engine starts. During starting, the bleed valves are open up to 94 percent RPM. At 94 percent RPM, the speed-sensitive valve ports compressor- discharge air to close the bleed valves. Engine Start Cycle Operation The following sequence of events is typical of a normal engine start cycle. While reading this section, you should assume that external compressed air and electrical power are being applied to the aircraft. Also, assume that all other system switches are in the proper position for an engine start. Refer to Figure 5-7 and Table 5-1 throughout this discussion. The operator places the ENGINE START SELECTOR switch to the engine number 1 position. Position FUEL and IGNITION switch, Engine 1, to the ON position, de- energizing the fuel control relay. This allows power to pass through the contacts of the fuel control relay, the 16 percent speed-sensitive control switch, and the fuel manifold pressure switch to energize the temperature datum relay. When the operator depresses the ENGINE START switch, current flows through the engine start switch, engine start selector switch, starter control valve, and speed- sensitive control 65 percent switch to ground. Current also flows through the engine start switch holding coil to ground through the same 65 percent switch in the speed- sensitive control. With power applied to the starter control valve, the valve opens. This allows compressed air to flow to the air turbine starter. It also closes the contacts of the air valve position switch. The yellow starter valve lights illuminate to show the operator the starter control valve is open. The air turbine starter now causes engine rotation. If fuel enrichment is needed, the operator depresses the PRIMER switch holding it in the ON position until the engine reaches 16 percent. Power then goes through the primer relay contacts and the temperature datum relay contacts, energizing the primer valve solenoid. When the engine reaches 16 percent RPM, the speed-sensitive control mechanically actuates the 16 percent switch from 16 percent to 65 percent. Power then goes through the 16 percent switch contacts, energizing the ignition relay. Power then flows through the ignition relay contacts, energizing the fuel pump paralleling solenoid, the drip valve solenoid, and the ignition exciter. Power also goes to the fuel control shutoff valve, allowing fuel to enter the engine fuel manifold. Extra fuel for fuel enrichment (if used) flows to the temperature datum valve. The temperature datum relay remains energized 5-19

by a holding circuit consisting of the lower temperature datum relay contacts and the fuel manifold pressure switch. As engine speed increases, the fuel pressure increases. When fuel manifold pressure reaches 50 psig, the fuel manifold pressure switch opens, de-energizing the temperature datum relay, stopping fuel enrichment. When the secondary fuel-pump pressure exceeds 150 psig, a paralleling light illuminates showing parallel operation of the fuel pumps to the operator. At 65 percent RPM, the speed-sensitive control 65 percent switches open to de- energize the ignition relay. The ignition relay removes power from the ignition exciter, the fuel pump paralleling solenoid, and the drip valve solenoid. The fuel pumps now operate in series, and fuel pressure now holds the drip valves closed. The starter control valve and the engine start switch lose their common ground, and current flow ceases through those circuits. The starter control valve closes, stopping airflow to the air turbine starter, and the engine start switch opens. The air valve position switch opens, causing the starter control valve light to go out. This completes the engine start cycle. When engine speed increases above 94 percent, contacts in the speed-sensitive control (circuit not shown) de-energize the temperature datum valve take solenoid. This reduces the fuel take capability from 50 percent to 20 percent. The fifth and tenth stage bleed air valves also close now. When the power lever advances above 66 degrees of coordinator travel, temperature datum system switches from temperature limiting range to temperature controlling range. POWER PLANT ANTI-ICING AND DEICING SYSTEMS Naval aircraft deicing lets planes fly in any type of weather by protecting the power plant from ice buildup in freezing conditions. There are two electrical systems that do this— the anti-icing and deicing systems. Anti-icing systems prevent ice from forming and deicing systems remove ice that has already accumulated. Many types of anti-icing systems are used today. All systems use heated air from the engine to perform the anti-icing function. The use of heated air causes engine power loss, so use anti-icing only when necessary. In some aircraft, a reversible electric motor opens and closes an air valve to supply the needed air. In other aircraft, an electrical solenoid positions a pneumatic valve to allow regulated heated air into the engine anti- ice system. When missions dictate that aircraft fly routinely in adverse weather conditions, a fail- safe anti-ice system is used. Fail safe means the solenoid-actuated air valve electrically actuates closed. If the switch is turned on, or if electrical power fails, the valve is spring loaded to the open position. Some systems anti-ice the complete inlet duct; in other systems, only the guide vanes are anti-iced. Guide Vane Anti-Icing System There are a variety of engine anti-icing systems in use today. The system covered in this RTM is representative of several systems designed for Navy aircraft. Look at Figure 5-8. Here, you see that the electrical portion of the circuit serves only to turn the system on or off.

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Figure 5-8 — Inlet guide vane anti-icing system.

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The guide vanes of a turbine-powered engine direct the flow of inlet air into the compressor section. At this point, the air is coldest and most subject to icing. The biggest problem caused by ice forming here is blockage of inlet air, causing air starvation and thus engine failure. Also, there is a possibility that chunks of ice can be inducted into the engine. Therefore, turn on the anti-icing system at the first indication of any icing condition or before entering an icing condition. Normally, icing does not occur in supersonic flight because friction of the aircraft passing through the air creates enough heat to prevent ice formation. The anti-icing valve is a solenoid-operated bleed valve. With no electrical input to the solenoid, the bleed valve closes, and there is no anti-icing airflow through the valve. When the engine is operating with the valve solenoid de-energized, the main poppet will remain in the closed position. When the solenoid energizes, the solenoid valve unseats and permits air pressure within the main poppet to escape through the overboard vent. With pressure decreasing in the poppet valve body, inlet pressure on the main poppet valve face overcomes spring tension and raises the valve from its seat. This permits high-pressure air to discharge through the outlet of the valve to the anti-icing manifold on the engine. The regulating piston and spring valve assembly control discharge air from the anti-icing valve to a preset pressure. Propeller Deicing Systems One method of preventing excessive accumulation of ice on the propeller blades of turboprop or reciprocating engines is using electric heaters. Figure 5-9 is a simplified schematic diagram of a system for a two-engine aircraft.

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Figure 5-9 — Electrical deicing for a propeller system.

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The propeller deicing system consists of a three-position, two-speed selector switch (propeller deicer switch) and an indicator light, a two-speed timer, and two propeller deicer relays (one for each propeller). Also, included (for each propeller) are the brush pad bracket assembly, slip-ring assembly, the aft portion of the propeller assembly, and a neoprene rubber heating element and connector for each blade. Abrasion strips protect the blade heaters. The deicing system operates at either a slow cycle of 40-75 seconds on, 120-225 seconds off; or a fast cycle of 17-22 seconds on, 51-66 seconds off. The icing conditions during flight determine the switch position. Setting the selector switch to SLOW or FAST permits dc power from the essential bus to energize the deicer timer motor and turn on the indicator light. Resistances in the timer determine the speed of the timer motor. The motor, through reduction gears, causes the camshaft to rotate. This rotation positions the cam switches alternately between the right and left contacts. Current flows through these contacts to cycle their respective propeller deicer relays. With the relays being energized alternately, current from the three-phase generator ac buses flows through propeller deicer circuit breakers to the propeller brush pad bracket assemblies. Carbon brushes contact the copper slip rings, transmitting ac power through the slip rings to the blade heating elements. Placing the selector switch in the OFF position stops the propeller deicing operation, and the indicator light goes out. The propeller deice timer is a two-speed, automatically controlled timer. It regulates, in cycles, the time duration and sequence of electrical impulses to the propeller blade heating elements. The unit is located in a moisture proof, airtight case, which isolates the unit from temperature extremes and vibration. The deice timer consists of a fractional horsepower, constant-speed dc motor, including reduction gear, camshaft with three cams, three cam switches, two fixed resistors, and variable resistor. The unit also includes a filter to minimize radio interference. With the propeller deice switch set to FAST, direct current flows from the left dc bus, through the propeller deicer circuit breaker and switch, to the timer. This current follows two paths in the timer. One path, from pins E and F that connect in the timer, directs the current flow to the control cam switch. The other path, from pin G, directs the flow through the variable resistor and one fixed resistor to the timer motor, the filter, and to ground. The adjustment of the variable resistor determines the speed of the motor. The motor, through the 3,000 to 1 reduction gear, rotates the camshaft and cams. Two single-lobed shift cams and a single two-lobed control cam are on the camshaft. Positioning on the camshaft is so the two single-lobed shift cams are on either side of the two-lobed control cam. As the control cam rotates, it alternately makes and breaks its right and left cam switch contacts. This permits the flow of current to the shift cam switches. As the current flows to the other cam switch, rotation of the single-lobed cam makes and breaks the shift cam switch contacts. This action cycles first the right and then the left propeller deicer relays. With the propeller deicer switch set to SLOW, the operation is the same as the fast cycle with the one exception. Direct current enters the timer through a different pin (pin H) in the plug and flows through the two fixed resistors and the variable resistor to the motor. (Dc power to the control cam switch is through the same pins as for the fast cycle.) Because of the increase in resistance, the motor operates at a slower speed. Thus, with motor speed reduction, rotation of the camshaft, through the reduction gear, is slower, and the timer now functions at the slower cycle. 5-24

Figure 5-10 — Propeller heating elements. Several aircraft have anti-icing and deicing system that prevents the formation of ice (anti-icing) on the forward portion of the propeller spinner. The system removes any ice formation (deicing) from the blades and cuffs, aft portion of the spinner, and spinner islands. This system operates similarly to the system described previously, except that the anti- icing elements are on continuously and the deicing elements cycle. Figure 5-10 shows the location of the heating elements. The system usually contains a safety feature for testing the propellers on the ground. This feature provides a low voltage to the heating elements, which prevents damage to the prop from overheating.

FIRE WARNING AND EXTINGUISHING SYSTEMS Some turbine engines operate at temperatures of more than 1,000 °C. Fuel and oil lines run within a few inches of these extreme temperatures. For this reason, you must closely monitor the engine, and immediately take corrective action when an abnormal condition occurs. Performance of precision work is necessary when maintaining fire warning systems to ensure their reliability. An undetected fire may cost the lives of the aircrew and possibly millions of dollars in aircraft and equipment. In multi‐engine aircraft, a fire warning usually dictates that the engine should be shut down and the fire extinguished. The least that can happen if there is an erroneous fire warning is the aircraft will abort its assigned mission. Warning System The engine fire detector system is an electrical system for detecting the presence of fire or dangerously high temperatures in the engine(s) areas. The system for each engine consists of a control unit, test relay, signal lamp, test switch, and several sensing elements. The system uses a continuous strip of temperature-sensing elements to cover the paths of airflow in the engine compartment. The same engine fire warning systems are in multiengine aircraft, one for each engine. Look at Figure 5-11. It shows the electrical schematic for a fire warning system that is representative of systems found in modern Navy aircraft. The system is of the continuous-element, resetting type. The s ensing element consists of two conductors separated by a semiconductor. The outer conductor is at ground potential, and the center conductor connects to an amplifier input in the fire detector control unit. The semiconductor portion of the element has an inverse 5-25

Figure 5-11 — Typical engine fire warning circuit schematic. temperature coefficient; as the temperature increases, the resistance of the sensing element decreases.

The fire detector control unit continuously monitors the electrical resistance of the fire detector system’s sensing element. The control unit activates a fire warning light (in some units an audible warning is also given) when one of the following conditions exists: 1. The sensing element resistance decreases to the predetermined level (established by the fire alarm setting) due to an increase in temperature. 2. The sensing element resistance decreases at a predetermined rate due to the rate of temperature increase in the sensing element.

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NOTE In the past, flights have aborted and crewmen actually ejected because of a fire warning light illumination caused by a short circuit in the system. Modern aircraft fire warning systems have a short discriminator circuit that differentiates between an overtemp (fire) and a short circuit. If there is a short in this system, the fire warning light will NOT illuminate.

The rest of this section contains a description of a dual jet engine aircraft fire warning sys tem. Both systems operate identically and are similar in operation to fire warning systems found on other Navy aircraft. The left side is discussed in this section. Refer to Figure 5-11 as you read this section. Electrical power from the L FIRE DET circuit breaker supplies the left fire detector and sensing element circuit through pin A of the detector control unit. The left fire-sensing element loop connects to pins L and C of the detector control unit. This completes the sensing circuit through normally closed contacts of the de-energized relay K1. At normal temperatures, the sensing element resistance is high, reverse biasing diode CR1. This allows current through resistor R2 to turn transistor Q1 on. With Q1 on, the base current at Q2 is shut off, turning Q2 off. With Q2 off, the current flows into the base of Q3. This turns Q3 on and Q4 off. Transistors Q3 and Q4 are relay-driving. With Q3 on, relay coil K2-A energizes, opening K2 contacts, de-energizing the warning circuit and turning out the L FIRE warning indicator lights. Normal temperature conditions energize relay coil K2-A. When the temperature rises, the sensing element resistance decreases, shunting the current from resistor R2 through diode CR1, turning transistor Q1 off. This switches transistor Q2 on, transistor Q3 off, and transistor Q4 on. With transistor Q4 on, relay coil K2-B energizes, and relay coil K2-A de-energizes. This transfers (switches) contacts of relay K2, energizing the warning circuit. Then, 28 volts dc powers the warning circuit through pin K of the detector, turning on the L FIRE warning indicator lights. If the temperature drops, the warning circuit de-energizes, causing the fire warning indicator lights to go out. Transistors, Q5 and Q6 make up the short discriminator circuit. The circuit measures the rate of change of sensor resistance. In a fire, the resistance rate of change is slow. In an electrical short condition, the resistance rate drops abruptly. There are two timing circuits— one is made up of resistors R5 and R6 and capacitor C2, and the other of resistors R15 and R16 and capacitor C5. These circuits are preset. This allows a slow change of sensor resistance to let transistor Q1 switch transistor Q2 before transistor Q5 switches transistor Q6. A fast change of sensor resistance allows transistor Q5 to switch transistor Q6 before transistor Q1 can switch transistor Q2. If transistor Q6 switches first, transistor Q2 is unable to switch. This action holds the rest of the circuits in the de-energized (no alarm) condition. If transistor Q2 switches first, the contacts of relay K2-B remove transistor Q6 from the discriminator circuit.

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Figure 5-12 — Container and dual valve assembly. WARNING

The cartridge mentioned in the above paragraph is an explosive device. Use extreme caution when working on or near this device. Refer to Cartridge Actuated Devices (CADS) and Propellant Actuated Devices (PADS) Manual, NAVAIR 11-100-1.1. You need to have an Ammunition and Explosives Handling Qualification and Certification before working on this system. Extinguishing System The fire-extinguishing system on many aircraft provides control for fires within the engines and nacelles. The extinguishing system is an electrically controlled, High Rate Discharge (HRD) system. Normally, you, the AE, will troubleshoot the HRD system electrical circuits only. The extinguishing agent container is a welded steel sphere, 9 inches in diameter and cadmium plated for corrosion prevention (Figure 5-12). Each container has a charge of Halon and is pressurized with nitrogen. Halon is nontoxic and classed as a nonpoisonous; however, it readily vaporizes, is odorless, and can be harmful. You should handle it carefully. Do not let it come in contact with your skin; frostbite or low temperature burns may result. As it leaves the system, vaporization changes the liquid to a gas, displacing the oxygen within the compartment. The lack of oxygen in the compartment will not support combustion or life. Do not enter an area where Halon has been discharged until it is safe to do so. Each container has two valve assemblies for discharging the agent. Each valve assembly contains an explosive, electrically controlled cartridge. When a fire-extinguishing discharge switch actuates, it completes a circuit (Figure 5- 13). The cartridge electrically fires, allowing the slug to rupture the frangible disk in the neck of the container. When the frangible disk is ruptured, nitrogen pressure expels the extinguishing agent.

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Figure 5-13 — Engine fire-extinguishing circuit schematic.

FUEL TRANSFER SYSTEMS The F/A-18 aircraft fuel transfer system is described in this section. The F/A-18 carries fuel internally in four interconnecting fuselage (bladder) tanks and two internal wing (wet) tanks. External fuel is carried in three 315 or 330 gallon tanks. All tanks may be refueled on the ground through a single-point refueling receptacle. Airborne, they can be refueled through the in-flight refueling probe. The internal wing tanks— tank 1 and tank 4—are transfer tanks. The tanks are arranged so internal fuel gravity transfers (at a reduced rate) even if the transfer jet ejectors fail. Regulated engine bleed air pressure is used to transfer fuel from the external tanks and also provides a positive pressure on all internal fuel tanks. Float-type fuel level control valves control fuel level during refueling of all tanks. Fuel level control shutoff valves in tanks 1, 2, 3, and 4 control fuel levels during external fuel transfer. During internal wing transfer, fuel level control shutoff valves control fuel levels in tanks 1 and 4. Fuel level sensors control the fuel level in tanks 2 and 3 (engine feed tanks) during fuel transfer from tanks 1 and 4. All internal and external fuel (except engine feed tanks) can dump overboard through flame arrester protected outlets in each vertical fin. All internal fuel tanks vent through outlets in the vertical fins. The external tanks vent overboard 5-29

through pressure relief valves in the individual external tanks. A fuel quantity indicating system provides fuel quantity indications in pounds. Feed Tanks The internal transfer system design keeps fuel in the feed tanks (tanks 2 and 3) at all engine power settings. Fuel being transferred from tanks 1 and 4 flows to the feed tanks, where the fuel level is maintained by fuel level sensors. Wing Tanks Wing tanks transfer fuel to tanks 1 and 4. They are an integral part of the wing structure. Wing tanks are sealed by filling channels with sealant injected through fittings on the outside of the wings. Transfer Motive Flow The internal fuel transfer system is powered by motive flow pressure, generated by two AMAD motive flow/boost pumps contained in a closed loop circuit. Flow pressure passing through the left and right engine motive flow check valves combines to create transfer motive flow pressure. Transfer motive flow pressure operates the wing transfer ejectors and tanks 1 and 4 transfer jet ejectors. Transfer motive flow pressure also closes the refuel/defuel shutoff valve and defuel valve. Wing Transfer Wing transfer starts when the fuel level drops below the high level pilot valves in tanks 1 and/or 4. This opens the fuel level control shutoff valve, allowing the wing transfer jet ejectors to transfer fuel through the refuel/transfer manifold. As the fuel level in the wings drops below the transfer motive flow pilot valves, the wing transfer motive flow shutoff valves close, stopping transfer from the wings. Refuel/transfer check valves in tanks 1 and 4 keep fuel from entering the refuel line and the feed tanks. As fuel from the hot fuel recirculation system increases wing fuel, the wing motive flow pilot valve and motive flow shutoff valve open, allowing wing transfer. A flapper check valve at each wing ejector inlet prevents transfer from one wing to the other. Transfer motive flow pressure to each wing ejector is controlled by the normally open wing damage shutoff valve in tank 4. If wing damage occurs, the pilot sets the INTR WING switch to INHIBIT on the cockpit EXT LT control panel. This closes the wing damage shutoff valve, preventing loss of fuel through a wing transfer motive flow line and stopping wing transfer. If normal wing transfer does not occur, all wing fuel can be gravity transferred to tank 4 with 5 degrees of roll. A check valve in each gravity transfer line prevents reverse flow. Fuselage Transfer Fuselage transfer (transfer from tanks 1 and 4 to tanks 2 and 3) starts when the fuel level sensors open the transfer shutoff valves in tanks 2 and 3. Fuel flow (transfer) from tanks 1 and 4 transfer jet ejectors enters a fuselage transfer manifold that supplies fuel to a transfer shutoff valve in feed tanks 2 and 3 and to the dump valve. Transfer from the tank 1 or 4 ejector alone is enough to keep both the feed tanks full at maximum engine demand. 5-30

When a transfer tank is empty, the transfer pilot valve and transfer shutoff valve close, preventing transfer motive flow pressure from entering the transfer line. Fuel transfer between fuselage transfer tanks is prevented by check valves in the inlets of each transfer jet ejector. Fuel levels in the feed tanks are maintained by a fuel level sensor and transfer shutoff valve within each feed tank. If transfer from tanks 1 and 4 to the feed tanks fails, fuel gravity transfers through an always open interconnecting line in the bottom of tank 4 and through an orifice in the interconnect valve in tank 1. If motive flow pressure to either engine fuel boost jet ejector or engine fuel turbine boost pump is interrupted, tanks 2 and 3 fuel gravity feed through the ejector to the engine. If the left engine shuts down and/or left motive flow boost pressure is lost, the tank 1 and tank 2 pressure-operated interconnect valves open. This allows fuel to gravity feed from tanks 1 and 2 to tank 3. Reverse flow from tank 3 is prevented by a flapper check valve on tank 3 interconnect valve. If the right engine is shut down and/or right motive flow boost pressure is lost, the tank 3 pressure-operated interconnect valve opens. Tank 4 interconnect line is always open. Fuel gravity feeds from tanks 3 and 4 to tank 2. Reverse flow is prevented by the flapper check valve on tank 2 interconnect valve and tank 3 flapper check valve. On some series aircraft, motive flow pressure to the tank 1 fuel low-level shutoff and pressure operated interconnect valve controls gravity feed to tank 2. The fuel low-level shutoff valve energizes closed when fuel in tank 2 is below 700 to 900 pounds. The closed fuel low-level shutoff valve stops motive flow fuel to the pressure operated interconnect valve, allowing the flapper to swing open. Once the flapper is open, fuel in tank 1 can gravity feed to tank 2. The positions of the tank 2 and tank 3 pressure- operated interconnect valves are tested using the fuel check panel. Center of Gravity (CG) Control System The fuel quantity gauging intermediate device continuously compares the ratio of fuel between tanks 1 and 4. When tank 1 transfers fuel at a faster rate than tank 4, the transfer control valve in tank 1 is energized closed, stopping transfer from tank 1. Tank 1 will not resume transfer until tank 4 transfers (depletes) fuel to within the parameters defined by the intermediate devices. If fuel distribution in tanks 1 and 4 has caused the aircraft CG to be further aft than desired, a CG caution will display on the left digital display indicator. Once tank 1 depletes below 150 pounds of fuel, the intermediate device will stop monitoring tanks 1 and 4 fuel ratios. Tank 1 will then transfer fuel until the transfer pilot valve closes the shutoff valve. OIL TEMPERATURE CONTROL SYSTEM The cooling capacity of the oil cooler system (Figure 5-14) in an aircraft depends on the airflow that passes through the cooler. Airflow is controlled by an oil cooler door actuator, which varies the oil cooler air exit duct. The door actuator is a split-field, reversible dc motor. It includes a magnetic brake for stopping it quickly when it reaches the limits of travel. The control switch has four positions— OPEN, CLOSE, AUTOMATIC, and OFF. In the OPEN or CLOSE position, electrical power goes to the actuator, which opens or closes the oil cooler door. When the switch is in AUTOMATIC, a thermostat controls the actuator. 5-31

Figure 5-14 — Oil temperature control circuit. The thermostatic control unit is in the oil return line. The unit contains two floating contact arms and a central contact arm that actuates by a bimetallic coil immersed in the oil return line. One of the floating contacts is in the door open circuit and the other is in the door closed circuit. The two arms rest on a cam, which a small motor constantly rotates. Thus, the floating contacts are constantly vibrating toward the central contact. When oil temperature rises above normal, the thermostatic element causes the central contact to move toward the door open contact. As the contact vibrates, it intermittently closes the door open circuit. As the actuator intermittently energizes, the door slowly opens. When the oil temperature returns to normal, the central contact moves back to a neutral position. When oil temperature falls below normal, the central contacts move in the opposite d irection, closing the door. To prevent excessive hunting of the system, a tolerance is maintained by an adjustment of the cam on the floating contact. When the oil temperature raises high above the normal value, the central contact lifts the floating contact clear of the cam, completing a continuous circuit. The door then moves to the full open position where a limit switch in the actuator breaks the circuit. Figure 5-15 shows another type of engine oil temperature regulator. This regulator has a mercury-filled thermostat, and relays automatically control the position of the engine oil cooler doors. When the engine temperature is low requiring more heat, the two relays energize, allowing the oil cooler door to close. As the temperature increases, the thermostat completes a path to ground, bypassing the relay coils and de-energizing them. Power then goes through the contact of one of the relays, opening the actuator coil and causing the oil cooler door to open and reduce engine temperature. 5-32

Figure 5-15 — Automatic oil temperature control circuit.

V ARIABLE EXHAUST NOZZLE CONTROL SYSTEM As you read this section, refer to the simplified schematic diagram of a typical variable exhaust nozzle (VEN) control system (Figure 5-16). This is the control system for the F/A-18 Legacy Hornet aircraft, and it is a converging-diverging nozzle system. When operating, this system varies the exhaust escape area size to obtain desired thrust, while maintaining safe operating conditions throughout the engine. The VEN control system consists of electrical, hydraulic, and mechanical components that position the VEN while maintaining exhaust gas temperature (EGT).

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Figure 5-16 — Variable exhaust nozzle system schematic.

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Major Components The VEN control system has nine major components. Locate each component in Figure 5-16 as you read about it in the text. 1. VEN power unit - The VEN power unit provides hydraulic power to actuators for positioning the VEN area. 2. Main Fuel Control (MFC) - The MFC provides a regulated flow of fuel to the fuel nozzles. 3. Electrical Control Assembly (ECA) - The ECA computes, schedules, and controls engine operation. 4. Fan speed transmitters - The transmitters are eddy-current sensors mounted in line with the second stage fan blades. A permanent magnet, rotating at the RPM of the fan blades, induces a voltage into a coil indicative of the fan speed. 5. Afterburner Control (ABC) - The ABC schedules fuel to the afterburner pilot and main spray bars. 6. Afterburner (AB) flame sensor - This sensor provides an electrical signal to the ECA. This signal must coincide with the AB no-light/light condition to start the afterburner. 7. Thermocouple harness - This device senses the Exhaust Gas Temperature (EGT). 8. VEN position transmitter - The transmitter provides feedback to the ECA to ensure the VEN is in the correct position. It also gives feedback to the Engine Monitor Indicator (EMI) to indicate percent of nozzle position. 9. VEN actuators - These actuators hydraulically operate to position the VEN. Operating Principles The VEN schedule is in response to movement of the throttle. Throttle setting repositions the Power Lever Angle (PLA) cam in the MFC, providing a Linear Variable Differential Transformer (LVDT) signal to the ECA. The ECA biases the VEN area schedule according to inputs from the following sources:  Fan inlet temperature transmitter  Air Data Computer (ADC) ambient pressure  Fan/low-pressure turbine speed  Compressor/high-pressure turbine speed  Main fuel control metering valve  LVDT  Afterburner control metering valve  Afterburner flame sensor  Afterburner permission signal  Thermocouples  VEN area position transmitter  AB permission switch

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The VEN area closes when the engine shuts down and the throttle is in the off position. As the throttle is moved to the idle position, the VEN area rapidly moves to an almost full open position. This aids engine starting and lowers engine thrust, allowing higher idle speeds and reducing engine acceleration time. As the throttle advances past the idl e position, the VEN area schedule closes the VEN area, increasing thrust. As the VEN area decreases, EGT increases. The ECA adjusts the VEN area for varying atmospheric conditions. When the ECA receives an ambient pressure signal from the air data computer, it means the aircraft is at a pressure altitude of 9,000 feet or greater. The ECA now increases low-pressure turbine discharge temperature and fan/low-pressure turbine speed limits, providing more thrust. Below 4,000 feet pressure altitude, the schedules trim back to reduce fuel consumption and to increase hot section life. This trim signal varies in size between 4,000 and 9,000 feet. To ensure correct positioning of the VEN area, the VEN position transmitter LVDT provides feedback to the ECA. Any error between the actual VEN area position and its required position goes to the bias signal to readjust the VEN area to its correct position. As the throttle advances into military (MIL) power (100 percent) and afterburner (AB) ranges, the VEN area maintains the low-pressure turbine discharge temperature within established limits. Throttle position establishes this limit. This limit is adjustable for ambient pressure for fan inlet temperature and for actual low-pressure turbine discharge temperature values from the thermocouple harness. As the throttle enters AB range, the VEN area reopens slightly above the normal throttle setting, and low-pressure turbine discharge temperature resets to a lower value. These conditions are held until the flame sensor signals the ECA of an AB light-off. The ECA then releases its hold on the VEN area and re-establishes actual low-pressure turbine discharge temperature values. The VEN area will adjust to maintain actual low-pressure turbine discharge temperature limits. The VEN power unit supplies hydraulic power for positioning the VEN. The power unit activates on an electrical signal from the ECA to the power unit torque motor. The torque motor drives a servo, which supplies high oil pressure to the synchronized actuators to open or close the VEN. PROPELLER SYNCHROPHASING SYSTEM The propeller synchrophaser system discussed here is common to the P-3 and C-130 aircraft and is similar to the E-2 aircraft. In this section, you will learn about the electrical operation of controlling and synchrophasing the hydromatic propellers of multiengine aircraft. Propeller Governor A propeller governor is a control device that controls engine speed by varying the pitch of the propeller. Increasing the propeller pitch adds load on the engine and increases propeller thrust. This load reduces engine speed. Conversely, decreasing the propeller pitch reduces engine load, which increases engine speed. Therefore, engine speed is a function of propeller pitch. Furthermore, if the propeller governor setting remains unchanged, any variation of power produced by the engine translates into a corresponding variation of propeller thrust. The system does this by varying the propeller pitch while engine speed remains constant. The best engine efficiency is when 5-36

engine speed is constant; therefore, because it controls engine speed, the propeller governor achieves engine efficiency. The pitch of hydromatic propeller blades changes by porting hydraulic fluid onto the propeller piston in the propeller dome. The action on the piston transmits through a geared cam mechanism that rotates the propeller blades to the pitch desired. The governor is the constant speed control device used with the hydromatic propeller. The output oil of the pump goes to either the inboard or the outboard side of the propeller piston. There are two separate ranges of propeller operation, the flight range and the ground operating range. The flight range includes the takeoff roll after the power levers advance forward for takeoff. For the ground operating range, power levers return aft of the flight- idle detent. In the ground operating range (taxi range), power lever position determines propeller blade angle. A hydro-mechanical system, with linkage to the power lever, meters oil pressure to either the increase or decrease side of the propeller dome. As the power lever moves forward toward FLIGHT IDLE, a simultaneous increase in blade angle and fuel flow occurs, providing increased power, As the power lever moves aft from FLIGHT IDLE, blade angle decreases and fuel flow decreases, reducing power. Fuel flow begins to increase when the blade angle decreases to the point that the propeller is delivering negative thrust. Reverse power continues to increase until the power levers reach the full aft position. During operation in the ground operating range, there is no electronic governing. In the flight range of operation, the power lever is forward of the flight-idle detent. In this range, a flyweight governor, driven by propeller rotation, mechanically controls propeller speed. In normal operation, the pitch-change oil goes through the feather valve to either the increase or decrease pitch portion of the propeller. Synchrophasing The synchrophaser has different functions, depending upon the mode of governing selected by the flight crew. The synchrophaser does not function in the mechanical governing mode, but the mechanical governor controls the blade pitch and so propeller RPM. In a normal governing mode, the synchrophaser helps the mechanical governor by limiting engine transient speed changes, or to changes in flight conditions affecting propeller speed. In a synchrophasing governing mode, the synchrophaser helps the mechanical governor by maintaining all propellers at the same RPM. It does this by maintaining a preset phase relationship between the master propeller number 1 blade and the number 1 blades of the slave propellers. This serves to reduce noise and vibration in the aircraft. In the synchrophasing governing mode, the synchrophaser also provides the limiting of transient speed changes as it does in normal governing mode. The synchrophaser consists of four main components, a pulse generator, a phase and trim control, a speed-bias servo assembly, and the synchrophaser. 1. Pulse generator - The pulse generator provides the information needed by the synchrophaser system to produce speed and phase control of the aircraft propellers. Each propeller has a pulse generator that consists of a permanent magnet on the propeller spinner and a stationary coil installation in the governor control. Each time the permanent magnet passes by the coil it generates a pulse. Each revolution of the propeller generates one pulse. 5-37

2. Phase and trim control - The phase and trim control functions as a means of setting phase relationships between master and slave propellers. It also trims the master engine. The phase and trim control consists of seven potentiometers that receive a fixed dc voltage from the synchrophaser. The wiper of the master trim potentiometer supplies a voltage through the master select switch to the synchrophaser to trim master engine speed. The other six wipers connect to relay contacts that separate the wipers into two groups of three per group. One group corresponds to engines 1, 3, and 4 when engine 2 is master. The other group corresponds to engines 1, 2, and 4 when engine 3 is master. These wipers supply bias voltages to the phase correction circuits of the synchrophaser to set propeller phase angles of other than 0 degree. 3. Speed-bias servo assembly - The speed-bias servo assembly functions as a means of translating synchrophaser electrical signals into a mechanical bias on the mechanical governor speeder spring. The synchrophaser supplies the servomotor with a reference voltage that is 90 degrees out of phase with the aircraft 400-Hz source. The synchrophaser also supplies a control voltage that is either in phase or 180 degrees out of phase with the aircraft 400-Hz source. Therefore, the in-phase control voltage lags the reference voltage by 90 degrees. This lag results in counterclockwise motor rotation when viewed from the output gear of the electric brake. The 180-degree out-of-phase control voltage leads the reference voltage and causes clockwise rotation. The amplitude of the control voltage determines motor speed and torque output. The motor drives a reduction gear train, which, in turn, drives a potentiometer wiper and the electric brake. The potentiometer receives a fixed dc supply from the synchrophaser across its resistive element. When the motor rotates, the wiper transmits a corresponding feedback voltage to signal winding number 2 of the magnetic modulator. The electric brake has clutch-controlled input and output shafts. The output shaft drives a lever, which biases the speeder spring in the propeller governor. Energizing the clutch decouples the two shafts, locking the output shaft and leaving the input shaft free to turn. 4. Synchrophaser - The synchrophaser has four channels, which correspond to the aircraft’s four engines. Figure 5-17 is a schematic of the synchrophaser. For explanation purposes, only two channels (one slave channel and one master channel) are shown. Each channel has a push-pull power amplifier feeding the control winding of its corresponding servomotor in the speed-bias servo assembly. Magnetic modulators using dc control current furnish a phase and amplitude controlled ac signal to the push-pull amplifier input. The synchrophaser changes all signal inputs to dc voltages proportional to the error before they are applied to the modulator. The modulators are the signal summing devices for the two operational modes of the synchrophaser.

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Figure 5-17 — Synchrophaser control schematic diagram.

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The magnetic modulators function on a core saturation basis. Each modulator consists of a dc bias winding, a 400-Hz excitation winding, and two control windings (signal winding number 1 and signal winding number 2). With no signals applied elsewhere, the 400-Hz excitation voltage appears as a 400-Hz output of negligible amplitude due to the bias winding current. Any current in either or both signal windings will change the output. The size of the signal windings current controls the amplitude of the output; the current direction controls the phase of the output. Thus, current from pin 10 to 9 in winding number 2 and current from pin 8 to 7 in winding number 1 of any modulator produces a voltage 180 degrees out of phase from the excitation voltage. Current in the opposite direction in the signal windings produces an in-phase voltage. Simultaneous currents flowing in opposite directions in the two signal windings produce a signal that is the algebraic sum of the two signals. Then, the modulator produces a 400-Hz signal, which is either in phase or 180 degrees out of phase with the excitation voltage. This signal is amplified and fed to the servomotor control winding. The 400-Hz voltage in the reference winding of the servomotor goes through a series capacitor, giving the voltage a 90-degree phase shift from the aircraft power source. Appropriate signals to the modulators cause clockwise or counterclockwise rotation of the motor because of phase difference in the speed bias motor windings. The use of the two signal windings in the modulators, along with appropriate relay switching, permits the two modes of synchrophaser operation. Operational Modes The normal governing mode provides improved engine response to transient RPM changes. In this mode, the synchrophaser receives signals from the power lever anticipation potentiometers and the engine tachometer generators. Signals from these result in a temporary resetting of the mechanical propeller governor. This resetting adjusts for power lever changes and engine speed changes, thus limiting engine overspeeds or underspeeds. The synchrophasing governing mode synchronizes engine speeds, regulates propeller phase angles, and maintains the limiting features of the normal governing mode. In the synchrophasing governing mode, one engine (2 or 3) is the master. The master engine operates in normal governing mode, while the other three engines (slaves) follow changes in speed or phase of the master within preset limits. Normal Governing Mode In normal governing mode, the propeller governor switch is in the NORMAL position and the power lever switch closes, providing reference voltages to the servomotors. The synchrophaser master switch is OFF and the PROP RESYNCH switch is in NORMAL. All relays are de-energized, resulting in the speed and phase error circuits being grounded. Each phase- and speed-error signal side of every magnetic modulator signal winding number 2 (pin 10) ends on a dummy load (Figure 5-18) within the synchrophaser. The other side (pin 9) connects to the feedback circuit in the speed-bias servo assembly. The controlling signals go to signal winding number 1 (pin 8) of each modulator. All channels function identically while in the normal governing mode. THROTTLE LEVER ANTICIPATION — Any power lever movement (Figure 5-18) causes a change in dc voltage at the anticipation potentiometer wiper, which serves as 5-40

Figure 5-18 — Power lever anticipation and speed derivative circuits in normal governing mode. a voltage divider for the RC circuit. The charging voltage for the capacitor is directly proportional to the position of the power lever.

The change in charge on the capacitor is directly proportional to the rate at which the power lever moves. If the power lever movement is to decrease engine power, the capacitor charges up to a more positive voltage value. This results in a current from pin 7 to pin 8 in signal winding number 1 of the magnetic modulator. A lagging voltage surge appears in the servomotor control winding, causing counterclockwise rotation. This rotation resets the mechanical governor towards decrease pitch to compensate for the reduced power setting. As the servomotor rotates, the feedback potentiometer begins canceling the error signal by causing a current in signal winding number 2. The magnetic field of this current is in opposition to the magnetic field of the signal current in 5-41

winding number 1. This stops the servomotor. As the anticipator capacitor continues to charge to its new peak value, the current in signal winding number 1 decays to zero. The feedback potentiometer is still applying voltage to signal winding number 2. This results in a leading voltage to the servomotor control winding that returns the motor to its original position. This position corresponds to a zero-volt feedback potentiometer position. In retarding the throttle lever very rapidly, the peak voltage will overcome the reverse bias on diode CR620. This will limit the signal value to prevent overcompensation toward a flat blade pitch. For an increase in engine power, the capacitor discharges. This causes a current in the opposite direction in signal winding number 1, which results in a temporary resetting toward increased pitch. The amount of reset in either case depends on the rate at which the lever moves. Mechanical stops in the speed-bias servo assembly limit speed resets to plus 10 and minus 10 percent, regardless of the applied signal. Furthermore, stops in the propeller control valve housing linkage reduce the limits to plus 6 and minus 4 percent. LIMITING ENGINE TRANSIENT SPEED CHANGES — The speed-derivative circuit in the synchrophaser (Figure 5-18) senses changes in engine RPM and produces output signals, which dampen the engine RPM changes. The speed-derivative circuit does this by translating the frequency changes received from one phase of the tachometer generator into signal voltages. The magnitudes of the signal voltages vary at the rate the tachometer generator frequency changes. The signal voltage goes to signal winding number 1 of the magnetic modulator, where it is summed and sent to the push-pull amplifier. After amplification, the signal goes to the servomotor control winding. The servomotor adjusts the speeder spring tension, which begins a change in propeller pitch, thus dampening the change in engine RPM. The action of the speed-derivative circuit is further described as follows: The voltage produced on the collector of transistor Q603 is proportional to the output frequency of the tachometer generator. When engine RPM is constant, the voltage on the collector is constant and capacitor C621 charges through resistor R634 and signal winding number 1 of the magnetic modulator. Current in signal winding number 1 decays to zero as the charge on capacitor C621 reaches the potential on the collector of transistor Q603. When engine RPMs change, a change in the collector voltage of transistor Q603, proportional to the change in tachometer generator frequency, occurs. This causes capacitor C621 to change its charge at the rate in which the tachometer generator frequency is changing. This produces a current in signal winding number one. The current size varies at the rate at which the engine is varying off-speed. The amplified signal in the servo-bias assembly control winding drives the servomotor in a direction to dampen the drift in engine RPM. Speed-error signals in signal winding number 1 from the speed-derivative circuit cancel in the same manner as anticipation signals from the anticipation circuit cancel. The speed-derivative and power lever anticipation circuits are much more sensitive to engine RPM changes than the mechanical governor flyweight speeder spring. The governing action of the flyweight and speeder spring improves the mechanical governor’s response to changes in power lever settings and engine RPM. Synchrophaser Mode In adding synchrophasing to normal governing mode, the master switch selects either engine 2 or 3 as the master engine (Figure 5-17). With master engine selection, relays energize, removing dummy loads from signal winding number 2 of all magnetic modulators, except the master channel modulator. Also, the outputs of the speed-error and phase-error circuits of each synchrophaser channel (except the master) are taken 5-42

NOTE Since the propellers are four-bladed, the relative blade position between the master and slave propellers is exactly the same when the slave propellers differ from the master by one-half revolution. Therefore, consider a 180-degree phase difference as an on-phase condition in pulse comparisons. from ground and connected to signal winding number 2 of their respective modulators. While the slave engines are in synchrophasing mode, the master engine remains in normal governing mode. Essentially, pulses from the master engine form into saw-tooth waves and are compared to pulses from each of the slave engines in the slave channel sampling circuits. If slave pulses are not in phase with the master pulse (saw-tooth), error detection occurs in the respective synchrophaser channel. The errors then go to signal winding number 2 of the channel magnetic modulator. Here, they are summed with any error signals that exist in signal winding number 1. The result of the error signals is amplified and fed to the control winding of the respective speed-bias servomotor. This alters the tension of the slave governor speeder spring, correcting for engine speed differences and propeller blade angle errors. Figure 5-19 shows one channel of the synchrophasing circuit. The following paragraphs describe phase and speed error sensing. SAW-TOOTH FORMER—The pulse from the master pulse generator is transformer- coupled to the saw-tooth former. The slave pulse generators are transformer-coupled to the channel sampler circuits. Figure 5-20, views A and B, shows the master pulse and the resultant saw-tooth formed in the saw-tooth former. SAMPLING CIRCUITS—Pulses from the three slave engines couple to the grids of the sampling circuit tubes, while the saw-tooth voltage goes to the plate of one tube and the cathode of the other tube in all sampling circuits. The positive-going portion of the slave pulse places the tubes in a conductive state. Sampling is the same in all channels. Refer to Figure 5-19. If the saw-tooth is at zero potential when the slave pulse occurs, neither tube conducts, so the phase difference and speed error circuits receive no signal. With the saw-tooth in the positive region, tube V205 conducts. The corresponding voltage changes go to the grid of phase difference tube V206A. You can see the nature of the sampling action by referring to Figure 5-20. The time interval between master pulses is the time interval for 360 degrees of propeller rotation (Figure 5-20, view A). The time interval of the saw-tooth, which the master pulse generates, is 360 degrees. The half-time interval, or 180-degree position, is the saw- tooth zero potential point (Figure 5-20, view C). When the slave pulse occurs at the zero point, the propellers are on phase with a 180-degree phase difference between them. All references to slave pulses are with respect to the 360-degree interval. The point of occurrence of the slave pulse determines the signal size in the sampler circuit (Figure 5- 20, view D). This signal represents the phase difference between propellers.

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Figure 5-19 — Synchrophaser schematic diagram, synchrophaser mode (slave channel).

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Figure 5-20 — Master and slave pulse comparisons.

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Figure 5-21 — Off-speed pulse comparison: (A) slave underspeed; (B) slave overspeed. PHASE DIFFERENCE CIRCUITS —Look at Figure 5-19. The phase difference circuits receive the voltages the sampling circuits generate at the grids of the respective tubes. With a 180-degree phase difference signal at the grid of V206A, the voltage at the wiper of potentiometer RP251 adjusts to null or zero volts. This voltage changes proportionally with the grid signal from the sampling circuit, and hence represents the phase difference between propellers. The following paragraphs discuss the effect of this voltage on synchrophaser output. OFF -SPEED CIRCUITS —When the propeller goes off-speed, the sampling circuit senses the condition as a sharp voltage change. When the slave propeller is on-speed, the slave pulses occur at the same time interval in each successive saw-tooth cycle. When an underspeed or overspeed condition develops, the slave pulse occurs at a different position for each saw-tooth cycle. At one point, the slave pulse falls up or down the saw-tooth (Figure 5-21). Think of this as a phase error that occurs too rapidly for the phase-error circuit to compensate. During an underspeed condition, the voltage going to the phase difference tube suddenly changes from positive to negative (Figure 5-21).

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For an overspeed condition, the voltage suddenly changes from a negative to a positive. As you read this paragraph, refer to Figure 5-19. The action of the off-speed circuit is as follows. The underspeed voltage change, registered in the plate circuit of the phase difference tube, couples to the off-speed circuit through capacitors C253 and C254. The voltage change in the plate circuit is positive and the reverse bias on diode CR251 is reinforced. However, the reverse bias on diode CR252 is momentarily overcome, allowing capacitor C255 to charge positively. Capacitor C255 discharges slowly through its parallel resistive network and maintains a grid bias on the off-speed tube V206B. For an overspeed, diode CR251 conducts and charges capacitor C255 negatively. Successive sharp voltage changes add to the charge on capacitor C255 until correction of the off-speed condition occurs. Like the phase difference circuit, the off-speed circuit has a zero-volt adjustment potentiometer in the cathode circuit. The pulsating voltage generated at the wiper by the changing grid bias on the tube is the speed-error voltage. The effect of this voltage is discussed later. SIGNAL SUMMING — As you can see by looking at Figures 5-19 and 5-22, the phase- difference and speed-error voltages couple to the signal summing point. Also connected to this point are the limiting circuit, the phase-control potentiometer, and magnetic modulator signal winding number 2. Any difference in potential between the signal summing point and the feedback potentiometer causes a current in the modulator winding. This results in motor movement, which, in turn, causes the feedback potentiometer to null whatever potential is at the summing point.

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Figure 5-22 — Simplified schematic of signal summing.

PHASE-ERROR CORRECTION —When the slave propellers are not on-phase, phase- difference voltage acts through the averaging circuit to the summing point causing speed-bias servomotor rotation to correct the off-phase condition. This action occurs provided the phase control potentiometer (Figure 5-22) is set for zero volts. In this case, the phase difference voltage actually represents phase error. However, it is desirable to have a slave propeller maintain a specific angle of lead or lag to the master propeller. The phase control potentiometer then adjusts to cancel the phase-difference voltage at the summing point. For example, to maintain a slave lead of 10 degrees from the 180- degree position of the master pulse, the phase control potentiometer adjusts to cancel the phase-difference voltage. The phase-difference voltage is generated when the slave propeller leads the master propeller by 10 degrees. Thus, when the slave propeller leads by 10 degrees, net potential at the summing point is zero, and the speed-bias servomotor does not move. When the slave propeller changes from the 10-degree lead condition, the phase-difference voltage changes. The net potential at the summing point is the difference between the phase control potentiometer setting and the phase- difference voltage. This is how phase-error voltage and its size depend on the slave propeller’s degree of lead or lag from the 10-degree lead condition. The manual phase control is adjustable to allow a slave lead or lag of 45 degrees. 5-48

SPEED-ERROR CORRECTION —During the phase-error correction, Zener diodes VR405 and VR408 isolate the speed-error potentiometer from the signal summing point. When an off-speed occurs, these diodes conduct and connect the speed-error circuit to the summing point. For large speed errors, the speed-error signal is a pulsating dc of very low frequency. At the same time, the phase-error signal is a rapid dc step voltage consisting of sharp potential changes, which trigger the speed-error circuit. The dc step voltage changes developed in the cathode circuit of V206A (Figure 5-19) are averaged and effectively blocked by resistor R268 and capacitor C251. The sharp potential changes passing resistor R268 and capacitor C251 leak to ground by resistor R272 and capacitor C256, leaving the speed-error signal in control. When the engine reaches on- speed condition, but is still out of phase, the speed-error signal drops off until a point where the Zener diodes stops conducting and the phase-error signal assumes control. TWO-PERCENT LIMITING — The potential at the summing point has a limit of plus or minus 5 volts. The limiting circuit R442, R443, R444, R445, CR409, and CR410 as shown in Figure 5-22 accomplishes this by clipping any signal outside of the range limit. In the off-speed circuit, this limited signal provides only enough output to drive the speed-bias servomotor to correct a two-percent speed change. This occurs because the motor movement results in a feedback voltage that cancels the speed-error signal. The slave propeller cannot follow large speed changes of the master engine. This prevents a slave from following an overspeeding or underspeeding master. Resynchrophasing The need for re-synchrophasing arises from the nature in which phase angle correction circuits operate. As phase errors occur, the servomotor rotates to correct the error. At the same time, the feedback potentiometer moves and cancels a portion of the error signal. As the phase-error correction occurs, the phase-error signal decreases until it matches the feedback signal. When this occurs, the potential at the summing point is zero and the motor stops moving, leaving a portion of the error uncorrected. This is insignificant for errors that occur about a set point, leading and lagging errors. However, for errors that continually occur in one direction, the error accumulates, and can become large enough to reduce system efficiency. To overcome this, a procedure provides a lock on the mechanical governor bias, while the servomotor re-centers at a zero-volt feedback position; this is re-synchrophasing. As you read this paragraph, refer to Figure 5-18. While in synchrophasing mode, the prop re-synchrophase switch is placed in the RESYNCH position. This energizes the electric clutch-brakes on the slave engine speed-bias servo assemblies. The brakes lock the output shafts and the clutches decouple the input and output shafts. This action leaves the input shafts free to turn. Locking the output shaft retains whatever mechanical bias is present on the mechanical governor. At the same time, relay K503 energizes. This opens the signal winding number 1 circuits on the magnetic modulators. The master channel remains in normal governing mode through a parallel ground provided by the master relay to signal winding number 1 of the master channel modulator. After a time delay sufficient to assure brake and clutch actuation before allowing the recentering action, relays de-energize, removing the phase- and speed- error circuits from the slave channel magnetic modulators and connecting the dummy loads. The only input to the slave channel magnetic modulators is now the feedback potentiometer acting into the dummy load through signal winding number 2. The generated signal then drives the motors until the feedback voltage is zero. Upon releasing the prop resynchrophase switch, all circuits return to the synchrophasing mode. Correction of the accumulated phase error can now occur. 5-49

PROPELLER CONTROL SYSTEM The propeller control system is similar to systems already discussed in this chapter. It is an integral part of the propeller synchrophaser system and the propeller governing system. The following discussion concerns propeller pitchlock, the negative torque system, feathering, unfeathering, and autofeathering operations. You should refer to Figure 5-23 throughout this discussion of the components in and operation of the propeller control system.

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Figure 5-23 — Propeller control circuit.

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Propeller The four-bladed propeller converts engine shaft horsepower to thrust. The propeller consists of two principal sections, the rotating section and the nonrotating section. The rotating section consists of the blades, hub, spinner, and the dome that houses the pitch changing mechanism. The non-rotating section contains the pressure and scavenge oil pumps, the governor control mechanism, and the spinner after body. It is a constant- speed, full-feathering, reversing propeller, having the added features of pitchlock and a combination synchronizing and synchrophasing system. Low-Pitch Stop Assembly A mechanical stop assembly in the propeller dome limits low pitch blade angle to maintain a minimum blade angle. When the power levers position is below 28 degrees, a cam-operated backup valve in the propeller housing collapses the low-pitch stop levers. The blades move toward the reverse position as directed by the power lever beta schedule. Beta Follow-up System The beta follow-up system provides a variable hydraulic low-pitch stop. At the FLIGHT IDLE power lever position, the beta follow-up stop is 10-degree blade angle; the mechanical low-pitch stop is 13 degrees. The purpose of the beta follow-up stop is to provide a secondary low-pitch stop. As the power lever moves toward the TAKEOFF position and the blade angle increases, the mechanical low-pitch stop remains at 13 degrees. The beta follow-up control programs the hydraulic stop in relation to power lever position to a maximum of 22.5 degrees of blade angle. If a sudden engine failure occurs and the negative-torque system fails to operate, beta follow-up prevents excessive reduction in blade angle and associated violent yawing. Negative Torque System (NTS) The NTS protects the aircraft from excessive drag by limiting the negative torque from the propeller to a predetermined value range of -150 to -500 shaft horsepower. During a negative torque condition, NTS provides a mechanical signal that overrides the propeller governing action to increase blade angle. When the propeller reaches a position where it is no longer developing negative torque, the propeller governor regains control of the propeller. If the negative torque condition persists, a cycling action continues from the mechanical signal to the propeller governor, and vice-versa, until some corrective action is taken. The NTS INOP warning light illuminates when the 45-degree air start blade angle circuit energizes. The air start blade angle system limits negative horsepower during in-flight restart operations if NTS fails. The blades, upon reaching 45 degrees, energize the air start blade limit circuit, which drives the feather valve to the feather position. The blades moving toward the feather position open the circuit, de-energizing the feather solenoid valve. The result of this action is a cycling of the blades around the air start blade angle switch. For ground unfeather operation, bypass the air start blade limit switch by actuating the feather pump pressure cutout override switch while unfeathering the propeller.

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Pitchlock A hydraulic, speed-sensitive pitchlock mechanism is in the propeller assembly to prevent overspeeding. It is completely automatic and prevents further decrease in blade angle when engaged. Automatic engagement occurs in two ways: by loss of control oil pressure and when RPM exceeds 103.5 percent. The pitchlock governor, sensing the overspeed, allows the pitchlock teeth to engage, which prevents further decrease in blade angle. Due to teeth design, the blade angle can increase if the normal governing control is restored. BLOCKOUT RANGES —There are two blade angle ranges where pitchlock does not operate. One range is from plus 17 degrees to minus 14 degrees. This allows for RPM surges as blade angles reduce during approach and landing. The second is blade angles between 57 degrees and 86 degrees (full feather). This is necessary to decrease blade angle for air starting. Blade angles will be less than 57 degrees at 405 knots or limit Mach speed. PITCHLOCK RESET —A reset system will reset pitchlock up to 109 percent RPM at blade angles above 10 degrees with the power lever below 28-degree coordinator position. This 109 percent RPM setting of the pitchlock is momentary. It is necessary when the power lever moves very rapidly into the beta range and blade angle reduction is not quick enough. Pitchlock reset comes into use during an aborted takeoff or landing. In this condition the increase in fuel scheduled will cause engine speed to increase above 103 percent RPM. FUEL GOVERNOR AND PROPELLER PITCHLOCK TEST SWITCH — Four two- position (TEST-NORMAL) switches, one for each propeller, are on the engine check panel. Upon placing a switch to TEST, the propeller speed-bias servomotor receives a continuous overriding signal. The signal drives the mechanism full travel toward decrease pitch. This resets the propeller governor to a speed of 106 percent RPM, permitting a ground check of the pitchlock and fuel governor functions. Propeller Feathering Feathering aligns the propeller to the airstream to minimize drag during engine shutdown conditions. Feathering is started by pulling the engine emergency shutdown handle, autofeathering, or depressing the feather switch button. The feather valve ports hydraulic fluid to increase pitch on the propeller blades. This action bypasses all other control functions. NORMAL FEATHERING — To accomplish normal feathering, pull the engine emergency shutdown handle. Pulling the handle mechanically positions the feather valve and electrically energizes the feather button solenoid. Also, current goes to the auxiliary pump and to the feather valve solenoid, which hydraulically positions the feather valve to feather the propeller. When the propeller fully feathers, oil pressure buildup operates a pressure cutout switch, causing the auxiliary pump and feathering solenoid to de-energize. The feathering button light then goes out. Feather Pump Pressure Cutout Override Switch —There is a push-button switch next to each feather button. Actuating the override switch bypasses the feather pump pressure cutout switch. This permits continued operation of the feather pump if the feather pump operation finishes before the propeller reach the full feather. Feathering Switches (Buttons) — Four guarded feather switches (buttons), one for each engine, provide an alternate method for feathering the propellers. Pressing a button to FEATHER cuts off fuel electrically, and energizes the feather solenoid and the 5-53

feather pump motor to feather the propeller. The propeller unfeathers when the button goes to the unfeather position. The center position is for normal propeller operation. A light in the feather button illuminates when the circuit to the feather pump is energized. AUTOMATIC FEATHERING — The automatic feathering system performs its function by energizing the feathering button holding coil (pulling in the feathering button). This occurs when the autofeather arming switch is in the ARMED position and the engine loses power, which results in a large decrease in propeller thrust. This system is for use during takeoff only, and it functions above 60°±2° coordinator setting. Only one propeller will feather automatically. Autofeather Arming Switch —There is an autofeather ARMED-OFF switch on the autofeather and RPM control panel. It provides the control for the autofeather system. During normal operations and when the switch is in the ARMED position, all power plants have autofeather protection. With the switch in the ARMED position and electrical power to the power lever quadrant, autofeather arming switches close and the autofeather armed lights illuminate. Both the arming switch and power lever quadrant switch (activated at 60 degrees) must close before the thrust-sensitive signal device can cause propeller autofeathering. Autofeather System Indicators — Four green indicator lamps, one for each propeller, illuminate to show the arming of each individual propeller autofeathering circuit. The lights are on the pilot’s overhead control panel. Also, should a propeller autofeather, its light will remain on and the others will go out. UNFEATHERING ON THE GROUND — The propeller is unfeathered by holding the feathering button in the unfeather position. The air start switch (in the propeller) limits the blade angle decrease to 45 degrees. For further blade angle decrease, the pressure cutout override button is held in. Propeller Control Operation In this section, you will read about a typical propeller control operation under the following conditions:  The engines are running, and  The engines are operating in the taxi range (power levers between 0 degree and 34 degrees on the coordinator). Each power lever controls the propeller blade pitch and engine fuel flow through hydro- mechanical linkages. If a power lever is below 28 degrees and the blade pitch angle is above 10 degrees, the pitchlock reset valve solenoid (Figure 5-23) activates, resetting pitchlock up to 109 percent RPM. As the power levers advance above 34 degrees into the flight-idle range, the flyweight propeller governor assumes propeller pitch control. This governor increases or decreases blade pitch to maintain 100 percent RPM by directing fluid through the feather valve. The temperature datum system controls proper engine temperature operation. When ready for takeoff, the pilot places the AUTOFEATHER switch in the ARMED position. As the power levers advance beyond 60 degrees, the number 3 power lever switches close to the HIGH POWER position. If a propeller’s thrust drops below 500 pounds, the thrust sensitive switch closes. This energizes the auto feather relay. Power goes through the auto feather relay contacts to energize the feather switch pull-in and hold-in solenoid. The auto feather relay loses its power from the feather switch, but the 5-54

feather switch remains energized through a holding circuit. Power runs from the energized feather switch through the contacts of the feather cutout relay to the feather pump power relay to the feather pump motor. Power also goes to the feather valve solenoid through the feather cutout relay. The feather valve solenoid moves the feather valve, which ports oil to increase pitch, causing the propeller to feather. The fail-safe- type fuel control relay activates by power from the feather switch, causing the engine fuel shutoff valve to close. As the propeller feathers, the beta cam switch closes when the blade angle goes past 74 degrees. The feather motor cutout switch closes when the propeller reaches full feather and the oil pressure increases. This completes a circuit that causes the feather cutout relay to energize, de-energizing the feather valve solenoid. Besides cutting out the feather pump power relay, the energized feather cutout relay maintains its own holding circuit. This completes the auto feathering cycle. When shutting down the engine with the EMERGENCY ENGINE SHUTDOWN handle, power from the emergency engine shutdown relays is directed to the feather switch pull- in and hold-in solenoid. Each emergency shutdown handle has two switches and two relays to make sure the electrical systems energize. The engine is shut down and feathered as described above. Also, pressing in the feather switch causes the feather cycle operation to take place. When the flight crew restarts an engine during flight, the feather switch is pulled out to the UNFEATHER position and held there. The feather cutout relay is de-energized, and power goes to the feather pump power relay. The feather pump motor pressure builds up and oil flows to the decrease pitch side of the propeller dome, causing the propeller to start unfeathering. During restarts the feather valve solenoid energizes during NTS inoperative conditions only. When blade angle decreases to less than 45 degrees, the beta cam air start switch closes, completing the path for current through the air start control relay. The air start control relay energizes the feather valve solenoid. This causes the propeller blades to increase toward the feather position. The beta cam air start switch opens at 48-degree blade angle, causing the air start control relay and feather solenoid valve to de- energize. This causes cycling of blade angle around the air start beta cam switch (45 degrees) and causes the NTS INOP warning light to flash. The light blinks because of the blade angle cycling action. At this indication, the flight crew pulls the emergency shutdown handle to abort the restart. APPROACH POWER COMPENSATOR SYSTEM The Approach Power Compensator (APC) system (Figure 5-24) automatically controls engine power to maintain the best angle of attack during landing approaches. This permits the pilot to focus on flying the approach. During an APC approach in light to moderate turbulence, the set can maintain airspeed within a range of ±4 knots.

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Figure 5-24 — Approach power compensator simplified schematic.

T he major components of the system are a control panel, accelerometer, computer, control amplifier, a potentiometer to detect changes in elevator position, and a rotary actuator, which moves the throttle linkage. The aircraft angle-of attack transducer supplies angle-of-attack signals to the computer. A compression switch on the landing gear breaks the APC circuit upon touchdown. With the landing gear down and weight off the gear, the APC is energized by placing the engage switch to ON. The switch is magnetically held on until the aircraft touches down, the throttle overrides the system, or the pilot places the engage switch to OFF. The APC will also release automatically if either the override switch or the compression switch fails. 5-56

With the set engaged, the accelerometer, angle-of-attack transducer, and the ambient temperature selector switch supply information to the computer. The APC computer also receives elevator position signals from a potentiometer mounted in the control valve linkage of the elevator actuator. When normal acceleration is 1 g and the angle of attack is optimum for a landing approach, the computer sends no corrective signal and the throttle position does not change. The computer interprets deviations from these values as either offsetting each other or as requiring a change of power setting. When a power change is necessary, the computer sends an electrical signal to the rotary actuator through the control amplifier. The rotary actuator motor then drives the engine control linkage to accomplish the required power change. The system is capable of driving the throttle linkage at speeds up to 25 degrees per second. The APC can vary the engine RPM from full throttle to 67 percent RPM at temperatures above 0 °F, or to 58 percent RPM at temperatures below 0 °F. Since engine performance varies with ambient air temperature, the APC includes a three-position temperature switch to compensate for this effect. At low ambient air temperatures, the thrust change per degree of throttle movement is greater than at high ambient air temperatures. When the APC temperature switch is in COLD, the APC operates with a 15 percent reduction in gains from those with the switch in standard (STD). Conversely, with the temperature switch in HOT, the APC operates with an approximate 11 percent increase in gains from STD. This results in APC performance that is essentially the same regardless of ambient air temperature, with the following temperature settings: COLD below 4 °C (40 °F), STD from 4 °C to 27 °C (80 °F), and HOT above 27 °C. BLADEFOLD SYSTEM Blade fold on the H-60 helicopter is performed by an electromechanical system. The bladefold index drive unit positions the main rotor head before folding. There are four bladefold actuators, one for each blade. A tail rotor blade positioner actuator positions the tail rotor for pylon and stabilator folding. The blade deice distributor assembly on the main rotor contains sequencing electronics and power switching circuits for bladefold and pitchlock actuators. Power goes to the main rotor distributor through the main rotor slip ring. The bladefold control panel contains system operating controls and status/cue light capsules. You can manually fold the pylon by using a folding pole. A lockpin switch on the pylon indicates when the pylon is locked in flight position. The tail cone-pylon fold point has a crack open switch. This switch indicates pylon disconnection from the tail cone. Two stabilator lockpin switches indicate when stabilator lockpins are locking the stabilator panels in flight position. In addition, switches on each rotor hub indicate fold or spread status, and pitchlock engage or disengaged status. The switch signals go to the bladefold control panel and distributor. Master Power Distribution DC power for the system is from the No. 2 dc primary bus. This power goes through the weight-on-wheels and low oil pressure relays. The bladefold system cannot operate unless the helicopter is on the ground and the rotor head not moving. 5-57

The transmission pressure switch energizes the transmission low oil pressure relay. With both weight-on-wheels and rotor head stopped, the 28 Vdc master power goes to the blade deice junction box. The 28 Vdc at the blade deice junction box goes to two hydraulic system interlock relays. These relays actuate by pressure switches in the No. 1 and No. 2 hydraulic systems. To protect the flight controls, pressure can go to only one stage of the main rotor servos during fold operation. One hydraulic system must be shut off by the SERVO switch on either collective stick grip. Pressure switches operate relays to ensure that only one hydraulic system is on. Relay operation permits master power to the blade fold control panel only when one hydraulic system is on. The 28 Vdc rotor brake pressure switch senses rotor brake condition. The pressure switch actuates a relay, which applies a signal to the bladefold control panel. This signal is for the bladefold control panel indexing logic. The ROTOR BRAKE-APPLY-RELEASE indicators on the bladefold control panel advises pilots when to apply or release the rotor brake. Fold Sequence The first step in folding the main rotor blades is to index the rotor head to a known position. The bladefold index drive unit accomplishes this task. A de-operated linear actuator in the unit engages the indexing motor output gear to the rotor brake disc gear. An index command from the bladefold index drive unit drives rotor head to indexed position. A 2.5-degree maximum cutaway segment of an index slip ring in the slip ring assembly senses this position. With the BLADEFOLD switch in the FOLD position, and the rotor brake disengaged, the rotor can drive in either the clockwise or counterclockwise direction. The rotor will drive in the direction of the shortest distance to the index position. To do this, index drive slip ring construction is in two segments separated at the index position, 180 degrees from index position, and by providing a reversing relay in the index drive unit. The index motor rotates at 0.5 RPM. As the rotor head reaches the 1.0-degree index position, index drive motor power switches off and the drive motor brake de-energizes, applying the brake. Power removal and brake application prevent the rotor head from overshooting the index position. When the rotor head position is on the insulated area between index drive slip ring segments 180 degrees from the index position, head rotation will be counterclockwise toward the index position. The index unit can also act as a backup system to the rotor brake. You can use the index unit to engage the brake disk and to hold the rotor head stationary when performing maintenance on the rotor brake. To do this, use the GUST LOCKLKD- UNLKD switch on the miscellaneous switch panel. The switch enables the index unit to extend and engage the rotor brake disk, without applying power to the index motor. A brake and gear train inside the motor prevents motor output shaft rotation with no power to the motor. The index unit effectively acts as a lock to hold the rotor head in place. Power goes to the GUST LOCK switch from No. 2 dc primary bus interlocked through weight-on-wheels and transmission low oil pressure relays, so the GUST LOCK can only engage on the ground. When the index unit actuator retracts fully, a limit switch within the actuator activates a second relay inside the panel. This relay turns on a GUST LOCK caution light on the caution advisory panel. 5-58

DC power extends or retracts the index unit. Three-phase ac power operates the index motor. When the rotor head is indexed, logic circuits in the bladefold control panel turn on an INDEXED status light. The logic also senses the condition of the main rotor blades, tail pylon, stabilator, and rotor brake. If main rotor blades are spread and indexed, pylon and stabilator are in flight position. With the rotor brake on, a pitch blades command goes to the DAFCS computer. The computer flashes the BAR ALT pushbutton legend on the AFCS CONTROL panel. When the operator presses this push-button for 2 seconds, the TRIM push-button legend goes on. This causes the computer to send flight control positioning signals to the pilot-assist servos. The pilot-assist servos move the flight controls to a computer memorized position. With controls positioned, the computer sends an enable signal to the pitchlock logic in the bladefold control panel indicated by the flashing TRIM push-button legend. The control panel delivers a pitchlock command through the slip ring to the bladefold deice distributor. This power goes to the distributor for pitchlock actuators and bladefold actuators. This 115-Vac power also goes to the pitchlock actuator on each rotor blade hub. Each actuator advances a Iockpin that locks each blade pitch control horn. When all control horns lock, a signal goes through the slip ring to the bladefold control panel. This turns on a PITCH LOCKED status light. As each blade’s pitch lockpin engages its respective pitch control horn, the bladefold actuator operates. Each actuator, through gearing, unlocks its respective blade by pulling out two blade lockpins. Once the blade lockpins pull out, worm gearing drives a segment gear, which folds back each blade. Switches on each blade sense when the blade lockpins pull out and each blade reaches a folded position. When all four blades actuate the bladefold switches, a signal is sent to logic circuits in the distributor. This signal flows through the slip ring to the bladefold control panel, turning on a FOLDED status light. Each blade has individual logic circuits in a bladefold module. Should any individual logic circuit fail, the associated blade will not fold. This does not affect the other blades in the system, and they will fold normally. A spread blade can fold manually. If a loss of command signals or power occurs, this will prevent all blades from folding. A cycle caution logic circuit incorporated in the distributor monitors failures that are common to all blades. The cycle caution output goes to test points in the bladefold module for troubleshooting and fault isolation of the system. Spread Sequence You accomplish main rotor blade spreading by setting the BLADEFOLD switch to SPREAD. The switch routes a spread command through the slip ring to the blade deice distributor. The spread command also goes to the junction box and energizes a relay. This causes the power to the distributor circuits to reverse phase, allowing the bladefold actuators to run in the spread direction, spreading each rotor blade. As each blade reaches the full spread position; two blade lockpins drive into the blade hinge lugs. Two switches sense lockpin positions. When both switches actuate, a relay in the distributor energizes and reverses the phase of power to the actuators. The actuators run in the fold condition again for a brief period. This relieves stresses built up on the segment gear in each bladefold actuator. 5-59

CAUTION Damage may occur to the fuselage if the stabilator is in the fold position. Running the actuators in the fold direction causes the blade lockpins to back slightly out of the hinge lugs. One of the blade lockpin switches actuates before the other. This switch remains actuated even though the lockpins are backing out of the hinge lugs. The other switch de-actuates when the lockpins back out. When one lockpin switch actuates and the other is de-actuated, distributor logic determines achievement of the full spread condition. This de-energizes the bladefold motor actuator on each blade. Because the lockpins pull out only a short distance, they still safely lock the blades in flight position. Logic circuits in the distributor sense the fully spread conditions, spread command, and pitch lockpin IN condition for each blade. At this point, a relay actuates that pulls out the pitch lockpin in each blade pitch control horn. Simultaneously, a signal goes to a control panel spread complete logic circuit. This circuit actuates a 2-second time-delay relay. The relay shuts off hydraulic pressure to the pilot-assist servos and provides a logic signal to the DAFCS computer. The logic signal to the DAFCS computer tells the computer to update flight control position information from the transducers for 2 seconds. This is accomplished when one stage of hydraulic pressure is off, preventing preloading of the flight controls. Any preloading of the controls would cause an error in position information to the computer. After the 2-second delay, the computer blinks the RDR ALT push-button legend on the AFCS CONTROL panel. This informs the operator that the computer has accepted the flight control position update. The computer will place the flight controls in the new memorized position during the next fold cycle. Pylon and Stabilator Fold The first step in manually folding the pylon is to complete bladefold operation. By doing this, you prevent rotor blade and pylon damage. The crack open switch receives 28 Vdc, which comes through contacts of the weight- on-wheels relay. When the pylon opens about 5 degrees, the crack open switch applies power to the tail rotor blade positioner actuator. The actuator extends, indexes, and gust locks the tail rotor. It extends until the internal extend limit switch disconnects power. At the same time, another internal switch applies an actuator extended signal to the bladefold control panel logic circuits. If main bladefolding is operating at the same time as pylon folding, the actuator extended signal stops main fold. This prevents the main blades from striking the pylon. Two stabilator lock pin switches close when the stabilator locks in the flight (spread) position. This applies 28 Vdc to the automatic engagement circuit in the stabilator amplifier. With the stabilator lock pin switches in non-flight position, power is removed from the No. 1 and No. 2 stabilator amplifier auto engage circuitry. The auto engage circuitry engages when the stabilator AUTO CONTROL PUSH TO RESET pushbutton is depressed.

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At the same time, the bladefold control panel logic circuits receive a ground signal. With the main rotor blades properly folded, the bladefold control panel logic circuitry causes the SPREAD INCOMPLETE caution capsule to light under the following conditions:  Power goes to the fold or spread relays without being commanded by the bladefold control panel.  Pylon or stabilator panel spread but not locked.  Tail rotor blade positioner actuator extends with the pylon in the flight condition. ENGINE INLET BLEED AIR SYSTEM The engine inlet bleed air system of the F/A-18C/D Hornet provides the best inlet airflow for subsonic and supersonic flight. The inlet bleed air doors control inlet airflow by opening at 1.33 Mach to bleed boundary layer air off the compression ramp and close at 1.23 Mach. This ensures the doors will not oscillate when the aircraft travels at 1.33 Mach. A bleed air channel controls the fuselage boundary layer airflow near the wing roots. The bleed air doors open when three-phase, 115-V power is applied, and the controlling FLIGHT CONTROL COMPUTER (FCC) has a loss of power. After normal engine shutdown, the bleed air doors normally close. In some models of the F/A-18 aircraft when external power is on or the APU is operating in the ground maintenance mode and FCCs off, the bleed air doors will open. When FCCs are turned on, the bleed air doors will close. The FLT CONTR GND PWR RELAYS No. 11 and No. 12 prevent the cycling of the bleed air doors during ground operations. The Air Data Sensor, DT-600/ASW-44, computes the free air Mach number and sends the signal to FCCA and FCCB. When the Mach number is below 1.23, the FCCs supply a ground, energizing the retract circuits. Each actuator torque motor energizes, and the actuator brakes release. The torque motors retract the bleed air doors (if the doors are not retracted) until the actuator retract limit switches close. The retract limit switches de- energize the retract circuits, and the actuator brakes are applied by spring pressure. When the Mach number is above 1.33, the FCCs remove the ground, de-energizing the control relays and energizing the extend circuits. Each actuator’s torque motor energizes and the actuator brakes release. The torque motors extend the bleed air doors (if doors are not extended) until the actuators extend limit switches close. The extend limit switches de-energize the extend circuits, and the actuator brakes are applied by spring pressure. The extend limit and retract limit switches in each bleed air door actuator makes a ground to signal the signal data converter when the inlet bleed door fully extends or retracts. When ground is removed, the Signal Data Converter measures the time in transient until the other limit switch closes. If the time in transient exceeds 8 seconds, the Signal Data Converter signals the digital computer. The digital computer commands the cockpit DDI to display a caution.

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End of Chapter 5 AIRCRAFT POWER PLANT ELECTRICAL SYSTEMS Review Questions 5-1. In the air turbine starter, what unit limits rotational speed?

A. Scroll assembly B. Reduction gear C. Overspeed switch mechanism D. Output shaft

5-2. The air turbine starter converts energy from what type of air to shaft power?

A. Compressed air B. Ram air C. Fuel air D. Inlet air

5-3. The external stop switch de-energizes the air turbine start control holding relay by ________.

A. holding the relay B. removing the positive potential C. momentarily closing D. actuating the relay box

5-4. Why is timing not a factor in jet engine ignition systems?

A. Engine only needs igniting when abnormal fuel is used B. Engines need ignition 45 percent of the time C. Fuel and oil pressure changes D. Jet engines run on a continuous burning fire

5-5. At what engine RPM does ignition normally stop?

A. Between 40 percent and 60 percent of the engine rated speed B. Between 45 percent and 65 percent of the engine rated speed C. At 75 percent of the engine rated speed D. At 100 percent of the engine rated speed

5-6. Engine ignite automatically reactivates when either of what two conditions occurs?

A. Centrifugal speed switches are energized or closed. B. Voltage is on the relays or switches are on. C. A flameout occurs or during aircraft weapons fire. D. Engine is operating below 5 percent or a flameout occurs.

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5-7. What unit in an ignition system develops enough voltage to produce a spark?

A. Coil plugs B. The ignition exciter C. Spark plugs D. Ignition relay switches

5-8. What are the four positions of the E-2 aircraft condition lever?

A. Feather, ground stop, run, and air start. B. Ground, run, ram air and vent C. Run, air, trim, and slat D. Star, air, run and feather

5-9. When is the temperature datum control in the temperature controlling range?

A. Power lever is above 60°F and temp switch is in NORM B. Power lever is above 66°F coordinator and the temp datum switch is in AUTO. C. Power is at AUTO and temp datum switch is at AUTO D. Temperature is below 66°F

5-10. List the main components of an engine coordinator.

A. Pushrods, sectors and cables. B. Cable sector, coordinator shaft, and a temperature control switch, C. Variable potentiometer, discriminating device, and a cam-operated switch. D. Fuel control, fuel shutoff valve and condition lever.

5-11. At what RPM is the engine limiting temperature 830°C?

A. Above 95 percent B. Above 98 percent C. Below 94 percent D. At 99 percent

5-12. What are the thermocouple leads made of?

A. Steel and iron B. Brass and Chromel C. Brass and Alumel D. Chromel and Alumel

5-13. What are the three switches in the speed sensitive control?

A. 12, 16, and 65 percent B. 16, 65, and 94 percent C. 65, 75, and 95 percent D. 75, 95, and 98 percent

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5-14. How much greater is the capacity of the fuel pump primary element than the secondary element?

A. 5 percent B. 8 percent C. 10 percent D. 15 percent

5-15. Why is the fuel control schedule 20 percent richer than nominal engine requirements?

A. To maintain higher temperature. B. To accommodate the temperature datum valve. C. To maintain engine temperature at 16 percent RPM. D. To accommodate the accessory drive housing.

5-16. What unit minimizes the amount of fuel dropping into the combustion liners during engine shutdown?

A. Drain valves B. Check valves C. Flow liners D. Fuel lines

5-17. During a P-3 engine start, at what percent RPM does the start control switch de- energize?

A. 65 B. 75 C. 85 D. 95

5-18. What are the two types of anti-icing/deicing systems used on aircraft?

A. Manual and automatic B. Hot bleed air and electrical. C. Sensor and manual D. Electrical and manual

5-19. What is the difference between propeller deicing and anti-icing?

A. Deicing removes ice already built up whereas anti-icing removes moisture in the air on the component. B. Deicing removes ice buildup properties and anti-icing is physically removing the ice. C. Deicing removes ice already built up whereas anti-icing prevents ice buildup. D. Deicing prevents ice buildup whereas anti-icing removes ice already built up.

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5-20. What point in the engine is the coldest and most likely to ice?

A. At the solenoid operated bleed valve. B. At the valve face. C. Air at the inlet guide vanes. D. At the regulating piston.

5-21. In supersonic flight, icing does not occur for what reason?

A. There is limited time for moisture to freeze. B. Air over the aircraft at the altitude to obtain speed is dry air. C. Speed breaks the ice when it is formed. D. Friction of the air over the aircraft creates enough heat to prevent icing.

5-22. The fire warning element has an inverse temperature coefficient. What does this statement mean?

A. As the temperature decreases, the resistance of the sensing element decreases. B. As the temperature increases, the resistance of the sensing element decreases. C. As the temperature increases, the resistance of the sensing element increases. D. As the temperature decreases, the resistance of the sensing element increases.

5-23. What is the circuit that differentiates between an overtemp (fire) and a short circuit in a fire warning system?

A. Sensing B. Relay warning C. Fire warning D. Short discriminator

5-24. What chemical is used in the engine extinguishing system HRD bottles?

A. CO2 B. AFFF C. PKP D. Halon

5-25. What fuel tanks feed the engines in the F /A-18 fuel transfer system?

A. 1 and 3 B. 1 and 4 C. 2 and 3 D. 2 and 4

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5-26. When does fuel transfer from the wing start?

A. After the fuel level drops below the low pilot valves in tanks 1 and 4. B. When fuel level drops below the high level pilot valves in tanks 1 and/or 4. C. Before the fuel level drops below 45 percent in tanks 1 and/or 4. D. Fuel transfers continuously at all times.

5-27. When does the fuel transfer center of gravity control system energize?

A. When fuel tank 1 transfers at a faster rate than fuel tank 3. B. When fuel tank 1 transfers at a faster rate than fuel tank 4. C. After fuel tank 2 transfers with fuel tank 3. D. After fuel tank 3 transfers with fuel tank 4.

5-28. When the engine is shut down and the throttle is in the off position, the variable exhaust nozzle is in what position?

A. Open B. Standby C. Halfway open D. Closed

5-29. What unit in the variable exhaust nozzle adjusts the VEN for atmospheric conditions?

A. Afterburner Flame Sensor B. Afterburner Control C. Electrical Control Assembly D. Main Fuel Control

5-30. What are the two ranges of propeller operation?

A. Flight and ground B. Low speed and high speed C. Takeoff and roll D. Normal and reverse

5-31. What mode of operation is the synchrophaser in when maintaining all propellers at the same RPM?

A. Pulse generator B. Speed-bias C. Phase control D. Governing

5-32. What unit provides a means of setting phase relationships between the master and slave propellers in the synchrophase system?

A. Phase and trim control unit B. Pulse generator C. Speed-bias D. Wipers 5-66

5-33. The synchrophaser speed derivative circuit is used to dampen changes in engine ________.

A. temperature B. fuel flow C. RPM D. torque

5-34. The slave propellers are prevented from following an overspeed or underspeed master propeller by what synchrophaser circuit?

A. Phase-error B. Off-speed C. Magnetic modulator D. Limiting (two percent)

5-35. The propeller is used to convert engine shaft horsepower to ________.

A. force B. thrust C. pressure D. torque

5-36. What mechanical unit maintains a minimum blade angle?

A. Low pitch stop B. Propeller housing C. Propeller dome D. Flight power lever

5-37. What does the beta follow-up system provide?

A. Engine power for takeoff B. Programming for the hydraulics stop C. Variable hydraulic low pitch stop D. Negative torque

5-38. What system resets pitchlock RPM from 103.5 percent to 109 percent during an aborted takeoff?

A. Block out range B. Feathering C. Override reset D. Pitch lock reset

5-39. What valve bypasses all other control functions to feather the propeller?

A. Auto pitch B. Feather C. Override D. Overspeed

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5-40. What propeller system is used only during takeoff?

A. Automatic feathering B. Thrust sensitive C. Pitchlock D. Cutout override

5-41. During ground propeller unfeathering, how do you decrease blade angle below 45 degrees?

A. Energize air start system B. Actuate propeller governor C. Actuate the pressure cutout override switch D. Energize negative torque

5-42. To maintain the best angle- of-attack during landing approaches, the approach power compensator automatically controls ________.

A. ambient temperature selector switch B. compression switch C. engine temperature D. engine power

5-43. The approach power compensator circuit de-energizes when the pilot turns it off, the aircraft touches down, and ________.

A. the throttles override the system B. the switch is magnetically held on C. the landing gear is down D. the compression switch fails

5-44. What does the approach power compensator use to compensate for differences in ambient temperatures?

A. A 11 percent increase in gain switch B. A 15 percent reduction switch C. A three-position temperature switch D. A cold and hot switch

5-45. In the H-60 bladefold system, the transmission low oil pressure relay is energized by a ________.

A. hydraulic interlock relay B. transmission pressure switch C. blade junction box D. control panel

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5-46. When the rotor head is on the insulated area between index drive slip ring segments 180 degrees from the index position, in what direction will the rotor head travel towards index?

A. Remains stationary B. Up and down C. Clockwise D. Counterclockwise

5-47. During bladefold operations, if one blade fails to fold, will the other blades fold?

A. No B. Yes C. Only the next blade can fold manually D. Two blades will not fold normally

5-48. During blade spread operations, the actuators run in the fold condition after lockpin position switches activate to ________.

A. ensure each blade reaches the full spread position B. relieve stress built up on the segment gear C. run the actuators D. actuate a relay

5-49. During pylon and stabilator fold, at what point does the crack open switch power the tail rotor blade position actuator?

A. Pylon completes bladefold operation B. Tail rotor is locked C. Pylon opens about 5 degrees D. Main blades stop folding

5-50. The ratio of the speed of an object to the speed of sound in the same medium and at the same temperature is known as ________.

A. subsonic B. supersonic C. speed D. mach number

5-51. The F /A-18 engine bleed air system duct doors are controlled by what subsystem?

A. Flight control computers B. Flight control ground C. Air data control D. Bleed air channel

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CHAPTER 6 AIRCRAFT INSTRUMENTS When the first aircraft came into existence, the main goal was to launch the aircraft and keep it airborne as long as possible. At first, it was not possible to keep the aircraft in the air for longer than a few minutes. However, as engines and aircraft structures were improved, the aircraft was able to remain aloft for a longer time. Along with these improvements came the need for instruments. The first aircraft instruments were fuel and oil pressure instruments. These instruments warned the pilot of engine trouble so the aircraft could be landed before the engine failed. Later, when the aircraft could fly over considerable distances, weather became a problem. This led to the development of instruments that helped pilots fly through snowstorms, thunderstorms, and other bad weather conditions. The instruments used in aircraft years ago are reasonably simple compared with those in current aircraft. The jet aircraft has brought many complex problems to instrument engineering. Instrumentation is basically the science of measurement. Measurements that are common on all aircraft are position, direction, speed, altitude, engine condition, fuel on board, and fuel consumption. In addition, jet aircraft instruments include Mach speed, angle of attack, and Exhaust Gas Temperature (EGT) indicators. There are two ways of grouping aircraft instruments, by their operating principles and by the job they perform. Instrument operating principles include gyroscopic, pressure or temperature sensing, magnetism, electrical energy, or a combination of any of these. This chapter deals with indicating systems and flight, engine, and equipment instruments in relation to the jobs they perform. The flight instruments discussed are those instruments that provide aircraft performance information to the pilot. These instrument systems include the airspeed, altimeter, vertical speed, attitude, turn and bank, and angle-of-attack. Along with the heading indicator, these instruments provide primary flight reference to the pilot. The heading indicator is a flight instrument, which is discussed in the Compass and Inertial Navigation Systems (INS) chapter. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Describe the operating principles and features of aircraft flight instrument systems, including the pitot-static, airspeed indicator, angle-of-attack, gyroscope, and miscellaneous flight instrument systems. 2. Explain the operating principles and characteristics of engine instrument systems, including tachometer, temperature indicating, fuel flow, oil pressure, fuel pressure, oil temperature, exhaust-nozzle indicating, and torquemeter systems. 3. Summarize the operating principles and characteristics of aircraft instrument systems, including fuel quantity, hydraulic pressure, and position indicating systems. 4. Identify the various procedures used to maintain, and test aircraft instruments. 6-1

Figure 6-1 — Mercurial barometer. FLIGHT INSTRUMENT SYSTEMS To maintain instruments properly, you, as an Aviation Electrician’s Mate (AE), must know the basic principles of the flight instrument systems. AEs frequently work with equipment and systems that use the principles of density and pressure. You must consider density and pressure when discussing altimetry and airspeed. Although very light, air has weight and is affected by gravity. By its weight, air exerts pressure on everything it touches. Since air is a gas, it exerts pressure in all directions. The weight of the air pressing down from above determines the air pressure at any given altitude. The weight of the atmosphere presses the molecules closer together, making them more numerous per unit of volume. This action takes place at the bottom of the atmosphere, or where it rests upon the earth’s surface. Therefore, the air at the bottom of the atmosphere is denser than at higher altitudes. Air pressure at sea level on an average day will support a column of mercury 29.92 inches high (Figure 6-1).

Atmospheric pressure is a force per unit area, and force is equal to mass multiplied by acceleration. Therefore, a pressure change occurs if either the mass of the atmosphere changes or the molecules within the atmosphere accelerate. Although altitude exerts the dominant control, temperature and moisture alter pressure at any given altitude. Figure 6-2 shows the standard pressure and temperature at given altitudes.

6-2

Figure 6-2 — The standard atmosphere. Conditions are very seldom standard for temperature or pressure; therefore, you must correct the formula to find density altitude or true airspeed. Let’s consider an airfield under the influence of a low-pressure climatic condition, where the temperature is very hot. Together, these two conditions may reduce the density of the air to such an extent that it affects aircraft engine performance. This reduction of air density makes takeoff capability marginal, especially for a helicopter. The density of air also directly affects aircraft movement through the air, and thus the true airspeed of the aircraft. The denser the air, the more difficult it is for the aircraft to move through it.

6-3

Figure 6-3 — Pitot-static system. PITOT-STATIC SYSTEM The aircraft pitot-static system (Figure 6-3) includes instruments that operate on the principle of the barometer. The system consists of a pitot tube, static air vents, and three indicators, which connect with pipelines that carry air. The three indicators are Airspeed, Altimeter, and the Vertical Speed. The airspeed indicator shows the speed of the aircraft through the air, and the altimeter shows the altitude. The Vertical Speed Indicator (VSI) indicates how fast the aircraft is climbing or descending. All of these indicators operate on air that comes in from outside the aircraft during flight. The pitot tube mounts on the outside of the aircraft (Figure 6-3) at a point where the air is least likely to be turbulent. The tube points in a forward direction parallel to the aircraft’s line of flight. One general type of airspeed tube mounts on a streamlined mast extending below the nose of the fuselage. Another type mounts on a boom extending forward from the leading edge of the wing. Although there is a slight difference in their construction, the tubes operate identically. The Pitot System measures impact pressure, which is the pressure of the outside air against the aircraft flying through it. The tube that goes from the pitot tube to the airspeed indicator applies the outside air pressure to the airspeed indicator. The airspeed indicator calibration allows various air pressures to cause different readings on the dial. The purpose of the airspeed indicator is to interpret pitot air pressure in terms of airspeed in knots. Generally, static air vents (Figure 6-3) are small, calibrated holes in an assembly mounted flush with the aircraft fuselage. Their position is in a place with the least amount of local airflow moving across the vents when the aircraft is flying. Static means stationary or not changing. The static part of the pitot-static system also introduces outside air. However, the outside air is at its normal outside atmospheric pressure as though the aircraft were standing still in the air. The static line applies this outside air to the airspeed indicators, the altimeter, and the vertical speed indicator. 6-4

Figure 6-4 — (A) Airspeed indicator; (B) maximum allowable airspeed indicator. Airspeed Indicators Readings from an airspeed indicator are used to estimate ground speed and to determine throttle settings for the most efficient flying speed. These readings also provide a basis for calculating the best climbing and gliding angles. They warn the pilot if diving speed approaches the safety limits of the aircraft’s structure. Since airspeed increases in a dive and decreases in a climb, the indicator is an excellent check for maintaining level flight. Figure 6-4, view A, shows a cutaway view of a typical airspeed indicator. An airspeed indicator has a cylindrical, airtight case that connects to the static line from the pitot-static tube. Inside the case is a small aneroid diaphragm of phosphor bronze or beryllium copper. The diaphragm is very sensitive to changes in pressure, and it connects to the impact pressure (pitot) line. This construction allows air from the pitot tube to enter the diaphragm. The side of the diaphragm fastens to the case and is rigid. The needle or pointer connects through a series of levers and gears to the free side of the diaphragm. The airspeed indicator is a differential pressure instrument. It measures the difference between the pressures in the impact pressure line and in the static pressure line. The two pressures are equal when the aircraft is stationary on the ground. Movement through the air causes pressure in the impact line to become greater than that in the static line. This pressure increase causes the diaphragm to expand. The expansion or contraction of the diaphragm goes through a series of levers and gears to the face of the instrument to regulate needle position. The needle shows the pressure differential in MPH or knots. All speeds and distances are in nautical miles. MAXIMUM ALLOWABLE AIRSPEED INDICATOR – Figure 6-4, view B, shows the face of a maximum allowable airspeed indicator. The dial face measurements are in 6-5

Figure 6-5 — Airspeed/Mach speed indicator. knots from 50 to 450 with an expanded scale below 200 knots. The dial has an indicating pointer and a maximum safe airspeed pointer. The maximum safe airspeed pointer moves as the maximum safe airspeed changes because of static pressure changes at different altitudes. No matter where the pitot-static tube is located, it is impossible to keep it free from all air disturbances set up by the aircraft structure. You must make allowances for this installation error when reading the indicator. Temperature is another cause of error. Also, imperfect scaling of the indicator dial with respect to the airspeed differential pressure relationship will cause an error in reading. You can make simple adjustments to the instrument mechanism to correct the tendency to read fast or slow. MACH SPEED INDICATORS – In some cases, the term Mach speed is used to express aircraft speed. The Mach speed is the ratio of the speed of a moving body to the speed of sound in the surrounding medium. For example, if an aircraft is flying at a speed equal to one-half the local speed of sound, it is flying at Mach 0.5. If it moves at twice the local speed of sound, its speed is Mach 2. Figure 6-5 shows the front view of a typical airspeed and Mach speed indicator. The instrument consists of altitude and airspeed mechanisms incorporated in a single housing. This instrument gives the pilot a simplified presentation of both indicated airspeed and Mach speed. Both indications are read from the same pointer.

6-6

Figure 6-6 — Airspeed/Mach speed indicator mechanical schematic. The pointer shows airspeed at low speeds, and both indicated airspeed and Mach speed at high speeds. Pitot pressure on a diaphragm moves the pointer, and an aneroid diaphragm controls the Mach speed dial. The aneroid diaphragm reacts to static pressure changes because of altitude changes. Figure 6-6 is a mechanical schematic of an airspeed and Mach speed indicator.

The range of the instrument is 80 to 650 knots indicated airspeed and from 0.5 to 2.0 Mach speed. Its calibrated operating limit is 50,000 feet of altitude. A stationary airspeed dial masks the upper range of the movable Mach dial at low altitudes. The stationary airspeed dial is graduated in knots. The instrument incorporates a landing speed index and a Mach speed setting index. You can adjust both indexes by a knob on the lower left-hand corner of the instrument. You can adjust the landing speed index over a range of 80 to 150 knots. The index operates with the knob in its normal position. You may adjust the Mach speed index over the entire Mach range. The index adjusts by depressing the knob and turning it.

6-7

Altimeter An altimeter is an instrument that measures static pressure. Before you can understand how the altimeter works, you need to understand altitude. Remember, even though the altimeter reads in feet, it is actually measuring pressure. The word altitude is vague, so it needs further defining. The term altitude includes altitude above Mean Sea Level (MSL) and altitude Above Ground Level (AGL). It also includes pressure altitude, indicated altitude, density altitude, and elevation. MEAN SEA LEVEL – Since about 80 percent of the earth’s surface is water, it is natural to use sea level as an altitude reference point. The pull of gravity is not the same at sea level all over the world because the earth is not perfectly round and because of tides. To adjust for this, an average (or mean) value is set; this is the mean sea level. Mean sea level is the point where gravity acting on the atmosphere produces a pressure of 14.70 pounds per square inch. This pressure supports a column of mercury in a barometer to a height of 29.92 inches. This is the reference point from which you measure all other altitudes. See Figures 6-1 and 6-2. The altitude you read from an altimeter refers to MSL. ELEVATION AND TRUE ALTITUDE – Elevation is the height of a land mass above MSL. Elevation is measured with precision instruments that are far more accurate than the standard aircraft altimeter. You can find elevation information on charts or, for a particular spot, painted on a hangar near an aircraft ramp or taxi area. True altitude is the actual number of feet above MSL. A ruler or yardstick is used to measure the altitude. In standard day conditions, pressure altitude and true altitude are the same. ABSOLUTE ALTITUDE – Absolute altitude is the distance between the aircraft and the terrain over which it is flying. It is referred to as the altitude Above Ground Level (AGL). Due t o variations in terrain, AGL is typically unreliable information. However, it is useful when flying near the ground, such as in a takeoff or landing pattern. You find AGL by subtracting the elevation of the terrain beneath the aircraft from the altitude read on the altimeter (MSL). A radar altimeter indicates actual altitude above the terrain; you call this indication radar altitude. PRESSURE ALTITUDE – To measure altitude, instruments sense air pressure and compare it to known values of standard air pressure at specific, measured altitudes. The altitude you read from a properly calibrated altimeter referenced to 29.92 inches of mercury (Hg) is the pressure altitude. Refer back to Figure 6-2. If a pressure altimeter senses 6.75 pounds per square inch pressure with the altimeter set to sea level and barometric pressure 29.92 inches of mercury, the altimeter indicates 20,000 feet. This reading does not mean that the aircraft is exactly 20,000 feet above MSL. It means the aircraft is in an air mass exerting a pressure equivalent to 20,000 feet on a standard day. You can see that pressure altitude is not true altitude. INDICATED AND CALIBRATED ALTITUDE – Unfortunately, standard atmospheric conditions very seldom exist. Atmospheric conditions and barometric pressure can vary considerably. A pressure change of one-hundredth (0.01) of an inch of mercury represents a 9-foot change in altitude at sea level. Barometric pressure changes between 29.50 and 30.50 are not uncommon (a pressure change of about 923 feet). Indicated altitude is the uncorrected reading of a barometric altimeter. Calibrated altitude is the indicated altitude corrected for inherent and installation errors of the 6-8

altimeter. On an altimeter without such errors, indicated altitude and calibrated altitude are identical. Assume that this is the case for the rest of this discussion. When flying below 18,000 feet, the aircraft altimeter must be set to the altimeter setting (barometric pressure corrected to sea level) of a selected ground station within 100 miles of the aircraft. Altitude read from an altimeter set to local barometric pressure is indicated altitude. The accuracy of this method is limited because you must assume a standard lapse rate; that is, for a given number of feet of altitude, an exact change in pressure occurs. This exact change seldom happens, which limits the accuracy of the altimeter. Above 18,000 feet, all altimeters are set to 29.92 (pressure altitude). Although the altimeter is not accurate, as long as all aircraft have the same barometric pressure setting, aircraft vertical separation is controlled. DENSITY ALTITUDE – A very important factor in determining the performance of an aircraft or engine is the density of the air. The denser the air, the more horsepower the engine can produce. Also, there is more resistance to the aircraft when flying resulting in airfoils producing more lift, and propellers producing more thrust. Pressure, temperature, and moisture content all affect air density. Measurements of air density are in weight per unit volume (for example, pounds per cubic foot). However, a more convenient measurement of air density for the pilot is density altitude. This is that altitude in the standard atmosphere which corresponds to a particular air density. Density altitude is the pressure altitude corrected for temperature deviations from the standard atmosphere. In basic terms, it is the altitude that the aircraft "thinks" it's at. An increase in density altitude corresponds to reduced air pressure felt by the aircraft. This results in airfields at higher elevations, particularly when warm temperatures are present to require more runway for aircraft to take off. Additionally, aircraft will have a reduced rate of climb and a faster approach and will experience a longer landing roll. Density altitude does not show on an instrument. It is usually taken from a table or computed by comparing pressure, altitude, and temperature. Although moisture content affects air density, its effect is negligible." Several kinds of altimeters are in use today. They are all constructed on the same basic principle as an aneroid. They all have pressure responsive elements (aneroid wafers) that expand or contract with the pressure changes of different flight levels. The heart of a pressure altimeter is its aneroid mechanism (Figure 6-7), which consists of one or more aneroid wafers. The expansion or contraction of the aneroid wafers with pressure changes operates the linkage. This action moves the indicating hand/counter to show altitude. Around the aneroid mechanism of most altimeters is a device called the bimetal yoke. As the name implies, this device is composed of two metals. It performs the function of compensating for the effect that temperature has on the metals of the aneroid mechanism. The altimeter discussed in the following paragraphs is a simple one. Several complex altimeters are discussed later in this chapter, along with the automatic altitude system. 6-9

Figure 6-7 — Simplified aneroid mechanism. Figure 6-8 — Counter pointer pressure altimeter.

COUNTER POINTER PRESSURE ALTIMETER – The purpose of the counter pointer pressure altimeter (Figure 6-8) is to show aircraft height. By studying the dial of the indicator, you can easily understand the procedure for determining the height of the aircraft. A description of the mechanical operation of this altimeter follows. As you read about the operation, refer to Figure 6-9.

6-10

Figure 6-9 — Mechanical schematic of a counter pointer pressure altimeter.

Atmospheric changes cause movement of the two aneroid diaphragm assemblies. These assemblies move two similar rocking shaft assemblies mutually engaged with the main pinion assembly. This movement goes to the handstaff assembly, which operates the hand assembly and drives the counter mechanism through a disk. Because of the special design of the hand assembly, the counter indication is never obscured. An internal vibrator minimizes friction during the instrument’s operation. You make barometric corrections by turning the externally located knob. The knob engages the barometric dial and the main plate assembly that supports the entire mechanism. You make adjustments so the reading on the barometric dial corresponds to the area barometric conditions in which the aircraft is flying. 6-11

Figure 6-10 — Vertical speed indicator (VSI). Figure 6-11 — Mechanical schematic of a VSI. Vertical Speed Indicator (VSI) A VSI shows the rate at which an aircraft is climbing or descending. It is very important for night flying, flying through fog or clouds, or flying when the horizon is obscured. Another use is to determine the maximum rate of climb during performance tests or in actual service. The rate of altitude change, as shown on the indicator dial, is positive in a climb and negative in a dive or glide. The dial pointer (Figure 6-10) moves in either direction from the zero point. This action depends on whether the aircraft is going up or down. In level flight the pointer remains at zero. The vertical speed indicator is contained in a sealed case, and it connects to the static pressure line through a calibrated leak. Refer to Figure 6-11. Changing pressures will result in expansion or contraction of the diaphragm, which in turn will move the indicating needle through the use of internal gears and levers. The instrument automatically compensates for changes in temperature. Although the vertical speed indicator operates from the static pressure source, it is a differential pressure instrument. The difference in pressure between the instantaneous static pressure in the diaphragm and the static pressure trapped within the case creates the differential pressure. When the pressures equalize in level flight, the needle reads zero. As static pressure in the diaphragm changes during a climb or descent, the needle immediately shows a change of vertical speed. However, until the differential pressure stabilizes at a definite ratio, indications are not reliable. Because of the restriction in airflow through the calibrated 6-12

NOTE To understand air data computer maintenance in modern aircraft, you must have knowledge of digital electronics, including logic diagrams and flow charts. Review NEETS, Module 13, Introduction to Number Systems and Logic Circuits, NAVEDTRA 14185, and Module 22, Introduction to Digital Computers, NAVEDTRA 14194 before continuing. leak, the differential pressure requires a 6 to 9 second lag for the pressures to stabilize. The VSI has a zero adjustment on the front of the case. You use this adjustment with the aircraft on the ground to return the pointer to zero. While adjusting the instrument, tap it lightly to remove friction effects. AIR DATA COMPUTER (ADC) SYSTEM Aircraft operating below 0.9 Mach airspeed use raw pitot and static pressures to develop accurate airspeed, altitude, and vertical speed indications. Aircraft operating in this speed range use the pressures that the pitot-static ports sense. Modern supersonic aircraft operate in a higher speed range and require more accurate pressures. At high speed, pressures build upon the external skin of the aircraft. These pressures cause a distortion of the normal flow of air, causing the pitot-static system to sense false pressures. The system then supplies erroneous information to the flight instruments. The altimeter, for instance, may show an error of more than 3,000 feet. A 3,000-foot error in altitude is intolerable and could put an aircraft in an extremely dangerous position. The system that compensates for altitude and other pitot-static errors is the Air Data Computer (ADC) system. Many variations exist in both the name of the systems and the method of data development. Purpose Many inputs are common to the various types of ADC systems. ADCs differ in how they process input data and distribute output data to the various systems using the data. Data requirements vary with the type and mission of the aircraft. Figure 6-12 shows the major distribution of systems that depend on all or part of the ADC. Notice that all inputs, such as pitot and static pressures, go to the ADC. The ADC receives pneumatic and electrical inputs to produce various outputs. Signals resulting from the processing of the inputs go to the using systems.

The ADC receives information from pressure-sensitive and temperature-sensitive units mounted on external points of the aircraft. Using this data, it compensates for errors and sends the corrected information to other systems in the aircraft. Concurrently, it detects any changes in pressure and temperature information. It converts these changes into usable signals and sends them along with the pressure and temperature signals. The electrical signal outputs are representative of altitude, Mach speed, true airspeed, angle of attack, total temperature, and impact pressure. There is also a pneumatic output of corrected static pressure. This output is used by the barometric altimeter, airspeed, VSIs, and some modules within the air data computer. 6-13

Figure 6-12 — Air Data Computer block diagram showing inputs and outputs.

6-14

There are four basic data inputs in every ADC system: 1. Total pressure (pitot) 2. Indicated static pressure 3. Indicated angle of attack 4. Total temperature Command and test signal inputs use the available raw and corrected primary data for making functional tests of various ADC outputs. Table 6-1 contains a list of symbols and their definitions. Since these symbols are used many times in this section, you should refer to this table for symbol meanings.

Table 6-1 — Symbols Used with an ADC System SYMBOL DEFINITION ADC Air data computer AOA Angle of attack BIT Built in test a i Indicated angle of attack a T True angle of attack B Constant ∆ Incremental change (delta) Hp Barometric or pressure altitude M Mach speed PD Pressure differential Ps Correct static pressure P Indicated static pressure Pt Correct total pressure Pti Indicated total pressure QA Actual impact pressure Qc Correct impact pressure Ts Free airstream temperature Tt Total temperature Tti Indicated total temperature Va True airspeed Vc Calibrated airspeed 6-15

Figure 6-13 — Angle of attack transmitter (probe). Major Components The major components that collect and distribute information used in the air data computer system are listed below. 1. Angle- of-attack transmitter 2. Pitot-static system 3. Total temperature probe Although the functions of these components are essentially the same on all aircraft the processing and distribution of air data information varies from aircraft model to model. When performing maintenance on any ADC system, you shall refer to the latest Maintenance Instructions Manual (MIM) for that particular aircraft model to ensure you use correct ADC system information. ANGLE-OF-ATTACK TRANSMITTER – Forces vary with the angle of attack. The angle of attack is the angle between the relative wind and the chord of the wing. The chord of the wing is a straight line running from the leading edge to the trailing edge. Increasing the angle of attack increases the pressure felt under the wing and vice versa. The angle-of-attack transmitter (Figure 6- 13) detects changes in the aircraft’s local angle of attack. It sends these changes, in the form of mechanical motion, to potentiometers within the transmitter. These potentiometers convert the mechanical motion to proportional electrical voltages. These voltages go to associated angle-of-attack indicating and interface equipment. The transmitter has a detector probe that senses changes in airflow. Changes in airflow cause the probe paddle to rotate. This rotation, in turn, drives the wiper arms of the three internally mounted potentiometers. The angle-of-attack system shows the pilot aircraft pitch attitude with respect to the surrounding air mass. 6-16

Figure 6-14 — ADC system airstream sensors. PITOT-STATIC SYSTEM – Figure 6-14 shows the airstream sensors of the pitot-static system. These sensors sense the air surrounding the aircraft and provide impact (pitot/Pt) pressure and atmospheric (static/Ps) pressure. These pressures go to the flight instruments and to the ADC. The pitot-static system is actually two separate systems with individual pitot-static probes (Figure 6-15), one on each side of the forward fuselage. The ADC receives static pressure (Ps) from both probes. However, it receives total pressure (Pt) from only one probe.

Indicated Static Pressure – This pressure (P) is the atmospheric pressure as sensed at a point on the aircraft that is relatively free from airflow disturbances. At subsonic speeds, static pressure error is small and of little significance. However, at transonic and supersonic speeds, the static ports sense extreme static pressure errors. Both Mach speed and angle of attack can cause significant errors in the static pressure system. Indicated static pressure (P), as detected by the aircraft static ports, deviates from true static pressure. These deviations have a definite relationship to Mach speed and angle of attack. The size of the error is the ratio of true static pressure to indicated static pressure, as related to Mach speed and angle of attack. 6-17

CAUTION Be sure to disable the pitot-static heater before working on this system. You may be seriously burned by touching the probes. Figure 6-15 — Pitot-static probe. Impact Pressure – As implied, impact pressure (Qc) is the force of the air against the aircraft. Qc is measured directly by use of a pitot-static probe (Figure 6-15) or calculated from the outputs of the static and total pressure transducers. The ADC calculates actual impact pressure (QA) as a function of Mach speed squared and static pressure.

Indicated Total Pressure – This pressure (Pti) is the sum of static air pressure and the pressure created by aircraft motion through the air. The pitot tube senses total pressure, which you also know by the familiar term pitot pressure. Corrected Static and Corrected Total Pressures – These pressures, Ps and Pt, contain errors that must be corrected to get true static and true total pressures. These errors are a result of slope and offset errors related to Mach speeds. The computer calculates the specified slope and intercept errors as functions of the indicated pressure ratio (Pti/P) and of the indicated angle of attack (a i).

6-18

Figure 6-16 — Total temperature circuit. TOTAL TEMPERATURE PROBE – Total temperature (Tt) is the temperature of ambient air plus the temperature increase created by the motion of the aircraft. Total temperature is sensed by a probe. This probe includes a platinum resistance element inside an aerodynamic housing placed in the airstream. The resistive element, whose resistance varies with temperature, acts as the variable portion of a bridge circuit. The total temperature probe provides the ADC with accurate outside air temperature information. The raw information is the indicated total temperature (Tti). The computer smooths and limits computations on the Tti before using the resultant output to calculate true T t. Figure 6-16 shows a typical temperature-sensitive bridge circuit that provides temperature data to the air data computer.

AUTOMATIC ALTITUDE SYSTEM In the past, the air traffic control system radar presented azimuth and distance information to the controller on horizontal radarscopes. Aircraft identification was done primarily by voice radio, the use of position reports over definite fixes, identifying turns of the aircraft to headings requested by the controller, or a beacon identification signal. Altitude information was given over the voice radio. After this information-gathering process, the information was recorded on a flight strip by the controller and updated as required. When the aircraft moved into another controller’s area, the handoff of the aircraft and the associated information was a manual process. Although the system was adequate, it became cumbersome during heavy traffic. 6-19

Figure 6-17 — The automatic altitude reporting system. The increase in air traffic since 1950 has caused serious problems of vertical separation, terrain clearance, and collision avoidance. Because of these problems, improved air traffic control techniques were developed. These techniques included the use of altitude-coded transponders for automatic altitude and position reporting. Automatic altitude reporting equipment that provides continuous automatic identification of aircraft on the ground controller’s radarscopes has been developed. This equipment cuts out many of the manual steps required in the old air traffic control system. An air data computer corrects static pressure errors and provides synchro-driven altitude information to the pilot’s altimeter. It also provides altitude in digital form to the aircraft transponder in high-performance aircraft. In low-performance aircraft, the equipment provides a direct readout of altitude to the pilot and digital altitude information to the aircraft transponder. The digital information then goes to the ground interrogator and shows on the radarscopes in alphanumeric form. The automatic altitude system operation is discussed in the following paragraphs utilizing the Identification Friend or Foe (IFF) system operation. An interrogation pulse group goes from the interrogator-transmitter unit through a directional interrogator antenna assembly. The pulse group triggers an airborne transponder, causing a multiple pulse reply group to be transmitted. The transponder transmission goes to the ground interrogator-receiver, which is processed through a computer. It is then displayed in alphanumeric form on the controller’s radar screen. The length of the round-trip transit time determines the range of the replying aircraft. The mean direction of the main beam of the interrogator antenna during the reply determines the azimuth. The encoded signal from the transponder provides, via mode C, the aircraft’s altitude in 100-foot increments. Refer to Figure 6-17, which shows the automatic altitude reporting system.

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Figure 6-18 — AAU-19/A, AAU-21/A, AAU-24/A altimeters. As you can see, a semi-automated air traffic control system includes the following improvements over the past system.  The automatic altitude reporting system automatically provides the air traffic controller with a radar presentation. It identifies, in three dimensions, every properly equipped aircraft within the controllers’ area.  Because of the three-dimensional presentation, the automatic altitude reporting system greatly reduces the use of voice radio. It also eases the workload of the air traffic controller, thus increasing air traffic control efficiency.  A transponder signal reinforces the radar signal normally seen on the radarscopes. It makes the signal stronger and much less susceptible to atmospheric interference.  The beacon system altitude reporting feature may reduce vertical separation in the higher flight levels.  The automatic altitude reporting system continuously updates aircraft altitude and records in 100-foot increments. Separations permits more accurate traffic control when aircraft are changing altitude rapidly, as they do in terminal areas. Altimetry The three altimeters that work with the automatic altitude reporting system are the AAU- 19/A, AAU-21/A, and AAU-24/A (Figure 6-18).

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Figure 6-19 — AAU-19/A altimeter. SERVOED BAROMETER ALTIMETER AAU-19/A – The counter-drum-pointer served barometric altimeter (Figure 6-19) consists of a pressure altimeter combined with an at- powered servomechanism. The altitude display is in digital form, using a 10,000-foot counter, a 1,000-foot counter, and a 100-foot drum. Also, a single pointer shows hundreds of feet on a circular scale. The barometric pressure setting (baroset) knob is used to insert the local pressure in inches of mercury. The baroset knob has no effect on the digital output (mode C) of the ADC. This digital output is always referenced to 29.92 inches of mercury. The altimeter has a servoed mode and a pressure mode of operation. The mode of operation is controlled by a spring-loaded, self-centering mode switch, placarded RESET and STBY. In the servoed mode, the altimeter displays altitude, corrected for position error, from the synchro output of the air data computer. In the standby mode, the altimeter operates as a standard altimeter. In this mode, it uses static pressure from the static system that is uncorrected for position error. The servoed mode is selected by placing the mode switch to RESET for 3 seconds. The ac power must be on. During standby operation, a red STBY flag appears on the dial face. The altimeter automatically switches to standby operation during an electrical power loss or when the altimeter or altitude computer fails. The standby operation is selected by placing the mode switch to STBY. An ac-powered internal vibrator 6-22

Figure 6-20 — AAU-24/A altimeter dial face. automatically energizes in the standby mode to lessen friction in the display mechanism. With the local barometric pressure set, the altimeter should agree to ±75 feet of field elevation in both modes. AAU-21/A ALTIMETER – AAU-21/A altimeter is used in low/slow aircraft. It has a counter-drum-pointer display similar in appearance to the AAU-19/A. The altimeter contains a servo-driven encoder. The encoder provides an altitude signal to the aircraft transponder for transmission to a ground station. AAU-24/A ALTIMETER – The AAU-24/A altimeter (Figure 6-20) contains a precision pressure sensing device, counter, and pointer drive mechanisms. It also contains a combination counter-drum and pointer for altitude display. The counter displays two digits, showing multiples of 10,000 feet and 1,000 feet respectively, and moves intermittently. The drum shows multiples of 100 feet and moves continuously. The pointer travels one revolution for each increment of 1,000 feet of altitude. The pointer scale is from 0 to 9, each step representing an increment of 100 feet. Each 100-foot step is split into two increments of 50 feet each.

The barometric setting (baroset) knob is located in the lower left corner of the bezel. It protrudes a maximum of 0.73 inch in front of the bezel. The baroset knob works with a four-digit counter, designated IN Hg, to set the altitude indication to the prevailing barometric pressure. It is adjustable from 28 to 31 inches of mercury. Next to the baroset knob is a locking screw. This screw is used only during calibration procedures to align the barometric pressure (IN Hg) indication with altitude indication. Two sets of internal lights, one red and one white, provide dial lighting. Each set consists of four lights. Controls for dial lighting are external to the altimeter. To overcome the effects of stop-and-jump friction in altimeter mechanisms, the altimeter has an internal, electrically operated mechanical vibrator. 6-23

Figure 6-21 — AOA indicators: (A) radial; (B) vertical scale. ANGLE-OF-ATTACK (AOA) INDICATING SYSTEM The Angle-of-Attack (AOA) indicating system detects aircraft angle of attack from a point on the side of the fuselage. It furnishes reference information for the control and actuation of other units and aircraft systems. It provides signals to operate an AOA indicator (Figure 6-21) on the pilot’s instrument panel. This indicator displays a continuous visual indication of the local angle of attack. A typical AOA system provides electrical signals for operating the rudder pedal shaker. The shaker warns the pilot of an impending stall when the aircraft is approaching the critical stall angle of attack. Electrical switches in the AOA indicator operating at various preset angles of attack energize colored lights in the approach light system and an approach index light in the cockpit. These lights furnish the landing signal officer and the pilot with an accurate indication of approach angle of attack during landing. An angle-of-sideslip system, consisting of an airstream direction detector, and angle-of-sideslip compensator, is installed on some aircraft. The outputs from these are used for controlled rocket firing.

The AOA indicating system consists of an airstream direction detector transmitter (Figure 6-22) and an indicator. The airstream direction detector measures local airflow direction relative to the true angle of attack. It does this by determining the angular difference between local airflow and the fuselage reference plane. The sensing element works with a servo-driven balanced bridge circuit, which converts probe positions into electrical signals.

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Figure 6-22 — AOA transmitter. Figure 6-23 — Mechanical schematic of airstream direction detector.

The AOA indicating system operation is based on detection of differential pressure at a point where the airstream is flowing in a direction that is not parallel to the true angle of attack of the aircraft. This differential pressure is caused by changes in airflow around the probe. The probe extends through the skin of the aircraft into the airstream. The exposed end of the probe contains two parallel slots (ports). These slots detect the differential airflow pressure (Figure 6-23). Air from the slots passes through two separate air passages to separate compartments in a paddle chamber. Any differential pressure, caused by misalignment of the probe to the direction of airflow, causes the paddles to rotate. The moving paddles rotate the probe, through mechanical linkage, until the pressure differential is zero. Alignment occurs when the slots are symmetrical with the airstream direction.

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Two potentiometer wipers, rotating with the probe, provide signals for remote indications. Probe position, or rotation, converts into an electrical signal by the po tentiometer that is the transmitter component of a self-balancing bridge circuit. When the angle of attack of the aircraft changes, the position of the transmitter potentiometer alters. The alteration causes an error voltage to exist between the transmitter potentiometer and the receiver potentiometer in the indicator. Current flows through a sensitive polarized relay to rotate a servomotor located in the indicator. The servomotor drives a receiver potentiometer in the direction required to reduce the error voltage. This action restores the circuit to a null or electrically balanced condition. The polarity of the error voltage determines the resultant direction of rotation of the servomotor. The indicating pointer is attached to, and moves with, the receiver potentiometer wiper arm to show on the dial the relative angle of attack. Figure 6-24 shows the relationship of the AOA indexer lights indication and stall warning. The AOA indexer lights mounted on the pilot’s Heads Up Display (HUD) Combiner Assembly has two arrows and a circle illuminated by colored lamps to provide the pilot with approach information. Two Angle of Attack Transmitters (AOATs) provide angle of attack information to the flight control computers, which in turn control the AOA indexer display. The upper arrow is for high angle of attack (green). The lower arrow is for low angle of attack (red). The circle is for optimum angle of attack (amber). An arrow and a circle together show an intermediate or optimum position for landing approach.

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Figure 6-24 — Angle-of-attack (AOA) indications.

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NOTE You should review NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187 before continuing. Figure 6-25 — Rudder shaker schematic (simplified). The indexer lights function only when the landing gear is down. A flasher unit causes the indexer lights to pulsate when the arresting hook is up with the HOOK BYPASS switch in the CARRIER position. STALL WARNING SYSTEM Many aircraft have stall warning indicators to warn the pilot of an impending aerodynamic stall. In the past, stall warning indicators were of a pneumatic control type. These devices activated either warning horns or flashing lights. Later, research found that a stall relates directly to the angle of attack, regardless of airspeed, power setting, or aircraft loading. The stall warning devices of most aircraft now in the fleet operate at a specified angle of attack. The devices operate through cams in the AOA indicator. The cam- driven switch activates a vibrator motor connected to either a rudder pedal or the control stick. Figure 6-25 shows a simplified schematic of the rudder shaker system. When the aircraft reaches stall angle of attack, the AOA indicator cam-actuated switch completes the rudder shaker motor circuit to ground. When the angle of attack returns below stall conditions, the cam de- actuates the switch. The switch action removes the ground from the rudder shaker motor. GYROSCOPIC INSTRUMENTS Early aircraft were flown by visually aligning the aircraft with the horizon. With poor visibility, it was not possible to fly the aircraft safely. The need for flight instruments to correct this condition led to the development of gyroscopic instruments. The gyroscopic properties of a spinning wheel made precision instrument flying, precise navigation, and pinpoint bombing practical and reliable. Some of the instruments that use this principle are the turn-and-bank indicator, directional gyro, gyro horizon, and drift meter. Systems that use the gyroscopic principle include the Automatic Flight Control System (AFCS), gyrostabilized flux-gate compass, and inertial navigation system. The following paragraphs contain a brief review of gyroscopic principles.

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Figure 6-26 — Simple gyroscope. A gyroscope is a spinning wheel or rotor with universal mounting. This mounting allows the gyroscope to assume any position in space. Any spinning object exhibits gyroscopic properties. The wheel, with specific design and mounts to use these properties, is a gyroscope. The two important design characteristics for instrument gyros are: 1. High-density weight for small size 2. High-speed rotation with low friction The mountings of the gyro wheels are gimbals. They can be circular rings or rectangular frames. However, some flight instruments use part of the instrument case itself as a gimbal. A simple gyroscope is shown in Figure 6-26. The two general types of mountings for gyros are the free or universal mounting and the restricted or semi-rigid mounting. The type of mounting the gyro uses depends on the gyro’s purpose. A gyro can have different degrees of freedom. The degree of freedom depends on the number of gimbals supporting the gyro and the arrangement of the gimbals. Do not confuse the term degrees of freedom, as used here, with an angular value as in degrees of a circle. The term degrees of freedom, as used with gyros, shows the number of directions in which the rotor is free to move. Some authorities consider the spin of the rotor as one degree of freedom, but most do not. A gyro enclosed in one gimbal, such as the one shown in Figure 6-26, has only one degree of freedom. This is a freedom of movement back and forth at a right angle to the axis of spin. When this gyro is mounted in an aircraft, with its spin axis parallel to the direction of travel and capable of swinging from left to right, it has one degree of freedom. The gyro has no other freedom of movement. Therefore, if the aircraft should nose up or down, the geometric plane containing the gyro spin axis would move exactly as the aircraft does in these directions. If the aircraft turns right or left, the gyro would not change position, since it has a degree of freedom in these directions. A gyro mounted in two gimbals normally has two degrees of freedom. Such a gyro can assume and maintain any attitude in space. For illustrative purposes, consider a rubber ball in a bucket of water. Even though the water is supporting the ball, it does not restrict the ball’s attitude. The ball can lie with its spin axis pointed in any direction. Such is the case with a two-degree-of-freedom gyro, often called a free gyro. In a two-degree-of-freedom gyro, the base surface turns around the outer gimbal axis or around the inner gimbal axis, while the gyro spin axis remains fixed. The gimbal system isolates the rotor from the base rotation. The universally mounted gyro is an example of 6-29

Figure 6-27 — Precession resulting from deflective force. this type. Restricted or semi-rigid mounted gyros are those mounted so one plane of freedom is fixed in relation to the base. Practical applications of the gyro are based upon two basic properties of gyroscopic action: 1. Rigidity in space 2. Precession Newton’s first law of motion states, “A body at rest will remain at rest, or if in motion will continue in motion in a straight line, unless acted upon by an outside force.” An example of this law is the rotor in a universally mounted gyro. When the wheel is spinning, it stays in its original plane of rotation regardless of how the base moves. The factors that determine how much rigidity a spinning wheel has are in Newton’s second law of motion. This law states, “The deflection of a moving body is directly proportional to the deflective force applied and is inversely proportional to its mass and speed.” To obtain as much rigidity as possible in the rotor, the rotor has great weight for siz e and rotates at high speeds. To keep the deflective force at a minimum, the rotor shaft mounts in low friction bearings. The basic flight instruments that use the gyroscopic property of rigidity are the gyro horizon, the directional gyro, and any gyrostabilized compass system. Therefore, their rotors must be freely or universally mounted. Precession (Figure 6-27) is the resultant action or deflection of a spinning wheel when a deflective force is applied to its rim. When a deflective force is applied to the rim of a rotating wheel, the resultant force is 90 degrees ahead of the direction of rotation and in the direction of the applied force. The rate at which the wheel precesses is inversely proportional to rotor speed and directly proportional to the deflective force. The force with which a wheel precesses is the same as the deflective force applied minus the friction in the gimbal ring, pivots, and bearings. If too great a deflective force is applied for the amount of rigidity in the wheel, the wheel precesses and topples over at the same time.

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CAUTION If pull-to-cage switch is locked in caged position, it must be pulled to extended position before rotating counterclockwise to un-caged position or damage will result. The pull-to-cage switch should not be locked in caged position if gyro is spinning or damage to gyro may result. Any spinning mass exhibits the gyroscopic properties of rigidity in space and precession. The rigidity of a spinning rotor is directly proportional to the weight and speed of the rotor, and inversely proportional to the deflective force. Attitude Indicator Pilots determine aircraft attitude by referring to the horizon when they can see it. Often, however, the horizon is not visible. When it is dark or when there are obstructions to visibility such as overcast skies, smoke, or dust, pilots cannot use the earth ’s horizon as a reference. When these conditions exist, they refer to an instrument called the att itude indicator. This instrument is also known as a Vertical Gyro Indicator (VGI), artificial horizon, Attitude Reference Indicator (ARI), or g yro horizon. From these instruments, pilots learn the relative position of the aircraft with reference to the earth’s horizon. Although attitude indicators (Figure 6-28, frame 1) differ in size and appearance, they all have the same basic components and present the same basic information. On the face of the indicator will always be a miniature aircraft that represents the nose (pitch) and wing (bank) attitude of the aircraft. The bank pointer on the indicator face shows the degree of bank (in 10-degree increments up to 30 degrees, then in 30-degree increments to 90 degrees). The sphere is always light on the upper half and dark on the lower half to show the difference between sky and ground. Calibration marks on the sphere show degrees of pitch in 5- or 10-degree increments. An OFF flag comes into view when the system has a loss of power or the pull-to-cage knob is pulled out (F igure 6-28, frame 3). Each indicator has a pitch trim adjustment or pull-to-cage knob for the pilot to center the horizon as necessary. When transporting the gyro, keep it in a locked and fixed position and use the pull-to-cage knob to protect the gimbals from damage. The knob must be pulled and turned clockwise.

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Figure 6-28 — Roll and pitch indications on the attitude indicator.

Some attitude indicators have a self-contained gyro. Other more modern indicators use pitch and roll information from the inertial system or the attitude heading reference system. These systems are accurate and reliable. They gain their reliability and accuracy from their larger size, which is not limited by the space of an instrument panel. Electrical signals from the remote gyro travel via synchros. The signal is amplified in the indicator to drive servomotors and position the indicator sphere. This positioning is the same as the vertical gyro position in the gyro case. In the newer attitude indicators, the sphere is gimbal-mounted and capable of 360-degree rotation. Also, a test function is provided to test the instrument landing system vertical and horizontal pointers. In contrast, the older gyros could only travel 60 degrees to 70 degrees of pitch and 100 degrees to 110 degrees of roll. Operation The attitude reference indicator receives 115vac 3-phase aircraft power through energized contacts of relay K1 located in the static power inverter. With 115vac 3-phase power applied to the attitude reference indicator, the OFF flag goes out of view and the gyro will spin up and erect. Also, 115vac phase A is applied to the dc power supply which develops dc voltages for the amplifiers and test circuits. The 115vac phase C is used to excite the attitude pick-off synchros. If 115vac 3-phase aircraft power is lost, 6-32

Figure 6-29 — All-attitude indicator (AAI). relay K1, in the static power inverter, de-energizes and 28-vdc is applied to a dc-to-ac inverter and develops the 115vac 3-phase power. An electrically driven vertical gyro (Figure 6-28, frames 5 through 10) maintains vertical orientation through use of an electronic erection system and provides a continuous attitude display. Attitude pick-off synchros are mechanically coupled to the gyro and their output signals are applied to pitch and roll amplifiers. Amplified pitch and roll analog signals are then sent to control-converter. In the control-converter, pitch and roll analog signals are applied through Scott-T transformers and an A/D converter to produce attitude signals. A software built in test BIT samples the attitude signals for reasonable content and a no-go produces not valid attitude pitch and roll output signals. The pitch and roll attitude and validity signals are then sent to the Mission Computer (MC) system and provide backup attitude signals for the various navigation routines and displays. Variations in aircraft angle-of-attack will cause differences in the caged position of the gyro spin axis relative to true vertical. Some aircraft incorporate an all-attitude indicator (Figure 6-29). In addition to pitch and roll, this indicator shows compass information along the horizon bar. It also shows turn- and-bank information on the bottom. An even more sophisticated instrument, the flight director, displays the above information plus radio navigation information, all on one instrument.

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NOTE Because of close association and interface with inertial navigation systems, Chapter 7 of this RTM contains information about attitude indicating systems. Figure 6-30 — Turn-and-bank indicator.

T urn-and-Bank Indicator The turn-and-bank indicator (Figure 6-30), also called the turn-and-slip indicator, shows the lateral attitude of an aircraft in straight flight. It also provides a reference for the proper executions of a coordinated bank and turn. It shows when the aircraft is flying on a straight course and the direction and rate of a turn. It was one of the first modern instruments for controlling an aircraft without visual reference to the ground or horizon. The indicator is a combination of two instruments, a ball and a turn pointer. The ball part of the instrument operates by natural forces (centrifugal and gravitational). The turn pointer depends on the gyroscopic property of precession for its indications. The power for the turn indicator gyro is either electrical or vacuum. BALL – The ball portion of a turn-and-bank indicator (Figure 6-30) consists of a sealed, curved, glass tube. The tube contains water-white kerosene and a black or white agate or common steel ball bearing. The ball bearing is free to move inside the tube. The fluid provides a damping action and ensures smooth and easy movement of the ball. The curved tube allows the ball to seek the lowest point when in level flight. This point is the tube center. A small projection on the left end of the tube contains a bubble of air. The bubble lets the fluid expand during changes in temperature. There are two markings or wires around the center of the glass tube. They serve as reference markers to show the correct position of the ball in the tube. The plate that holds the tube and the references are painted with luminous paint. The only force acting on the ball during straight flight (no turning) with the wings level is gravity. The ball seeks its lowest point and stays within the reference marks. In a turn, centrifugal force also acts on the ball in a horizontal plane opposite to the direction of the turn. The ball assumes a position between the reference markers when the resultant of centrifugal force and gravity acts directly opposite to a point midway between the 6-34

reference markers. When the force acting on the ball becomes unbalanced, the ball mo ves away from the center of the tube. In a skid, the rate of turn is too great for the angle of bank. The excessive centrifugal force moves the ball to the outside of the turn. The resultant of centrifugal force and gravity is not opposite the midpoint between the reference markers. The ball moves in the direction of the force, toward the outside of the turn. Returning the ball to center (coordinated turn) calls for increasing bank or decreasing rate of turn, or a combination of both. In a slip, the rate of turn is too slow for the angle of bank. The resultant of centrifugal force and gravity moves the ball to the inside of the turn. Returning the ball to the center (coordinated turn) requires decreasing the bank or increasing the rate of turn, or a combination of both. The ball instrument is actually a balance indicator because it shows the relationship between angle of bank and rate of turn. It lets the pilot know when the aircraft has the correct rate of turn for its angle of bank. TURN POINTER – The turn pointer operates on a gyro. The gimbal ring encircles the gyro in a horizontal plane and pivots fore and aft in the instrument case. The major parts of the turn portion of a turn-and-bank indicator are as follows:  A frame assembly used for assembling the instrument.  A motor assembly consisting basically of the stator, rotor, and motor bearings. The electrical motor serves as the gyro for the turn indicator.  A plate assembly for mounting the electrical receptacle, pivot assembly, choke coil, and capacitors for eliminating radio interference.  A damping unit that absorbs vibrations and prevents excessive oscillations of the needle. The unit consists of a piston and cylinder mechanism. The adjustment screw controls the amount of damping.  An indicating assembly composed of a dial and pointer.  The cover assembly.

The carefully balanced gyro rotates about the lateral axis of the aircraft in a frame that pivots about the longitudinal axis. When mounted in this way, the gyro responds only to motion around a vertical axis. It is unaffected by rolling or pitching. The turn indicator takes advantage of one of the basic principles of gyroscopes— precession. Precession, as already explained, is a gyroscope’s natural reaction 90 degrees in the direction of rotation from an applied force. It is visible as resistance of the spinning gyro to a change in direction when a force is applied. As a result, when the aircraft makes a turn, the gyro position remains constant. However, the frame in which the gyro hangs, dips to the side opposite the direction of turn. Because of the design of the linkage between the gyro frame and the pointer, the pointer shows the correct direction of turn. The pointer displacement is proportional to the aircraft rate of turn. If the pointer remains on center, it shows the aircraft is flying straight. If it moves off center, it shows the aircraft is turning in the direction of the pointer deflection. The turn needle shows the rate (number of degrees per minute) at which the aircraft is turning. By using the turn-and-bank indicator, the pilot checks for coordination and balance in straight flight and in turns. By cross-checking this instrument against the airspeed 6-35

Figure 6-31 — Miscellaneous flight instruments. indicator, the pilot can determine the relation between the aircraft lateral axis and the horizon. For any given airspeed, there is a definite angle of bank necessary to maintain a coordinated turn at a given rate. MISCELLANEOUS FLIGHT INSTRUMENTS The pilot uses several other indicators to control the aircraft. These indicators are not always useful, but they are beneficial under special flight conditions. As you read this section, refer to Figure 6-31.

Accelerometer Indicators The pilot must limit aircraft maneuvers so various combinations of acceleration, airspeed, gross weight, and altitude remain within specified values. These operational limits cut out the possibility of damaging aircraft as a result of excessive stresses. The accelerometer shows the load on the aircraft structure in terms of gravitation (g) units. It presents information that lets the aircraft be maneuvered within its operational limits. The forces sensed by the accelerometer act along the vertical axis of the aircraft. The main hand moves clockwise as the aircraft accelerates upward and counterclockwise as the aircraft accelerates downward. 6-36

Figure 6-32 — Accelerometer mechanical schematic. The accelerometer indications are in g units. The main indicating hand turns to +1 g when the lift of the aircraft wing equals the weight of the aircraft. Such a condition prevails in level flight. The hand turns to +3 g when the lift is three times the weight. The hand turns to minus readings when the forces acting on the aircraft surfaces cause the aircraft to accelerate downward. The accelerometer operates independently of all other aircraft instruments and installations. The activating element of the mechanism is a mass that is movable in a vertical direction on a pair of shafts (Figure 6-32). A spiral-wound main spring dampens the vertical movement of the mass. The force of the mass travels by a string-and-pulley system to the main spring and main shaft. From here, it goes to the plus and minus assemblies. The hand assemblies mount on the plus and minus assemblies. Changes in vertical acceleration cause movement of the mass on the shafts, which translates into a turning motion of the main shaft. The turning motion pivots the indicating hands around the dial. The hand travels a distance equivalent to the value, in g units, of the upward or downward acceleration of the aircraft.

The accelerometer operates on the principle of Newton’s third Law of Motion. During level flight, no forces act to displace the mass from a position midway from the top and bottom of the shafts. Therefore, the accelerometer pulley system performs no work, and the indicating hands remain stationary at +1 g. When the aircraft changes from level 6-37

flight, forces act on the mass. This action causes the mass to move either above or below its midway position. These movements cause the accelerometer indicating hands to change position. When the aircraft goes nose down, the hands move to the minus section of the dial. When the nose goes up, they move to the plus section. The main hand continuously shows changes in loading. The two other hands on the accelerometer show the highest plus acceleration and highest minus acceleration of the aircraft during any maneuver. The indicator uses a ratchet mechanism to maintain these readings. A knob in the lower left of the instrument face is used to reset the maximum- and minimum-reading hands to normal. Thus, the accelerometer keeps an indication of the highest accelerations during a particular flight phase or during a series of flights. Clocks The standard Navy clock is a 12-hour, elapsed-time, stem-wound clock with an 8-day movement. This type of clock is in the cockpit for use by the pilot or copilot. Clocks may be located elsewhere for use by other crewmembers as well. The pull-to-set winding stem is at the lower left of the dial. The dial has 60 divisions, which you read as minutes or seconds, as appropriate. The face has standard minute and hour hands, a sweep- second hand, and an elapsed-time minute hand. You may start, stop, or reset the elapsed-time minute hand by pressing a single button at the upper right of the dial. Direct-Reading Magnetic Compasses During the early days of aviation, direction of flight was determined chiefly by direct- reading magnetic compasses. Today, the direct-reading magnetic compass (Figure 6- 31) is used as a standby compass. Direct-reading magnetic compasses used in Navy aircraft mount on or near the instrument panel for use by the pilot. They are read like the dial of a gauge. A nonmagnetic metal bowl, filled with liquid, contains the compass indicating card. The card provides the means of reading compass indications. The card mounts on a float assembly and is actually a disk with numbers painted on its edge. A set of small magnetized bars or needles fasten to this card. The card-magnet assembly sits on a jeweled pivot, which lets the magnets align themselves freely with the north-south component of the earth’s magnetic field. The compass card and a fixed-position reference marker (lubber’s line) are visible through a glass window on the side of the bowl. An expansion chamber in the compass provides for expansion and contraction of the liquid caused by altitude and temperature changes. The liquid dampens, or slows down, the oscillation of the card. Aircraft vibration and changes in heading cause oscillation. If suspended in air, the card would keep swinging back and forth and be difficult to read. The liquid also buoys up the float assembly, reducing the weight and friction on the pivot bearing. Instrument-panel compasses for naval aircraft are available with cards marked in steps of either 2 degrees or 5 degrees. Such a compass indicates continuously without electrical or information inputs. You can read the aircraft heading by looking at the card in reference to the lubber line through the bowl window. Standby Attitude Indicator The standby attitude indicator (Figure 6-31) on the pilot instrument panel consists of a miniature aircraft symbol, a bank angle dial, and a bank index. It also includes a two- colored drum background with a horizon line dividing the two. 6-38

The indicator roll index is graduated in 10-degree increments to 30 degrees, with graduation marks at 60 degrees and 90 degrees. The indicator is capable of displaying 360 degrees of roll, 92 degrees of climb, and 79 degrees of dive. Because of the high spin rate of the gyro, the indicator displays accurate pitch-and-roll data for 9 minutes after electrical failure. The attitude indicator incorporates a pitch trim knob to position the miniature aircraft symbol above or below the horizon reference line. The pitch trim knob also cages the gyro. When the pitch trim knob is pulled out, the gyro will cage. Rotating the knob clockwise while extended will cause the gyro to lock, in the extended position. The attitude indicator also incorporates an OFF flag. The flag appears if electrical power fails, or if you cage the gyro. Outside Air Temperature Indicator An indicator displaying uncorrected outside air temperature is located on the pilot’s instrument panel (Figure 6-31). A temperature-sensitive resistor (temperature bulb) is exposed to the slipstream. This resistor measures changes in temperature. The temperature of the air measurement is in the form of changing resistance. The outside air temperature indicator displays this change in resistance. The graduated indicator dial is marked in Celsius, from +50 degrees to –50 degrees. ENGINE INSTRUMENT SYSTEMS Engine instruments provide indications of tail pipe temperature, oil and fuel pressure, engine RPM, oil temperature, and fuel flow rate. The pilot must be aware of engine operation at all times. If oil pressure falls below the normal operating limit or tail pipe temperature becomes excessively high, the engine instruments provide these indications to the pilot.

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Figure 6-33 — Tachometer indicators: (A) jet engine (radial); (B) jet engine (vertical scale). TACHOMETER SYSTEMS The tachometer indicator is an instrument that shows the speed of a gas turbine engine (jet) main rotor assembly. Figure 6-33 shows tachometer indicators for various types of engines. The dials of tachometer indicators used with jet engines are shown in percentage of Revolutions per Minute (RPM), based on takeoff RPM.

Several types and sizes of generators and indicators are used in the tachometer systems of naval aircraft. As a rule, they all operate on the same basic principle. This section introduces you to information on tachometer systems. A typical generator and a typical indicator are described because it is not practical to describe all the generators and indicators. For detailed information on a particular system, you should refer to the manufacturer’s manuals. Essentially, the tachometer system consists of an ac generator coupled to the aircraft engine and an indicator consisting of a magnetic-drag element on the instrument panel. The generator transmits electric power to a synchronous motor, a part of the indicator. The frequency of this power is proportional to the engine speed. An accurate indication of engine speed is obtained by applying the magnetic-drag principle to the indicating element. The problem of changes in generator output voltage is cut out by the generator and synchronous-motor combination. These units make a frequency-sensitive system for sending an indication of engine speed to the indicator with absolute accuracy. For many installations, it is desirable to send a single engine-speed indication to two different stations in the aircraft. The frequency-sensitive system is ideal for this application because there is no change in indication when a second indicator connects in parallel with the first. Synchronous motor operation in each indicator depends only on the availability of enough power in the generator to operate both indicator motors. 6-40

Figure 6-34 — Cutaway view of a tachometer generator. Tachometer Generator Tachometer generator units are small and compact (about 4 inches by 6 inches). The generator is constructed with an end shield designed so the generator can attach to a flat plate on the engine frame or reduction gearbox, with four bolts. Figure 6-34 shows a cutaway view of a tachometer generator. You should refer to it while you read this section. The generator consists essentially of a permanent magnet rotor (callout 1) and a stator (callout 8) that develop three-phase power as the rotor turns.

The armature of the generator consists of a magnetized rotor. The rotor is cast directly onto the generator shaft. The generator may be of either two- or four-pole construction. The two- and four-pole rotors are identical in appearance and construction. They differ in that the two-pole rotor is magnetized north and south diametrically across the rotor, while the four-pole rotor is magnetized alternately north and south at each of the four pole faces. The key (callout 2) that drives the rotor is a long, slender shaft. It has enough flexibility to prevent failure under the torsional oscillations originating in the aircraft drive shaft. It will also accommodate small misalignments between the generator and its mounting surfaces. This key goes into the hollow rotor shaft. A pin (callout 3) at the end opposite 6-41

Figure 6-35 — Cutaway view of a tachometer indicator (radial). the drive shaft secures the key in place. An oil-seal ring (callout 4) is located inside the hollow shaft and over this key. This seal prevents oil from leaking into the generator through the hollow shaft. The shaft runs in two ball bearings (callout 5) set in stainless steel inserts. The inserts are cast directly into the generator end shields (callout 6). An adjusting spring (callout 7) at the receptacle end of the shaft maintains the proper amount of end play. The stator consists of a steel ring with a laminated core of ferromagnetic material. A three-phase winding goes around this core and is insulated from it. The winding is adapted for two- or four-pole construction, depending on the generator in which it is used. The two end shields are made of die-cast aluminum alloy. They serve to support the generator stator and rotor by means of a receptacle (callout 9). The receptacle attaches to the junction box (callout 10) of the generator. Tachometer Indicators Tachometer indicators mount on the cockpit instrument panel. They are relatively small in size. The type of unit varies. Depending on the particular installation, some are single element and others are dual element. The operating principles of the two types are basically the same. Figure 6-35 shows a cutaway view of a single element tachometer indicator (radial). The unit consists of two parts, a synchronous motor and an indicating element. The motor runs in synchronism with the tachometer generator. It also drives the indicating element through a magnetic-drag coupling. The indicating element indicates the speed of the synchronous motor, and, therefore, the speed of the aircraft engine.

The synchronous motor (callout 3) consists of a three-phase stator winding that goes in, and is insulated from, a laminated circular core. Within the circular core is a shaft. The rotating parts attach to this shaft. A cotter pin secures a hysteresis disk (callout 1) to the 6-42

shaft. A permanent magnet rotor (callout 2) is free to move on the shaft. The hysteresis disk at one shaft end and a spring at the other restrain longitudinal motion of the permanent magnet rotor. The spring is secured to the shaft to transmit torque from the rotor to the shaft. Ball bearings in the motor end shields support the shaft. These end shields also serve to locate the stator. This combination secures all parts of the motor and maintains their proper position with respect to each other. The armature of the synchronous motor consists mainly of the permanent magnet and the hysteresis disk. The purpose of the permanent magnet material is to provide starting and running torque at low speeds. The hysteresis disk provides starting torque at high speed. High speeds are necessary because the magnitude of flux is great, but the permanent magnet, by itself, cannot pull into step. At higher speeds, the hysteresis disk moves the rotor up to near synchronism, and then the permanent magnet pulls it into exact synchronism. One end of the motor shaft extends through the front end shield and supports the drag- magnet assembly (callout 9). The drag-magnet assembly, which is driven by the synchronous motor, consists of two plates to which small permanent magnets attach. The arrangement of the magnets concentrates the flux near the outside edge of the drag disk. This arrangement obtains maximum torque with minimum weight. Between the two plates, carrying the magnets is a drag disk (callout 4) of conducting material. This material is an alloy with a low-temperature coefficient, which prevents temperature changes from affecting the material’s resistance. The magnet assembly spinning around the disk of conducting material produces torque on the disk. The drag disk connects to the lower end of the indicator assembly shaft. When the disk rotates, the indicator pointer moves to show the speed of the aircraft engine. The indicating element is supported by three posts. These posts have adjustable nuts (callout 8) for leveling the assembly as necessary. You can obtain further positioning by moving the adjusting arm (callout 7). The scale plate (callout 5) is calibrated in either RPM or percentage and shows engine speed. The cover assembly (callout 6) serves as a protective container for the mechanism. The receptacle (callout 10) at the rear of the indicator provides electrical connection to the tachometer generator. Dual Indicators With the increasing requirement for more instruments for efficient flight, the combination of several instruments in one has become very common. The dual tachometer is an example of a combination of instruments. Some multi-engine aircraft use dual tachometers. The dual tachometer consists of two synchronous-motor, magnetic-drag tachometer indicat ors housed in a single case. The indicators show the speed of rotation of the engines simultaneously on a single dial. There is one tachometer indicator for each pair of engines on the aircraft. Vertical Scale Indicators Vertical scale indicators are used on some models of naval aircraft. A vertical scale shows engine performance data such as fuel flow, engine speed, exhaust gas temperature, and accelerometer readings. Vertical scale indicators are compact, light in weight, and easily read. 6-43

Figure 6-36 — Basic engine instrument vertical scale indicators. All vertical scale indicators (Figure 6-36), consist of a vertical tape that operates by an amplifier, motor, gears, and sprockets. These systems are the same as systems used on other aircraft. Vertical scale indicators may be utilized individually or in clusters and may be analog or digital based displays depending upon the Type/Model/Series aircraft. The engine indicating groups consist of cockpit indicators and associated sensing devices required to monitor left and right engine performance. Dual indicators display percentage of engine compressor rotor rpm (RPM indicator), turbine inlet temperature (TIT indicator), and engine fuel flow (FF indicator).

TACHOMETER INDICATOR – The electrical tachometer (RPM) indicator displays percentage of engine rotor speed on two vertical scales (one each for the left and right engines). The indicator scales are linear from 0 to 6, and from 6 to 11 multiplied by 10 to get percent of RPM. Upper left and right limit range markers and OFF failure flags appear between the 10.4 and 11 points of the scales. The absence of the OFF failure flags confirms the indicator channels are receiving power. The indicator receives variable frequency signals proportional to compressor speed from each engine tachometer generator. The signals go to solid-state circuitry to produce a proportional drive signal for a servomotor. The servomotor drives gears and sprockets to position a tape on the indicator face, showing percentage of engine rotational speed.

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NOTE If there is an open in the voltage supply circuit, the galvanometers will also read zero. Figure 6-37 — Wheatstone bridge thermometer. TACHOMETER GENERATOR – The compressor rpm tachometer generator is a two-phase alternator type generator. The generator supplies electrical signals directly propo rtional to engine-compressor rotation speed. The tachometer generator is driven by the high-pressure compressor through the engine accessory gearbox. Signals from this generator go directly to the electrical tachometer indicator (RPM indicator) at the crew station. The indicator displays percentage of rotor speed. TEMPERATURE INDICATING SYSTEMS To properly monitor the operation of an aircraft engine, you must know various temperature indications. Some of the more important indications include the temperatures of the engine oil, free air, and exhaust systems of jet engines. Various types of thermometers, such as the bimetal and resistance types, collect and present this information. The main parts of resistance thermometers are the indicating instrument, the temperature sensitive element (resistance bulb), and the connecting wires leading from the bulb. Wheatstone Bridge System A schematic diagram of a Wheatstone bridge thermometer circuit is shown in Figure 6- 37. You should refer to it as you read this section. The resistance bulb element is one side of the Wheatstone bridge circuit. The other three sides are resistors in the indicating meter. The circuit receives voltage from the aircraft dc power supply. When the temperature bulb senses a temperature of 0 °C, its resistance is 100 ohms. The resistance of arms X, Y, and Z are also 100 ohms each. At this temperature the Wheatstone bridge is in balance. This means the sum resistance of X and Y equals the sum resistance of the bulb and Z. Therefore, the same amount of current flows in both sides of this parallel circuit. Since all four sides are equal in resistance, the voltage drop across side X equals the drop across the bulb. Since these voltages are equal, the voltage from A to B is zero, and the indicator reads zero.

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Figure 6-38 — Radiometer type temperature indicator. When the temperature of the bulb increases, its resistance also increases. This unbalances the bridge circuit causing the needle to deflect to the right. When the temperature of the bulb decreases, its resistance decreases. Again, the bridge circuit goes out of balance. However, this time the needle swings to the left. The galvanometer is calibrated so the amount of deflection causes the needle to point to the number of the meter scale. This number corresponds to the temperature at the location of the resistance bulb. This instrument requires a constant and steady supply of dc voltage. Fluctuations in the power supply affect total bridge current, which can cause an unbalanced bridge. Unless excessive heat damages the bulb, it will give accurate service indefinitely. When a thermometer does not operate properly, check carefully for loose wiring connections before replacing the bulb. Radiometer System The radiometer is a temperature indicator that uses two coils in a balanced circuit. In some instruments, the coils turn between the poles of a permanent magnet. In other instruments, a small permanent magnet rotor turns between stationary coils. Radiometer circuits vary in design, but the principle of operation is very much the same for all. Figure 6-38 shows a simplified circuit with a permanent magnet rotor. The two coils are stationary in the instrument, and the indicator needle fastens to the permanent magnet rotor. The needle position is determined by how the permanent magnet aligns itself with the resultant flux of the two coils. For an understanding of how the circuit operates, let’s trace the current through the circuit. Starting at ground, current flows up through the bulb, centering potentiometer R5 and R6, to point D. Current through the left leg of the bridge is from ground through R1 to point A. Current then goes from point A through the lower part of the expansion and contraction potentiometer R2. It also goes from pin 2 of R2 through R4 to point D. Here, the currents of the two legs combine and flow through R7 to the positive 28 volts. Note that restoring coil L2, resistor R3, and upper part of potentiometer R2 forms a parallel path for current flow from point A to pin 2 of R2. Deflection coil L1 connects between points A and B. Therefore, any difference in potential between these two points will cause current to flow through L1. 6-46

Figure 6-39 — Thermocouples: (A) gasket type; (B) rivet type. The radiometer temperature indicator uses a fixed permanent magnet to pull the pointer to an off position when the indicator is not operating. Thus, current through restoring coil L2 must compensate for the pull off magnet when the indicator is operating. Variations in the resistance of the bulb, because of temperature changes, will cause a change in volta ge at point B. Variations also causes the resulting change in current through deflection coil L1. Thermocouple System Thermocouple temperature indicators show the air temperatures in the heater duct of anti-icing systems and in the exhaust systems of jet engines. A thermocouple is a junction or connection of two unlike metals; such a circuit has two junctions. When one of the junctions becomes hotter than the other, an electromotive force is produced in the circuit. By including a galvanometer in the circuit, this electromotive force can be measured. The hotter the high temperature junction (hot junction) becomes, the greater the electromotive force. By calibrating the galvanometer’s dial, in degrees of temperature, the galvanometer becomes a thermometer. The galvanometer contains the cold junction. The thermocouple thermometer systems used in naval aircraft consist of a galvanometers indicator, a thermocouple or thermocouples, and thermocouple leads. Some thermocouples consist of a strip of copper and a strip of constantan pressed tightly together. Constantan is an alloy of copper and nickel. Other thermocouples consist of a strip of iron and a strip of constantan. Others may consist of a strip of Chromel and a strip of Alumel. The hot junction of the thermocouple varies in shape, depending on its application. Two common types, gasket and rivet, are shown in Figure 6-39. In the gasket thermocouple, the rings of two dissimilar metals are pressed together, forming a spark plug gasket. Each lead that connects back to the galvanometers must be of the same metal as the thermocouple part to which it connects. For example, a copper wire connects to the copper ring, and a constantan wire connects to the constantan ring. Thermocouple leads are critical in makeup and length because the galvanometers are calibrated for a specific set of leads in the circuits.

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Figure 6-40 — Turbine inlet temperature indicator system. TURBINE INLET TEMPERATURE INDICATOR SYSTEM – Some aircraft have a Turbine Inlet Temperature (TIT) Indicator System (Figure 6-40) to provide a visual indication of temperatures entering the turbine. The temperature of each engine turbine inlet is measured by 18 dual-unit thermocouples in the turbine inlet casing. These dual thermocouples are connected in parallel. One set sends signals through a harness and aircraft wiring to an indicator. The other set of thermocouples provides signals to the temperature datum control. Each circuit is electrically independent and provides dual system dependability.

All parts of the engine temperature measurement system, including welds, are made of Chromel and Alumel material. Special wiring and wire identification are in the aircraft from the thermocouple harness terminal block to the indicator. Plugs in the thermocouple circuits are also of a special type. The thermocouple harness mounts on the turbine unit aft of the thermocouple. The harness includes separate leads for each of the 18 thermocouples, and it maintains two electrically separate circuits. The harness is located inside a rigid metal, channel type of housing and cover. The leads and terminals project through holes in the front side of the housing wall. Electrical signals from the 18 dual-junction thermocouples are averaged within the harness. The thermocouple assemblies mount on pads provided around the turbine inlet case. Each thermocouple incorporates two electrically independent junctions within a 6-48

sampling-type probe. AL identifies Alumel terminal studs, and CR identifies Chromel terminal studs. Since the average voltage of the thermocouples at the thermocouple terminal blocks represents the turbine inlet temperature, it is necessary that no interference with the signal take place while the signal goes to the indicator. Therefore, the wiring from the thermocouple terminal block to the indicator goes through the harness. The harness wiring goes separately from other interference producing wiring. The indicator contains a bridge circuit with cold junction compensation, a two-phase motor to drive the pointer, and a feedback potentiometer. Also included in the indicator are the Zener voltage reference circuit, a chopper circuit, an amplifier, a power supply, a power-off flag, and an over-temp warning light. Output of the bridge circuit goes to the chopper circuit, so the bridge circuit is not loaded. The chopper output goes to the amplifier. Output of the amplifier feeds the variable field of a two-phase motor. This field positions the indicator main pointer and the digital indicator. The motor also drives the feedback potentiometer to provide a nulling signal. The signal is relative to the temperature signal and stops the drive motor upon reaching the correct pointer position. The Zener diode circuit provides a closely regulated reference voltage in the bridge. This signal avoids the error caused by voltage variation from the indicator power supply. The indicator power supply powers the Zener circuit, the chopper, and the amplifier. It also powers the power-off warning flag and the fixed field of the two-phase motor. The over-temperature warning light in the indicator comes on when the TIT reaches 1,082 °C. At this point, a switch in the indicator closes to energize the warning light. One test switch installed external to the indicators lets the crew test all the indicator over- temperature warning lights at once. The test switch simulates an over-temperature signal in each indicator’s temperature control bridge circuit. When power to an indicator fails, a red warning flag becomes visible. Also, the indicator pointers maintain their position, and the over-temperature warning light becomes inoperative. The indicator scale is calibrated in degrees Centigrade from 0 to 12 (times 100 °C). The digital indicator goes from 0 °C to 1,200 °C in 2-degree increments. The aircrafts engine also uses the thermocouple principle for indicating engine turbine inlet temperatures. Each engine has 10 thermocouple probes, distributed at three stations on the engine. They measure and average engine turbine inlet temperature. There are three types of thermocouple probes, compressor inlet temperature , compressor discharge temperature , and exhaust gas temperature thermocouple pressure ( – ). Each thermocouple probe has one or more Alumel and Chromel junctions. When the junctions become hot, a reaction between the dissimilar metal generates a dc voltage. A thermocouple harness connects the thermocouples in parallel to provide an average heat signal from each station. The thermocouple temperature indicator displays turbine inlet gas temperature on two vertical scales (Figure 6-36), one for each engine. The scales are linear from 0 to 6, segmented in tens from 6 to 14, and multiplied by 100 °C when read. OFF failure flags appear at the upper left and right of the indicator to show loss of signal input or electrical power. Internally, the indicator has two channels, one for each engine. The channels consist of a cold junction compensator, rebalance potentiometer, chopper, servo amplifier, 6-49

Figure 6-41 — Exhaust gas temperature indicating system. servomotor, and gear train. Thermocouple signal voltage from the engines goes to the cold junction compensator in each channel. The compensator provides corrective voltages to counteract the effect of secondary thermocouple junctions in the indicator when Alumel and Chromel leads connect to copper ones. A stable voltage goes to the old junction compensator and rebalance potentiometer. The feedback of the potentiometer and output of the compensator go to the chopper, where it compares the inputs. The chopper provides a 400-Hertz error signal to the servo system. The chopper output (signals relative to temperature change and potentiometer versus compensator difference) goes to a servo amplifier. The servo amplifier modifies the signals to drive the servomotor. The shaft of the motor couples to the rebalance potentiometer and indicator tape through the gear train. As the amplifier error drives the motor, the rebalance potentiometer goes in a direction that reduces the error signal, nullifying the condition. The tape shows temperature, on the front scale of the indicator, relative to thermocouple output. When supplied with 28 volts dc, a test circuit in the indicator energizes a relay, disconnecting the thermocouple input. Then, it substitutes a test signal of specific value to be processed and to drive the indicator tape. EXHAUST GAS TEMPERATURE INDICATING SYSTEM – The Exhaust Gas Temperature (EGT) indicating systems provide a visual temperature indication in the cockpit of the engine exhaust gases. The following is a discussion of a typical EGT indicating system. The aircraft contains two separate but identical EGT indicating systems (Figure 6-41), one for each engine. Each system has 12 dual thermocouples, a combination indicator and transistorized amplifier, and the interconnecting Chromel and Alumel leads. Power for the indicator-amplifier is from the essential 115-volt ac bus.

Both exhaust gas temperature indicators are on the pilot’s main instrument panel. They provide a visual indication of the engine exhaust temperatures. Each instrument is a 6-50

hermetically sealed unit with a single receptacle for a mating plug electrical connection. The instrument scale ranges from 0 °C to 1,200 °C. There is a vernier dial in the upper right corner of the instrument face. A power-off warning flag is in the lower portion of the dial. Internally, the indicator contains a simulated thermocouple cold junction with compensating resistors, a reference voltage source, and a dc-to-ac modulator. It also contains a transistor power output stage, miniature ac servomotor, and the power-off warning flag. The temperature indicator contains range markings on the instrument face. The thermocouples convert engine exhaust gas temperature into millivolts. The voltage from the thermocouples goes directly to the indicator amplifier through the Chromel and Alumel leads. The voltage is amplified and drives a small servomotor. The motor, in turn, drives the indicator pointer. The thermocouple harness consists of two halves, each containing six dual-loop thermocouples. The assembled halves make up two independent thermocouple systems, each consisting of 12 thermocouples connected in parallel. The harness mounts on the turbine frame aft of the turbine rotor. FUEL FLOW SYSTEMS Fuel flow indicating systems provide a continuous indication of the rate of fuel delivery to the engine. The rate of flow is in pounds per hour. In some systems, the indicator also shows the amount of fuel remaining in the tanks. A typical flow meter consists of two units, a transmitter and an indicator. The measurements are transmitted electrically to the panel-mounted indicator. Thus, use of electrical transmission ends the need for a direct fuel-filled line from the engine to the instrument panel. Removing the fuel line minimizes the chance of fire and reduces mechanical failure rate. The fuel flow meter system is quite similar to other synchro systems discussed in Navy Electricity and Electronic Training Series (NEETS) Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187. The following discussion describes a typical fuel flow indicating system to acquaint you with flow meters in general. However, you should always refer to the manuals for the particular system you are maintaining. Fuel Flow Transmitter Figure 6-42 shows a cutaway view of a fuel flow transmitter. It is a two-in-one unit, a fuel -measuring mechanism (or meter) and a synchro transmitter. You can separate these parts from one another for maintenance purposes, but they join as a single assembly for installation. The fuel enters the inlet port of the transmitter and flows against the vane (callout 1), causing the vane to swing. The spiral fuel chamber design allows the distance between the vane and chamber wall to become increasingly larger as fuel flow increases. A calibrated hairspring (callout 2) retards the motion of the vane. The vane ceases motion when the forces exerted on it by the hairspring and by the fuel are equal.

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Figure 6-42 — Cutaway view of a fuel flow transmitter. The rotor shaft of the synchro transmitter (callout 3) connects to a bar magnet (callout 4). Attached to the vane shaft is a ring magnet (callout 5). The ring magnet moves as the vane shaft moves. The transmitter mounting frame is between the bar magnet and the ring magnet, forming a liquid-tight seal. This is the seal between the fuel-metering section of the mechanism and the synchro. However, the bar magnet moves in unison with the ring magnet because the two magnets are magnetically coupled. The south pole of the ring magnet is opposite the north pole of the bar magnet. The two magnets send vane movement, caused by the fuel flow, to the synchro rotor. This action results in a corresponding movement of the rotor. Therefore, the angular displacement of the vane in relation to the fuel chamber housing determines the synchro rotor movement with respect to the stator. The fuel flow transmitter has a relief valve, which automatically opens and bypasses the instrument when the fuel flow exceeds the capacity of the instrument. At such time, only part of the fuel flows through the metering portion. As the pressure across the instrument falls below the value at which the relief valve opens, the valve closes. This lets the flowmeter again operate normally. The transmitter unit location is in the fuel line between the fuel pump and fuel nozzle. Fuel Flow Indicator The fuel flow meter indicator is located on the instrument panel. It is a remote-indicating instrument. This indicator consists of a synchro receiver, a step-up gear train, a magnetic drag cup, and a calibrated spring. When fuel flows through the fuel flow transmitter, an electrical signal goes to the indicator receiver. This signal drives the synchro rotor to the proper position. Thus, the indicator pointer shows the rate of fuel flow. 6-52

Figure 6-43 — (A) Fuel flow indicator; (B) fuel flow totalizer indicator. Figure 6-44 — Single fuel flow indicating system. Figure 6-43, view A, shows the face of the single flow indicator. To determine the amount of fuel consummation per hour, multiply the scale reading by 1,000. Figure 6-44 shows a schematic diagram of the single fuel flow indicator.

The following discussion uses the basic indicating system as an example of a fuel flow system. Other aircraft fuel flow systems operate in a similar way. The fuel flow transmitter consists of a synchronous motor, drum assembly, impeller assembly, spiral spring, and pickup coils. The transmitter housing has fuel inlet and outlet attachment flanges. The drum and impeller assemblies have two miniature 6-53

permanent magnets, 180 degrees apart. The motor runs at a constant 120 RPM. The motor connects through a shaft to the drum assembly. The impeller assembly rotates over the motor drum shaft and mechanically couples to the drum with the spiral spring. The pickup coils, one for each assembly, are in line with the drum and impeller assembly magnets. As the motor rotates the drum, the impeller also rotates. When there is no fuel flow, the magnets of the drum and impeller assemblies align. As they pass their respective coil, they generate simultaneous output signals. As fuel flow starts and increases through the transmitter housing, it goes through straightening vanes. These vanes eliminate the swirling motion of fuel. The fuel then passes through straight drilled passages of the rotating impeller. As fuel flow increases through the impeller, a proportional drag factor, or resistance to rotation, is imposed on the impeller assembly. This resistance causes the spring to deflect, equalizing the loading. The impeller magnets then deflect out of alignment. This action produces a later signal than that of the drum magnets and coil. Thus, an increased time span between signals of the rotating assemblies becomes relative to increased fuel flow. The fuel rate-of-flow power supply consists of a power transformer and power supply, two signal-conditioning channels, and a motor driver. The transformer receives 115 volts ac, 400 Hertz, and it feeds the power supply. The power supply provides low dc vo ltage to operate the motor-driven logic and signal-conditioning channels. Using a stepping signal to drive control logic, the motor driver controls positive and negative 8- Hert z ac signals between phases of both fuel rate-of-flow transmitter motors. The signal-conditioning channel for each engine system receives pulses from the coils of its transmitter. For each channel, a pulse shaper converts the time between transmitter drum and impeller coil pulses into a rectangular pulse width signal. An averaging filter processes this converted signal, which provides a low-ripple dc signal input to the fuel rate-of-flow indicator. The size (0 to 5 volts dc) of this signal is proportional to flow rate. A test circuit permits testing the power supply. A tap-off motor, phases A and B from the motor driver, routes an 8-Hertz signal through an external test switch to the signal-conditioning channels for processing. The results of the processed signal are displayed on the indicator. The fuel rate-of-flow indicator, a vertical scale indicator, displays rate of fuel flow for each engine on parallel scales. The scales are from 0 to 13. The scale reading, multiplied by 1,000, shows the rate (pounds per hour) at which the engine is consuming fue l. The upper left and right OFF failure flags show a loss of power, or signal to the indicator. The indicator has two separate channels, one for each engine. The channels include a control transformer servo amplifier, servomotor, gears, and sprockets. The indicator channels receive 115 volts of ac for signal processing. An input of 0 to 5 volts dc from the fuel rate-of-flow power supply produces an output. This output is from the control transformer rotor winding to the servo amplifier. The servo amplifier modifies the signal to the proper impedance and power level to drive the channel servomotor. Shaft rotation controls transformer output, and the rotation reduces transformer output voltage. When the output is null, the motor, gear train, and sprockets come to rest at a rate equivalent to the input. The test selector switch on the MASTER TEST panel tests the indicators. The self-test circuit in the indicators disconnects the fuel rate-of-f low transmitter input circuits. It then connects to an appropriate test signal within the two indicator channels. The channel circuitry processes the test signal and drives the indicator tapes to indicate 4,200 to 4,400 pounds per hour. 6-54

Figure 6-45 — Engine gauge unit. Fuel Flow Totalizing Systems Figure 6-43, view B, shows the indicator of a fuel flow totalizing system. The pointer of this instrument usually shows the combined rate of fuel flow into two or more engines. Also, if only one engine is operating, the pointer gives a true indication. A continuous reading of the pounds of fuel remaining in the aircraft fuel cells appears in the small window. Before starting the engines, you set the total amount of fuel in the aircraft on the pounds-fuel-remaining indicator by using the reset knob on the front of the instrument. As soon as the engines are running, the fuel flow pointer shows the rate of fuel consummation. The fuel-remaining indicator starts counting toward zero, giving a continuous reading of fuel remaining in the cells. Numbers rotate past the window like those of the mileage indicator of an automobile speedometer. The entire fuel flow totalizing system consists of two or more fuel flow transmitters, an amplifier, and an indicator. FUEL FLOW TRANSMITTERS – The fuel flow transmitters are almost identical to those already discussed in the single system. In the fuel flow totalizing system, the transmitters connect electrically, so their combined signals go into the fuel flow amplifier as one. FUEL FLOW AMPLIFIER – The fuel flow amplifier is an electronic device that supplies power of the proper size and phasing to drive the indicator. The speed at which the indicator motor runs depends on the transmitter signal going into the amplifier. FUEL FLOW TOTALIZER INDICATOR – The fuel flow totalizer indicator contains a two-phase variable speed induction motor. This motor travels in one direction only; however, the speed varies. As the rate of fuel consumption increases, more and more power goes to the indicator motor. This causes the speed of the motor to increase proportionally to the rate of fuel consumption. The motor turns a magnetic drum-and- cup linkage (similar to the tachometer indicator hysteresis disk), which causes pointer deflection. The deflection is proportional to the motor speed, and thus proportional to the rate of fuel consumption. At the same time, a linkage with a friction clutch drives the pounds-fuel-remaining indicator. The clutch is disengaged when using the reset knob to set the reading on the pounds-fuel- remaining indicator. OIL PRESSURE SYSTEM Oil pressure instruments show whether oil is circulating under proper pressure. An oil pressure drop warns of impending engine failure due to lack of oil, oil pump failure, or broken lines. Oil pressure shows on an engine gauge unit (Figure 6-45). This unit consists of three separate gauges in a single case, oil pressure, fuel pressure, and oil temperature. The gauge has a 6-55

Figure 6-47 — Bourdon tube oil pressure instrument. Figure 6-46 — Bourdon tube oil pressure gauge. Bourdon tube mechanism for measuring fluid under pressure (Figure 6-46). The instrument’s oil pressure range is from 0 to 200 pounds per square inch (psi). You read the scale in graduations of 10 psi. There is a single connection on the back of the case leading directly into the Bourdon tube.

In some aircraft, the oil pressure gauge is a separate instrument (Figure 6-47). This instrument operates on the Bourdon tube principle. The synchro system is another method of measuring oil under pressure. This type of oil pressure system is used on most modern aircraft. Essentially, it is a method of directly measuring engine oil pressure. After the measurements are taken, they go electrically from the point of measurement to the synchro indicator on the instrument panel. The synchro system ends the need for direct pressure lines from the engine to the instrument panel. It also reduces the chances of fire, loss of oil or fuel, and mechanical difficulties. The synchro system consists of a synchro indicator and transmitter. The synchro transmitter consists of a permanent magnet moving within a stator. The stator is a circular core of magnetic material wrapped with a single, continuous toroidal winding. Taps divide the winding into three sections. Voltages in each of the sections vary with the position of the permanent magnet. As the magnet moves, the ratio between the three signal voltages varies accordingly. 6-56

Figure 6-48 — Schematic of a synchro oil pressure indicating system. Look at Figure 6-48. Here, you can see the transmitter and indicator connect in parallel. When excited by the same fundamental source, the signal voltages in corresponding sections of the two stators are equal and balanced. The signal voltages remain equal and balanced as long as the magnets are in the same relative positions. However, if the transmitter magnet moves to a new position, the voltages in the three sections of the transmitter are no longer the same. They now differ from the voltages in the corresponding sections of the indicator. Because of this imbalance, current flows between the two units. This circulating current sets up additional magnetic lines of force in each stator, which establishes a magnetic force between the stator and the magnet of each unit. Since the indicator magnet is free to turn, it moves to a position corresponding to the position of the transmitter magnet. The indicator magnet connects to the indicator pointer by a shaft to provide a visual indication. The electrical leads between the transmitter and the indicator may be any reasonable length without noticeable effect on the indication.

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NOTE You can see that the transmitter has a vent to the atmosphere. This allows the transmitter to accurately measure the differential between pump pressure and atmospheric pressure. Figure 6-49 — Fuel pressure synchro system. FUEL PRESSURE SYSTEM The fuel pressure gauge provides a check on the operation of the fuel pump and fuel pressure relief valve. The pilot must check the gauges often to ensure that the fuel pressure is correct. With the fuel pressure correct, the engines have a full range of power at all altitudes. The fuel pressure gauge operates on the same principle as the oil pressure gauge. Fuel pressure indicators may be located in the cockpit by means of synchro systems. This type of system is the same for both fuel and oil pressure indications. However, the oil system transmitter is NOT interchangeable with the fuel system transmitter. Look at Figure 6-49. This synchro system is used to show fuel pressure. A change in fuel pressure introduced into the synchro transmitter causes an electrical signal to go through the interconnecting wiring to the synchro receiver. This signal moves the receiver rotor and the indicator pointer a distance proportional to the amount of pressure exerted by the fuel.

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Figure 6-50 — Exhaust nozzle position indicating system. OIL TEMPERATURE SYSTEM Two types of oil temperature gauges are available for use in the engine gauge unit. One unit consists of an electrical resistance type of oil thermometer, supplied with electrical current by the aircraft dc power system. The other unit, the capillary oil thermometer, is a vapor pressure thermometer. It consists of a bulb connected by a capillary tube to a Bourdon tube and a multiplying mechanism connected to a pointer. The pointer shows the oil temperature on a dial. EXAUST NOZZLE INDICATING SYSTEM The exhaust nozzle position indicating system shows the pilot engine variable exhaust nozzle position. This indication, in turn, provides a measure of percentage of afte rburning, since constant temperatures are indicated throughout the afterburner range. Each engine has a separate but identical nozzle position indicating system. Each system consists of a transmitter potentiometer in the nozzle area control unit and an indicator on the main instrument panel. Power for the system is from the essential 28- volt dc bus. Each indicator is a hermetically sealed unit containing a single receptacle for a mating plug electrical connection. The instrument scale ranges from OPEN to CLOSE, with markings at the ¼, ½, and ¾ positions (Figure 6-50).

The transmitter potentiometer consists of a resistance winding with a movable brush. This brush connects to a linkage within the nozzle area control and moves in relation to the variable exhaust nozzle. Current in the resistance winding is picked up by the movable rush, and it varies according to the location of the brush. Then, this signal current goes to one of the indicator field coils. The indicator contains two field coils and 6-59

Figure 6-51 — Torquemeter system: (A) pickup assembly; (B) indicating system. a rotor. The polarized rotor mounts on a free-moving shaft. The shaft is located in the center of the magnetic field created by the two coils. One coil connects to the transmitter potentiometer in the nozzle area control. The second receives a constant current to give smooth indicator operation. The rotor aligns itself with the magnetic field. The magnetic field varies as the signal received from the potentiometer varies. A pointer mounted on the rotor shaft shows rotor position in relation to nozzle position. TORQUEMETER SYSTEMS The electric torquemeter system in turboprop aircraft measures the torque (horsepower) produced by the engine at the extension shaft. Each system consists of a transmitter (part of the engine extension shaft), a phase detector, and an indicator (Figure 6-51). The system measures the torsional deflection (twist) of the extension shaft as it sends power from the engine to the propeller. Magnetic pickups detect and measure this deflection electronically. The indicator registers the amount of deflection in shaft horsepower. 6-60

Figure 6-52 — Fuel quantity indicator. The extension shaft of the engine consists of two concentric shafts. The inner shaft is the power-transmitting shaft. The outer shaft attaches to the inner shaft at the rear end. Toothed flanges on the front end of each shaft rotate in the field of magnetic pickups. When the engine is running, the teeth on the flange of the driving (inner) shaft move in relation to the teeth on the flange of the outer reference shaft. The displacement between the shafts is proportional to the torque load on the driving shaft. This displacement causes a phase displacement between the pickup signals. The phase angle of the resultant signal is linearly proportional to the shaft deflection. The phase detector and amplifier in the indicator convert the signal to current. The current goes to a servomotor, which drives the indicator pointers. The servomotor also drives the rotor of a synchro control transformer in the indicator. The synchro control transformer balances the synchro system when the pointer registers the measured torque. MISCELLANEOUS INSTRUMENT SYSTEMS On most Navy aircraft, another group of instruments does not fall under flight instruments or engine instruments. Usually referred to as miscellaneous instruments, this group consists of instrumentation for such systems as fuel, hydraulics, flap and gear positions, and cabin pressure. Capacitive-Type Fuel Quantity Indicating System The capacitive-type fuel quantity system electronically measures fuel weight (not gallons) of the fuel in the tanks of an aircraft. The main units of the system are an indicator, tank probes, a bridge unit, and an amplifier. In some systems, the bridge unit and amplifier are one unit mounted in the same box. In the design of newer systems, the bridge and a transistorized amplifier are contained in the indicator. Fuel Quantit y Indicator The fuel quantity indicator (Figure 6-52) is a hermetically sealed, self-balancing, motor- driven instrument. It contains a motor, pointer assembly, transistorized amplifier, bridge circuit, and adjustment potentiometers. As the quantity of fuel in the tank changes, the capacitance value of the tank probe changes proportionately. The tank probe is one arm of a capacitance bridge circuit. The change of capacitance of the probe unbalances the bridge circuit of the amplifier power unit. The unbalance in the circuit causes an error 6-61

Figure 6-53 — Fuel quantity transmitter. v oltage. The amplified error voltage goes to the motor. The motor drives the pointer mechanism and the rebalancing potentiometers to restore the bridge to a balanced condition. The direction of change in the capacitance of the probe unit determines the phase of the error voltage. The phase determines the direction of motor rotation and, therefore, the direction of pointer movement. Tank Probe A tank probe and a simplified version of a tank circuit are shown in Figures 6-53. The capacitance of a capacitor depends upon three factors, the area of the plates (A), the distance between the plates (d), and the dielectric constant (K) of the material between the plates, or C=

Where A = the area of the plates d = the distance between the plates K = the dielectric constant of the materials between the plates

The only variable factor in the tank probe is the dielectric of the material between the plates. When the tank is full, the dielectric material is all fuel. Its dielectric constant is about 2.07 at 0 °C, compared to a dielectric constant of 1 for air. When the tank is empty, there is only air between the plates, and capacitance is less. Any change in fuel quantity between full and empty produces a corresponding change in capacitance.

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Figure 6-54 — Fuel quantity indicator schematic. S ys tem Operation Look at Figure 6-54. As the tank unit capacitance increases or decreases, it is necessary to maintain the bridge circuit in a balanced condition. This prevents the indicator motor from continually changing the position of the indicating needle. To balance the bridge circuit, a balancing potentiometer (R128) connects across part of the transformer secondary. The indicator motor drives this potentiometer wiper in the direction necessary to maintain a continuous balance in the bridge.

The circuit shown in Figure 6-54 is a self-balancing bridge circuit. An empty-calibrating potentiometer and a full-calibrating potentiometer connect across portions of the 6-63

Figure 6-55 — Fuel quantity tank units: (A) non-characterized; (B) characterized.

transformer secondary winding. You can adjust these potentiometers so the bridge voltage balances over the empty-to-full capacitance range of a specific system. A test switch (not shown) unbalances the bridge circuit momentarily when checking the operation of the system. When the switch actuates, pin F connects to ground, unbalancing the circuit. As a result, the indicator drives toward the empty end of the dial. Opening the switch should restore the bridge to balance and return the indicator pointer to its original position. This test proves that the system is operating correctly. In installations where the indicator shows the contents of one tank, and the tank is fairly symmetrical, one probe is sufficient. However, for increased accuracy in peculiarly shaped fuel tanks, use two or more tank units in parallel. This configuration minimizes the effects of changes in aircraft attitude and sloshing of fuel in the tanks. Two classes of tank units (Figure 6-55) are used in a typical capacitance fuel quantity measuring system, non-characterized and characterized.

Non-characterized tank units are variable capacitors, vertically mounted in the fuel cell. As the fuel level changes, the capacitance of the tank unit changes. This change in capacitance is uniform the entire length of the tank unit. Characterized tank units are similar in construction and are mounted to the non- characterized tank units. As the fuel level changes, the capacitance of the tank unit changes. This change in capacitance is NOT uniform the entire length of the tank unit. Since neither electrode of the tank unit goes to ground, and one lead to the amplifier is shielded, capacitance to ground does not enter into the circuit. Therefore, the length of the tank unit leads does not affect the accuracy of the system. FUEL CHARACTERISTICS – The characteristics of fuel are such that the dielectric constant and density deviate because of temperature change. Also, the variable factor in the fuel composition changes the dielectric constant and density. The weight by 6-64

volume of aircraft fuel depends on its density, which, in turn, depends on its temperature. As the temperature of the fuel goes down, the density increases. As the temperature of the fuel goes up, the density decreases. Any change in the dielectric constant or density of the fuel affects the movement of the indicator pointer. For example, assume that the indicating system is at balance with the tank unit immersed to a given depth. The fuel it is immersed in has the density and dielectric constant of the fuel for which the system is calibrated. The indicator pointer will then reflect the correct amount of fuel in terms of pounds. Then, the tanks are drained and then refilled to the same level. This time the fuel has a greater density and higher dielectric constant, which causes the pointer to show a greater weight. The new reading is correct only if the effect of the changes in density and dielectric constant are proportional. However, the effect of the increase in dielectric constant is greater than the effect of the increase in density. The results are an incorrect indication. The system reduces this error by varying the capacitance of the reference capacitor in the bridge leg opposite the immersed tank unit. COMPENSATION – The reference capacitor is varied by connecting a compensator unit in parallel with it. The compensator, like the tank probe, is a variable capacitor. However, the compensator mounts at the lowest level of fuel so it is completely immersed until the tank is almost dry. Its capacitance depends on the dielectric content of the fuel rather than the quantity. The compensator connects into the common reference leg for both phases of the bridge circuit. This connection allows it to become a part of the reference capacitance. A change in the dielectric constant of the fuel affects both the tank probe and the compensator capacitance. Therefore, the current change in the tank probe leg of the bridge is counteracted by a similar change in the reference leg of the circuit. Various capacitor-type fuel quantity systems operate on the principle just described. Indicators, tank probes, and power units may differ as to shape, size, and specifications from system to system. For this reason, you should always consult the manufacturer’s manuals for specific information on a particular system. HYDRAULIC PRESSURE INDICATORS In most naval aircraft, the hydraulic system operates the landing gear, flaps, speed brakes, bomb bay doors, and certain other units. Aircraft hydraulic pressure gauges show either the pressure of the complete system or the pressure of an individual unit in the system. A typical direct reading gauge contains a Bourdon tube and a gear-and- pinion mechanism. The mechanism amplifies and transfers the tube’s motion to the pointer. The position of the pointer on the calibrated dial shows the pressure in pounds per square inch. The pumps supplying pressure for operating the aircraft’s hydraulic units are driven by an aircraft engine, an electric motor, or both. Some installations employ a pressure tank or accumulator to maintain a reserve of fluid under hydraulic pressure. In such cases, the pressure gauge registers continuously. With other installations, operating pressure builds up only when needed, and pressure registers on the gauge only during these periods. The pressures of hydraulic systems vary for different models of aircraft. In older pressure systems, the gauges registered from 0 to 2,000 psi. With later model aircraft, the pressure ranges have increased. Some aircraft have systems with pressure ranges as high as 4,000 psi. The trend is away from the direct reading pressure gauge and towards the synchro (electric) type of gauge. 6-65

Figure 6-56 — (A) Hydraulic pressure indicator; (B) hydraulic pressure indicating schematic. Figure 6-56, view A, shows the hydraulic pressure indicator of a late model naval aircraft. This aircraft has two hydraulic systems. The indicating system shows the hydraulic system (HS) No. 1 and HS No. 2 system pressures. The hydraulic pressure indicato r pointers, driven by signals from the hydraulic pressure transmitters, respond to signals from matching synchronous motors in the pressure transmitters. The indicating system consists of two remote hydraulic pressure transmitters and a dual pointer indicator. The system uses 26-volt, 400-Hertz, single-phase alternating current from the 26-volt, single-phase bus.

Each hydraulic pressure system line contains a Bourdon tube type of pressure transmitter. Expansion and contraction of the Bourdon tube travels by mechanical linkage to the rotor of the transmitter synchro. The pressure transmitter synchro sends an electrical signal to the receiving synchro within the indicator. The receiving synchro’s rotor links mechanically to the indicator pointer. The pressure indicator contains two synchros that attach mechanically to two separate pointers. PNEUMATIC PRESSURE SYSTEMS The cabin pressure altitude indicator (Figure 6-57) is a sensitive altimeter that measures cabin pressure. The instrument contains a sensitive diaphragm that expands or contracts with changes in cabin pressure. The altitude equivalent of cabin pressure shows on the dial, in increments of 1,000 feet. The range is from 0 to 50,000 feet. An opening in the back of the instrument case allows it to sense cabin pressure. You can also use this instrument to reflect pressure suit altitude rather than cabin altitude when wearing a pressure suit. 6-66

CAUTION Remove the cabin pressure altitude indicator if the aircraft is to undergo a cabin pressurization test on the ground. Failure to remove the indicator will result in damage to the instrument from excessive pressure. Figure 6-57 — Cabin pressure altitude indicator.

POSITION INDICATING SYSTEM (DC SYNCHRO SYSTEM) The dc synchro system shows remote mechanical conditions, specifically the movement and position of wing flaps, cowl flaps, oil cooler doors, and similar movable parts of the aircraft. The system consists of a transmitter, an indicator, and connecting wires. A dc voltage from the aircraft’s electrical power system supplies the voltage to operate the system. The transmitter mechanically connects to the movable device that is actually supplying the positioning data. The indicator repeats this information on a properly calibrated scale in the indicator on the instrument panel. Figure 6-58 illustrates a complete three-wire system.

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Figure 6-58 — Three-wire dc synchro system.

Transmitter The three-wire system transmitter consists of a continuous circular toroidal resistance winding with two diametrically opposite brushes continuously touching the winding. These brushes apply dc to the winding. The brushes rotate with the movement of the aircraft part to which they are mechanically attached. Indicator The three-wire system indicating element consists of an annular core, a permanent magnet rotor, a damping cylinder, and three field coils. The leads between the coils connect to the three taps of the transmitter winding. As voltages at the transmitter taps vary through brush rotation, the distribution of current in the indicator coils varies. This variation causes the resulting magnetic field of the three coils to position the pointer’s permanent magnet rotor. A copper cylinder provides a damping effect. The induced eddy currents in this cylinder oppose movement of the rotor. This reduces the tendency of the pointer to oscillate. 6-68

Landing Gear An example of a position indicating system is a wheel position indicator. The system consists of three back-mounted indicators. A series of limit switches in the landing gear control circuit control these indicators. The wheel instrument (Figure 6-59, view A) gives a ready indication of the landing gear position. There are three positions on a drum that move about an axis. One position of the drum has a landing gear wheel symbol. The center position has a diagonal barber pole (black and white warning lines). The other position has the word UP. The wheel symbol shows that the wheel is down and locked. The UP indicator shows the wheel is up and locked. The barber pole shows the wheel is somewhere between up and down, or not locked in position. The position indicator operates through the landing gear limit switches in the wheel wells of the aircraft. When the landing gear moves the contact of S1 (Figure 6-59, view A) to the up position, the indicator shows the wheel is up and locked. When the landing gear moves the contact of S2 to the down position, the indicator shows the landing wheel, indicating the wheel is down and locked. Each switch connects to a solenoid in the indicator. As the solenoids energize, the indicator element moves to reveal the position of the aircraft landing gear. When the landing gear is moving, both the up lock and down lock switches release. This opens both circuits to the corresponding indicator, causing the black and white barber pole to show. It also causes the warning light in the landing gear handle to illuminate. Figure 6-59, view B, shows a schematic for a complete aircraft landing gear position-indicating system.

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Figure 6-59 — Landing gear position indicating system: (A) typical; (B) system schematic.

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Figure 6-60 — Flap position indicating system. Figure 6-61 — DC synchro flap position indicating system. Flaps You can show the aircraft flap position in a manner similar to the landing gear. These indicators show UP, ½, DN (down), and barber pole. The indicator (Figure 6-60) energizes through limit switches. Other aircraft show the position of the flaps using the dc synchro system of remote indication. This system consists of a transmitter and an indicator (Figure 6-61). A change in flap position moves the transmitter rotor, and a similar rotor movement occurs in the indicator on the instrument panel. A pointer attached to the indicator rotor shows the amount of travel of the flaps in percent of full extension. The transmitter mounts on the flap drive control unit and actuates by the control actuating mechanism. 6-71

INSTRUMENT SYSTEM MAINTENANCE You need a wide variety of skills to maintain aircraft instruments and their associated equipment. It is important for you to check, inspect, and maintain these instruments because the aircraft will not perform properly unless the instruments present reliable information. Instruments used in high-speed aircraft must give correct indications. The accuracy of instruments such as percent-type tachometers, tailpipe temperature indicators, and gyro attitude indicators require preventive maintenance. The existence of excessive errors in instrument systems directly relates to flight safety and efficient aircraft performance. You cannot assume that borderline instrument errors are acceptable. This is particularly important in high-speed aircraft. As an AE, you will perform functional tests on aircraft instruments to make sure they give accurate indications. Operational and functional tests take time. When you perform an inspection, you need to know how the particular aircraft instrument operates. Also, you need to know what tools and test equipment you will need. Without this knowledge and these skills, you cannot properly perform the tasks assigned to your rating. GENERAL MAINTENANCE General maintenance of instruments falls into two categories, scheduled and unscheduled maintenance. The material condition of the systems and reliability of instruments are ensured by the day-to-day maintenance routine. Cases Instruments come in one of four different kinds of cases: 1. One-piece phenolic composition cases 2. Two-piece phenolic composition cases 3. Nonmagnetic all-metal cases 4. Metallic-shielded cases Th e cases come in several different sizes so instruments can be easily removed and maintenance simplified. Special instruments that contain mechanisms too large for adaption to a standard case come in specially designed cases. Instruments easily mount on the instrument panel with locking devices molded into the instrument flange assembly, by spring locknuts or mounting clamps. You can easily remove instruments that use a mounting clamp by unscrewing the tension screw in the instrument’s lower right corner. You do not have to remove the tension screw to release the tension on the clamp assembly. Markings and Graduation Markings on the glass instrument face cover help flight personnel confirm instrument operation to within the prescribed ranges of the equipment. The markings usually consist of a white arc on the outer edges of the instrument glass. They show the normal operating range. A red mark shows the operating limit that should not be exceeded. 6-72

WARNING

An index marking of white paint, not over one-sixteenth inch wide by three-sixteenths inch long, is at the bottom center of all instruments color-marked for operating ranges. Place this index mark at the point between the glass and the case. This mark will show whether or not the glass cover moves at any time after marking the ranges. Obtain the proper range markings for the aircraft instruments from the Naval Air Training and Operating Procedures Standardization (NATOPS) flight manual for the particular aircraft.

Panels Instrument panels are made from sheet aluminum alloy, with sufficient strength to resist flexing. The panel is nonmagnetic and painted a dull black or gray to eliminate glare and reflection. Some panels are constructed in two layers, and the instrument faces are flush with the rear panel. The front panel is a reflector panel that mounts over the rear panel with sufficient clearance to supply an indirect lighting effect. The indirect lighting system is not standard for all aircraft. Some aircraft have spotlights, edge lighting, or a combination of these. Some instruments have their own internal lighting system. Instrument panels are shock mounted to absorb low frequency, high-amplitude shocks. The mounts consist of square-plated absorbers in sets of two, each secured to separate brackets. You should inspect the mounts periodically for deterioration; if the rubber is cracked, replace the pair. As an AE, your instrument maintenance duties include certain inspections that you should conduct at regular intervals. Your daily inspection includes the following checks in accordance with applicable Periodic Maintenance Instruction Cards, Phase Cards, Daily Inspection Requirements Cards or MIM as required:  Check pointers for excessive errors. Some indicators should show existing atmospheric pressure, existing temperatures, etc. Others should indicate zero.  Check instruments for loose or cracked cover glasses. Replace pitot-static instruments if damaged.  Check instrument lights for proper operation.  Check caging and setting knobs for freedom of movement and correct operation.  Carefully investigate any irregularity the pilot reports. When performing a phase/calendar inspection, make the following checks in accordance with applicable Periodic Maintenance Instruction Cards, Phase Cards, Daily Inspection Requirements Cards or MIM as required:  Check the mounting of all the instruments and their dependent units for security.  Check for leaks in instrument cases, lines, and connections.  Check for dull or marred luminous paint on dial markings and pointers.  Check the condition of operation and limitation markings. 6-73

 Check for contact and condition of bonding on instruments.  Check shock mountings for condition of rubber and security of attachment.  Check for freedom of motion of all lines and tubing behind the instrument panel. Also, check that they are properly clamped or taped to avoid chafing, and that they are free from moisture, crimps, etc. After starting the engine, check the instrument pointers for oscillation. Also, check the readings for consistency with engine requirements and speeds. On multi-engine aircraft, check the instruments for the various engines against each other. Investigate any inconsistency; it may indicate a faulty engine, component, or instrument. After you have diagnosed a particular discrepancy and found the instrument to be faulty, remove and turn it into supply. Remember the following precautions when removing and installing instruments:  Handle instruments carefully at all times. Additional damage may result if you abuse the instrument.  Do not change the location of an indicator.  Do not force the mounting screws. If the screw is cross-threaded, replace it. Do not draw the screws up too tight against the panel. This may distort the case enough to affect the operation of the instrument, crack the case, or break off the mounting lugs.  When removing or installing tubing of a pressu re-operated instrument, use a backup wrench to avoid twisting the tubing or fitting. Do not exert undue force while tightening the connection.  Install all electrical plugs hand tight.  Before connecting an electrical plug to an instrument, check the plug for bent or broken pins.  Cap the open electrical receptacles, plugs, and hose connections to prevent foreign material from entering the instrument or system. Aircraft Plumbing Rigid and flexible tubing is extensively used in aircraft. These tubes come in many different sizes. Sizing is usually determined by the outside diameter of the tube and ranges from one-eighth inch to 2 inches in diameter. The type of material and wall thickness determines the amount of pressure that a tube can safely withstand. When replacin g or repairing tubing, you should use caution to make sure that you use the proper type. You can find detailed information on tubing and tubing repairs in the aircraft-specific technical manual. RIGID TUBING – Rigid metal tubes are widely used in aircraft for fuel, oil, coolant, oxygen, instrument, hydraulic and vent lines. Corrosion-resistant steel (stainless steel) and aluminum alloy tubing are the most commonly used tubing. You may identify the basic tube material by either visual inspection or by the alloy designation stamped on the tubing. Tube fittings connect tubes together and connect tubes to instruments. The shape of the fitting determined by the particular installation; some are straight, while others have various angles. The fittings secure to the tube by a beaded or flared joint. The beaded joint (or upset joint) is used in low-pressure lines. Systems that use low-pressure lines 6-74

Figure 6-62 — Electrical line identification application. include vacuum, deicer, and oil systems that use rubber hose fittings. All high-pressure and some low-pressure lines use flared joints. Grip dies and flaring or beading tools are used to form flared and beaded joints. When working rigid tubing, you should consult the manuals on the beading and flaring tools to use the tools properly. If piping and lines are damaged, you should replace them with new parts. To repair tubing, you must determine how much tubing to remove. Consider the following factors when deciding how much tubing to remove:  The location of the tubing  The extent of damage  The most convenient location for tool manipulation There is a tendency to over tighten tubing nuts to prevent high-pressure fluid from escaping. Such over tightening may severely damage or completely cut off the tube flare. When you remove a tube, check the flare. If you find a flare with less than 50 percent of its original wall thickness, reject the tube. When bending tubing, you must be careful to prevent collapsing of the tube at the bend. When making bends for fluid tubing, make sure that you use the proper bending radius. These specifications can be found in the aircraft-specific technical manual. Bands of paint or strips of tape around the line near each fitting identify each rigid line in the aircraft. There is at least one identifying marker in each compartment. All lines less than 4 inches in diameter have identification tape on them. The exceptions to this rule are cold lines, hot lines, and lines in an oily environment. Another except ion is any line in engine compartments where there is a chance of the tape going into the engine intake. In these cases, and all others where you do not use tape, use paint to identify the lines. Identification tape codes show the function, contents, hazards, direction of flow, and pressure in the fluid line. When applying these tapes, refer to MIL-STD-1247C. MIL- STD -1247C standardizes rigid line identification throughout the Department of Defense (Figure 6-62).

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The function of the line is identified by a 1-inch-wide tape that contains printed word(s), color(s), and a geometric symbol. Functional identification markings (see MIL-STD- 1247C) are the subject of an international standardization agreement. Three-fourths of the total width on the left side of the tape is a color or color code. This code shows one function only per color or colors. The function of the line is printed in English across the colored portion of the tape. Color-coding is utilized to provide universal recognition and understanding regardless of language for aircraft maintenance and troubleshooting. The right-hand one-fourth of the functional identification tape contains a geometric symbol. This symbol is different for every function. The symbol allows a colorblind person to identify the line function by means of the geometric design rather than by the color. Refer to Figure 6-63. Here, you see a listing of the functions and their associated identification media as used on the tapes.

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Figure 6-63 — Functional identification tape data.

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The identification-of-hazards tape shows the hazard associated with the contents of the line. Tapes that show hazards are one-half inch wide, with the abbreviation of the hazard printed across the tape. There are four general classes of hazards found with fluid lines. 1. Flammable Material (FLAM). The hazard marking FLAM identifies all materials ordinarily known as flammables or combustibles. 2. Toxic and Poisonous Materials (TOXIC). TOXIC identifies lines containing materials that are extremely hazardous to life or health. 3. Anesthetics and Harmful Materials (AAHM). AAHM identifies all materials producing anesthetic vapors and all liquid chemicals and compounds hazardous to life and property. However, they do not normally produce dangerous quantities of fumes or vapors. 4. Physically Dangerous Materials (PHDAN). PHDAN marks a line carrying material that is not dangerous within itself. However, the material is asphyxiating in confined areas or is generally handled in a dangerous physical state of pressure or temperature. Table 6-2 lists some of the fluids that you may work with and the hazards associated with each.

Table 6-2 — Hazards Associated With Various Fluids CONTENTS HAZARD Air (under pressure) PHDAN Alcohol FLAM Carbon dioxide PHDAN Freon PHDAN Gaseous oxygen PHDAN Liquid nitrogen PHDAN Liquid oxygen PHDAN Liquid Petroleum Gas (LPG) FLAM Nitrogen gas PHDAN Oils and greases FLAM JP-4 FLAM Trichloroethylene AAHM

FLEXIBLE TUBING (HOSE) – Flexible hose assemblies consist of lengths of hose coupled with threaded end fittings. There are two classes of flexible tubing, high pressure and low pressure. 6-78

You can get the specifications for flexible hose by interpreting the identification code on the hose. This identification, a series of dots and dashes, gives hose size, temperature range, and date of manufacture. The date of manufacture is in quarter of year and year. Refer to the aircraft-specific technical manual for a detailed discussion of flexible hose identification. You cannot construct high-pressure flexible hose at the organizational level. You must order it through supply, as organizational maintenance cannot perform high-pressure tests. You can order the parts for a low-pressure flexible hose through supply and make the hose locally. You can reuse fittings from a damaged hose if they meet the required specifications. When you install hose, make sure it will not twist under any operating condition. This type of installation lessens the tendency for connecting fittings to loosen. When you replace a hose in hydraulic, fuel, oil, alcohol, and pneumatic systems, make sure the new hose is an exact duplicate of the old hose. Specifically, the length, outside diameter, inside diameter, material, type, and shape (except on directed modifications) must be the same as the old hose. If a bend is necessary when installing hose in fluid systems, the radius must not be smaller than the minimum specified in the aircraft-specific technical manual or as Naval Air Systems Command (NAVAIR) instructions may direct. When practical, use a radius that is larger than the specified minimum. When you install hose through holes in brackets and when you use supporting clips, ensure there is no reduction in hose diameter. When these conditions are present, they redu ce flow, and damage to the hose may occur. The hose must have support every 24 inches. Closer supports are desirable when practical. The flexible line support should never cause deflection of the rigid connecting lines under any possible relative motion that may occur. Flexible hose between two rigid connections may restrain excessive motion where necessary, but never be rigidly supported. To avoid chafing, use suitable bulkhead-type grommets or cushioned clips. Protect hose installations from excessive temperature, such as exhaust blasts and supercharger ducts, by either shrouding or relocating. Use flame-resistant hose forward of the firewall on certain aircraft as specified in the aircraft-specific technical manual or NAVAIR instructions may direct. Where hoses connect to an engine or to engine-mounted accessories, provide 1½ inches of slack between the last point of support and engine attachment. This prevents the chance of the hose pulling off the nipple due to engine movement. Whenever possible, install the hose so all hose markings are visible. All hose materials deteriorate from exposure to heat, sunlight, excessive moisture, and ozone. Therefore, you should stow hoses in a cool, dry place and away from electrical equipment. You can obtain age limits of shelf items based on the manufacturer’s code from current accessory bulletins. Stow hose in straight lengths to prevent it from setting in a curved position. Replace the hose if cover material is peeling or flaking, or if the braid reinforcement is open to the elements. PITOT-STATIC INSTRUMENT MAINTENANCE Maintenance of the pitot-static system consists of checking the lines for integrity, water, and miscellaneous obstructions. A pressure check is of the utmost importance as any 6-79

CAUTION Do not touch the pitot tube with your bare hand with the heat on. The extreme heat may cause your skin to stick to the surface. slight leak will result in erroneous indications of the instruments during flight. You should also check the pitot heater for proper operation. You can do this by monitoring a voltage drop when the heat is turned on, or by checking the pitot tube for the presence of heat.

The following is an outline for water and debris removal from the pitot-static system. For specific procedures, you should refer to the MIM. 1. Disconnect all altimeters, airspeed indicators, rate- of-climb indicators, and any other systems receiving information from the pitot-static system. Disconnect the lines from pitot or pitot-static tubes. 2. Remove all drain caps in the system. 3. Circulate a stream of clean, dry, filtered air at medium pressure through the complete system. Be careful not to include the cabin pressurization system static vent. Be certain that air is flowing from the exit end of each line. 4. Inspect all static vents and the pitot tube water removal drain holes for damage and evidence of foreign matter and obstructions. Check all low points in the lines for possible cracks due to icing in the lines. 5. Replace and secure all system drain caps. 6. Reconnect all instruments. Tighten connections properly; do not kink or bend the lines. 7. Using a field test set or other approved tester, thoroughly check the system for proper operation and leaks. The maintenance of the pitot-static system is relatively simple when compared to more complex systems. However, its maintenance is not a minor task. INSTRUMENT TESTING Operation of most aircraft instruments is entirely automatic. Once installed, the units require no further maintenance or servicing other than routine and periodic inspections. If a system or instrument malfunctions, you must first localize the source of trouble. Develop a systematic troubleshooting procedure. The procedure should include the possible service troubles and their remedies for each type of instrument. You will find most of this information in applicable aeronautic publications, such as the MIM for specific aircraft and the service instruction manual for specific instruments. An instrument that does not function properly or that is suspected of being unserviceable must first be checked to determine if the instrument or the installation is at fault. Usually, instrument problems fall into three groups, trouble in the power supply, trouble in the unit, or trouble in the connections to units, either electrical or mechanical. If the installation is faulty, line maintenance can correct the problem. If the instrument is 6-80

the fault, you can remove and replace it with a serviceable unit. The defective unit goes to a qualified instrument overhaul depot for detailed inspection, overhaul, and repair. Only authorized instrument shops can open instrument cases and make repairs or adjustments. When making ground tests of electrical instruments, you should connect an external power supply to the aircraft. Do not use the battery when conducting ground tests of equipment. When performing ground testing, you should use portable field test sets, such as the ADTS405-8325 Air Data Test Set, the CA-11-FCS RPM Test Set, or appropriate authorized Individual Material Readiness List (IMRL) asset. Test sets are discussed in chapter 2 of this RTM. Always use a precision voltmeter to check instrument power. You can check most electrical instruments with a test indicator to determine where the trouble lies. For example, you can check synchro indicators using a synchro test indicator.

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End of Chapter 6 AIRCRAFT INSTRUMENTS Review Questions 6-1. What are the two ways aircraft instruments are grouped?

A. Mechanical and electrical B. Digital and manual C. Operating principles and jobs they perform D. Aircraft type and model

6-2. What determines the air pressure at any given altitude?

A. Temperature and moisture at that altitude B. Weight of the air above that altitude C. Density at that altitude D. Gravity at that altitude

6-3. At 5,000 feet of altitude, what is the standard atmospheric pressure, in pounds per square inch?

A. 12.23 B. 13.05 C. 15 D. 29

6-4. What are the three indicators that use the pitot-static system?

A. Line of flight, impact pressure, and airspeed B. Barometer, static air vent, and pitot tube C. Temperature, climbing, and descending D. Airspeed, altimeter, and vertical speed indicator

6-5. Which of the following is a description of impact pressure?

A. The force of air against the aircraft B. Various air pressure that causes different readings C. Pressure of the air closest to the terrain D. The normal outside atmospheric temperature

6-6. What instrument uses both pitot and static air pressure?

A. Airspeed indicator B. Altimeter C. Counter pointer pressure altimeter D. Density altitude

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6-7. What does the acronym “AGL” stand for?

A. Above Ground Level B. Above Gain Latitude C. Altitude Gain Level D. Absolute Gravity Level

6-8. What is “absolute altitude”?

A. Altitude of the aircraft in flight B. Altitude of the terrain the aircraft is flying over C. Distance between the aircraft and the terrain it is flying over D. Distance between the terrain and sea level

6-9. What altitude does gravity acting on the atmosphere produce a pressure of 14.70 psi and support a column of mercury to a height of 29.92 inches?

A. Pressure altitude B. Mean sea level C. Absolute altitude D. Density altitude

6-10. What is the name for the aneroid mechanism used in most altimeters?

A. Wafer B. Linkage C. Pinion D. Bimetal yoke

6-11. If an aircraft is in level flight, what will the Vertical Speed Indicator (VSI) indicate?

A. -0.5 B. 0.0 C. +0.5 D. +1.0

6-12. What are the four inputs to the Air Data Computer (ADC) system?

A. Total static pressure, angle of attack (AOA), impact pressure, and altitude B. Total pressure (pitot), indicated static pressure, indicated AOA, and total temperature. C. AOA, temperature, speed, and altitude D. Indicated AOA , altitude, speed, and temperature

6-13. Which of the following is the meaning of the term angle of attack (AOA)?

A. The difference between the leading edge of the wing and nose of the aircraft relative to the air through which it is passing B. The angle at which the leading edge of the wing must pass to provide adequate lift for sustained flight C. The angle at which the leading edge of the wing encounters the air mass D. The angle of the air passing over the elevators to provide more lift 6-83

6-14. Static pressure errors become a significant factor in the accuracy of pressure indications at what relative speeds?

A. Supersonic only B. Transonic only C. Supersonic and transonic D. Subsonic and transonic

6-15. Total temperature equals the ambient temperature plus the ________.

A. temperature of the engine intake air B. temperature increase created by the motion of the aircraft C. temperature decrease created by the surrounding air D. temperatures of the ram air external to the aircraft and the engine intake air

6-16. What system actuates the rudder pedal shaker to warn the pilot of an impending stall?

A. AOA B. Angle of sideslip compensator C. Airstream detector D. Approach index lights

6-17. If the red arrow on the pilot’s AOA indexer illuminates, which of the following conditions exists?

A. The aircraft is nose low B. The aircraft is nose high C. The aircraft is at optimum AOA D. There is a failure in the system

6-18. What determines the degrees of freedom for a gyro?

A. The size of the gimbals B. The number of gimbals supporting the gyro C. The material that is used for the gyro D. The space where the gyro is mounted

6-19. What are the two basic properties of gyroscopic action?

A. Rigidity in space and precession B. Angular value and degree of freedom C. Restricted and semi rigid D. Free and universal

6-20. What indicator shows the lateral attitude of an aircraft?

A. Roll and pitch B. Ball and turn pointer C. Accelerometer D. Turn-and-bank 6-84

6-21. An aircraft weighs 10,000 pounds and has a lift equal to 20,000 pounds. Which of the following readings will the accelerometer show?

A. -2 g B. 0 g C. +1 g D. +2 g

6-22. What is the reading on the accelerometer of an aircraft weighing 40,000 pounds and traveling in straight-and-level flight?

A. +2 g B. +1 g C. 0 g D. -1 g

6-23. What tachometer generator output is proportional to the engine speed?

A. Frequency B. Temperature C. Power D. Torque

6-24. When there is an open in the voltage supply circuit, what will the galvanometer of a Wheatstone bridge read, in ohms?

A. Zero B. 10 C. 50 D. 100

6-25. In a radiometer-type temperature indicator, what component determines needle po sition?

A. How the poles of the permanent magnet align to external magnetic source B. How much current is measured by the contraction potentiometer C. How much resistance is indicated by the center potentiometer D. How the permanent magnet aligns to the flux of the two coils

6-26. What is a thermocouple?

A. Copper and strip of iron material B. A junction of two unlike metals used to sense temperature C. Hot junction and cold junction used to measure temperature differences D. Pressed constantan used in aircraft instruments

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6-27. In the fuel flow transmitter, what component senses vane movement and sends it to the synchro motor?

A. Synchro transmitter B. Relief valve C. Chamber housing D. Magnet (ring magnet and bar magnet)

6-28. What are the two ways of indicating oil pressure?

A. Bourdon tube and synchro system B. Synchro system and stator C. Synchro system and visual D. Visual and manual

6-29. What is the reference shaft of the torque meter system?

A. Power transmitting B. Outer C. Torque load D. Driving

6-30. What are two classes of tank units used in a typical capacitance fuel quantity measuring system?

A. Ridged and parallel B. Non-characterized and characterized C. Immersed and shielded D. Dielectric constant and compensation

6-31. What does the density of fuel depend on?

A. Temperature B. Altitude C. Humidity D. Size of tank

6-32. What unit in the fuel quantity system depends on the dielectric content rather than the quantity?

A. Reference capacitor B. Pressure indicator C. Probe leg D. Compensator probe

6-33. When installing rigid tube markings, what is the minimum number per compartment?

A. One B. Two C. Three D. Four 6-86

6-34. What is the color of the identification tape on electrical conduits?

A. Red and brown B. Brown and orange C. Black and white D. Green and red

6-35. What are the four classes of hazards identified by hazard tape?

A. Corrosive, toxic, flammable, and gas B. Flammable material, toxic, explosive, and inert C. Flammable material, toxic and poisonous materials, anesthetics and harmful materials, and physically dangerous materials D. Poisonous, corrosive, explosive, and physically dangerous materials

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6-88

CHAPTER 7 COMPASS AND INERTIAL NAVIGATION SYSTEMS The material in this chapter is about aircraft navigation systems. The basic systems discussed are the aircraft compass system and Inertial Navigation System (INS). Also, this chapter presents a discussion of the calibration of these two systems. The way electrical signals are detected, amplified, and delivered to various indicators and systems is highly sophisticated. Before you begin this chapter, you might need to read Navy Electricity and Electronics Training Series (NEETS), Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. State the navigation-related terms and definitions basic to compass and inertial navigation system operation. 2. Explain the operating principles and features of compass systems, the attitude reference system, and associated sensors and indicators. 3. Summarize the operating principles and characteristics of the inertial navigation system, to include Schuler loops and tuning; and identify navigation errors and aligning and calibration procedures. 4. Describe the two types of Inertial Navigation Systems and discriminate between systems within those two types. NAVIGATION TERMS AND DEFINITIONS Any purposeful movement in the universe involves an intention to proceed to a definite point. Navigation is the business of proceeding so you will arrive at that point. Air navigation is defined as the process of directing the movement of an aircraft from one point to another. The function of air navigation is to locate positions and measure distance and time along the intended direction of flight. Position Position is a point defined by stated or implied coordinates. You will frequently qualify this term by such adjectives as estimated, dead reckoning, no wind, etc. However qualified, the word position always refers to some place that you can identify. One of the basic problems of the navigator is that of fixing his position. If he does not know where he is, he can’t direct the movement of the aircraft to its intended destination. Direction Direction is the position of one point in space relative to another, without reference to the distance between them. Direction may be either three-dimensional or two- dimensional, the horizontal being the usual plane of the latter. For example, the direction of San Francisco from New York is approximately west (two-dimensional). However, the direction of an aircraft from an observer on the ground may be west and 20° above the horizontal (three-dimensional). Direction (for example, east) is not itself 7-1

an angle, but it is often measured in terms of its angular distance from a reference direction. Course Course is the intended horizontal direction of travel. For example, the direction of NAS Jacksonville from NAS Pensacola is east. This should be the intended direction of flight. However, because of wind conditions aloft, the aircraft might not head straight toward Jacksonville, but somewhat to one side. No matter what the aircraft heading is, the course, the intended direction, is still east. Heading Heading is the horizontal direction in which an aircraft is pointing. In the previous example, you can see the difference between course and heading. Heading is the actual orientation of the aircraft’s longitudinal axis at any instant, while course is the direction of travel intended. True heading uses the direction of the geographic North Pole as the reference. Magnetic heading uses the direction of the earth’s magnetic field at that location as the reference. Magnetic heading differs from true heading by the amount of Magnetic Variation (MAGVAR) at that location. Compass heading differs from magnetic heading by the amount of magnetic deviation. Compass heading differs from true heading by the amount of compass error (deviation ± variation). Bearing Bearing is the horizontal direction of one terrestrial point from another. Bearings can be expressed by reference to two terms—true north or the direction in which the aircraft is pointing. If true north is the reference direction, the bearing is a true bearing. If the reference direction is the heading of the aircraft, the bearing is a relative bearing. If you get a bearing by radio, it is a radio bearing; if visual, it is a visual bearing. Thus, the direction between two objects on (or near) the surface of the earth can be described concisely by saying: THE (RADIO, VISUAL) BEARING OF A FROM B IS X ± (RELATIVE, TRUE). Distance Distance is the separation between two points. To measure distance, you measure the length of a line joining the two points. This seems understandable enough. However, suppose that the two points are on opposite sides of a baseball. How do you draw the line? Does it run through the center of the ball, or around the surface? If around the surface, what path does the line follow? You must qualify the term distance used in navigation to show how to measure the distance. The shortest distance on the earth’s surface from San Diego to Sydney, Australia, is 6,530 miles. However, via Honolulu and Guam, a frequently used route, it is 8,602 miles. You can express the length of a chosen line in various units, such as miles, kilometers, or yards. Time Time has many definitions. The two definitions used with navigation are as follows: (1) the hour of the day and (2) an elapsed interval. The first appoints a definite instant, as when takeoff time is 0214. The second definition appoints an interval, such as time of flight of 2 hours 15 minutes. 7-2

Figure 7-1 — Schematic representation of Earth showing axis of rotation and equator. Figure 7-2 — The equator is a great circle whose plane is perpendicular to the polar axis. Poles The earth’s geographic poles are the extremities of the earth’s axis of rotation. Look at Figure 7-1. Here, Pole north (Pn), East (E), Pole south (Ps) and West (W) represent the surface of the earth at sea level. Line PnPs is the axis of rotation. The earth’s rotation is such that all points in the hemisphere, PnWPs, approach the viewer. Those points in the opposite hemisphere will recede from the viewer. The extremities of the axis, points Pn and Ps are the north and south poles, respectively. A man on the surface of the earth, facing in the direction of rotation, has the North Pole on his left. East will be in front of him, the South Pole on his right, and west behind him. The earth has some of the properties of a bar magnet. The magnetic poles are the regions near the ends of the magnet. This is where the highest concentration of magnetic lines of force exists. However, the earth’s magnetic poles a re not at the geographic poles, nor are they antipodal (opposite) to each other.

Great Circles and Small Circles The intersection of a sphere and a plane is a circle. The intersection is a great circle if the plane passes through the center of the sphere. It will be a small circle if it does not. Parallels and Meridians Look at Figure 7-2. Here, the earth’s equator is a great circle.

7-3

Figure 7-3 — The plane of a parallel is parallel to the equator. Figure 7-4 — Great circle through the poles form meridians.

If a second plane (Figure 7-3) passes through the earth parallel to the equator, its intersection is a small circle. Small circles don’t always have planes perpendicular to the polar axis. However, if they are perpendicular, then all points on the small circle are equidistant from the equator; that is, the circles are parallel to the equator. Such small circles, together with the equator, are PARALLELS. They provide one component of a system of geographical coordinates.

Now, suppose that planes pass through the earth’s poles (Figure 7-4). Such planes contain the axis, and since they also contain the center, they form great circles at the surface. Great circles through the poles of the earth are MERIDIANS. All meridians are perpendicular to the equator. Meridians form the second part of a system of geographical coordinates commonly used by navigators.

7-4

Figure 7-5 — Latitude of M is angle QOM or arc QM. Latitude and Longitude You can identify any point on Earth by the intersection of a parallel and a meridian. It is the same as locating an address at the corner of Fourteenth Street and Seventh Avenue. You are just using different names for identifying the parallels and meridians. The circumference of a circle is divided into 360 units. This unit is the degree. It is the same unit you use to measure an angle. In Figure 7-5, circumference equator (Q)PnQPs represents a meridian. QQ´ represents the equator, whose plane passes through the axis of rotation. Let M be some position north of the equator on a meridian. The number of degrees in arc QM is the measure of angle QOM. If arc QM is 30°, then angle QOM is 30°. Thus, you measure a central angle by measuring its subtended arc. Let MM´ be the plane of a small circle parallel to QQ´, the equator. Then arc QM measures the distance of any point on MM´ from the equator. You can describe the whole parallel MM´ by saying that it is 30° north of the equator. Similarly, you can say any point on NN´ is 45° south of the equator. The angular distance of a position north or south of the equator is the position’s latitude. You measure latitude northward or southward through 90° and label it N or S to show the direction of measurement. You express latitude in terms of the angle at the center; see angle QOM in Figure 7-5. Latitude, then, is the north-south geographical coordinate. The east-west geographical coordinate is longitude. You can define longitude in three ways: 1. As an arc of the equator or a parallel. 2. As the angle at the pole or the angle at the center between the planes of the prime meridian. 3. As the meridian of a point on Earth. You measure this point eastward or westward from the prime meridian through 180°. Label it E or W to show the direction of measurement. You measure latitude from a standard great circle (the equator). You also use a standard great circle when measuring longitude. This great circle is the meridian. The standard meridian is the prime meridian. By international agreement in 1884, the meridian adopted as the prime meridian was the one on which Greenwich Observatory (near London, England) was located. This is was the 0° longitude. The longitude of a position is also described as being east or west of Greenwich. You can subdivide the degree into smaller units, as in the decimal system. The more 7-5

Figure 7-6 — Easterly magnetic variation. common method of subdivision is to divide each degree into 60 minutes (´) of 60 seconds (´´) each. Another method is to divide the degree into 60 minutes and tenths of minutes. To convert minutes into decimals of degrees, or to convert seconds into decimals of minutes, divide by 6. Thus: 15°30´ = 15.5´, and 15°30´24´´ 15°30.4´. Variation As stated under the definitions of poles, the earth’s true (geographic) poles and its magnetic poles are not at the same locations. Also, the location of the magnetic poles changes slightly over the years. In 1960, the north magnetic pole was at latitude 74.9°N and longitude 101.0°W. The southern pole was at latitude 67.1°S and longitude 142.7°E. Thus, a given line will have a different direction to the true North Pole than to the magnetic North Pole. In addition, lines of magnetic force are not generally straight li nes because of irregular iron deposits near the earth’s surface. Since a compass needle aligns to the lines of force at its location, it may not point to true or magnetic north. The locations on the earth where the compass does point to true north, when connected together, form an irregular line. This is the agonic line. At other locations, the earth’s magnetic field direction may not be the same as the direction of the magnetic poles. The angle between the direction of true north and the direction of the earth’s magnetic field is the location’s variation. This same angle is also often called the angle of declination. You label variation (or declination) east or west as the magnetic field direction is east or west, respectively, of true north. See Figures 7-6 and 7-7. Lines connecting locations having the same variation are isogonic lines.

7-6

Figure 7-8 — Deviation changes with heading. Figure 7-7 — Westerly magnetic variation.

Deviation Deviation is the error in a magnetic compass caused by nearby magnetic influences. These influences may be caused by magnetic material in the structure of the aircraft and to electrical (electronic) circuits. These magnetic forces deflect a compass needle from its normal alignment with the earth’s magnetic field. You express the amounts of such deflections in degrees. The deflection will be east or west as the compass points east or west, respectively, of the earth’s magnetic lines of force. Deviation varies with the heading of the aircraft. Figure 7- 8 shows one reason for this deviation.

7-7

For example, suppose that you represent the net result of all magnetic forces inherent in an aircraft by an arrowhead in the aircraft’s longitudinal axis and aft of the compass. If the aircraft is heading toward magnetic north, the magnetic forces (arrowhead) attract the south-seeking end of the compass needle. However, they don’t change the needle’s direction because the inherent magnetism has the same polarity as the earth’s field. Now, suppose that the aircraft takes an east magnetic heading. The aircraft’s magnetic forces now repel the north end of the compass needle and attract the south end, causing easterly deviation. The figure also shows that the deviation when heading south is zero and when heading west is westerly. You can reduce deviation by changing the position of small compensating magnets in the compass case. However, it is usually not possible to remove all the deviation on all headings. You must determine the residual deviation for each compass installation and record it on a deviation card. The card shows the actual deviation on various headings or, more frequently, the compass headings for various magnetic headings. You can accomplish this using a process known as compass swinging. Compass Error The net result of both variation and deviation is the compass error. If variation and deviation have the same name (east or west), you add to get compass error. If they have different names, subtract the smaller from the larger. Give the difference given the name of the larger. See Figure 7-9. You can label variation and deviation plus (+) if east, and minus (–) if west. In this case the compass error is the algebraic sum of the two. Example 1 Given: Variation 7˚ west (W), deviation 2˚ west (W). Required: Compass error. Solution: 7˚W + 2˚W = 9˚W. To fly a true course of 135˚, this aircraft over this spot on the earth would fly a compass heading of 144˚.

Example 2 Given: Variation (–)2˚, deviation (+)5˚. Required: Compass error. Solution: (–)2˚ + 5˚ = (+)3˚.

7-8

Figure 7-9 — Effect of compass error.

Magnetic Dip At the magnetic poles, the direction of the earth’s magnetic field is vertical (perpendicular to the earth’s surface). Along the aclinic line (sometimes called the magnetic equator), roughly half way between the poles, the field’s direction is horizontal (parallel to the earth’s surface). The difference between the direction of the earth’s field and the horizontal at any location is the magnetic dip. The magnetic dip varies from very small angles near the equator to very large angles near the poles. You can measure the angles with a dip needle, which is a magnetic needle free to turn about a horizontal axis. At San Francisco the dip angle is about 62°. A line connecting all locations having equal dip angles is an isoclinic line.

7-9

The total intensity of the magnetic field is along the dip angle. However you can show it as two components—vertical and horizontal (Figure 7-10).

Only the horizontal component is effective as a directive force for a magnetic compass (wet compass). It loses its effectiveness near the magnetic poles because of the weak horizontal component there. The vertical component causes errors in a magnetic compass during aircraft maneuvers that tilt the compass card east or west. If an aircraft heading east increases its speed, or one heading west decreases its speed, the compass card tilts. A turn to the east from a north or south heading will also tilt the floating compass card In both cases, the east side of the card sinks and the west side rises. The vertical component of the earth’s field causes the compass card to rotate to the east when in the northern hemisphere. It will cause the card to rotate west when in the southern hemisphere. The amount of error is zero at the aclinic line, and it increases toward the magnetic poles. Therefore, precise turns are difficult if referenced to such a compass. Pilotage Pilotage is the directing of aircraft from point to point by visual or radar observation of landmarks. These landmarks are either previously known or recognized from a chart. It is similar to taking a trip by automobile where the highway is the course taken and the Figure 7-10 — The earth’s magnetism. 7-10

towns are the check points. This method has obvious limitations if the flight is made over a large body of water or a poorly charted area, or in darkness, rain or fog. Therefore, whenever possible, use pilotage in conjunction with other methods of navigation. Dead Reckoning Dead reckoning is the process of determining a current position from the record of a previously known position, course, speed, and time traveled. To be accurate, you must consider every change of course and speed during the flight. It does not matter whether the pilot or the air mass (wind) through which the aircraft is flying makes the changes. Radar Navigation Modern radar can be a valuable aid to navigation. Some radars present a map like display of the terrain around the aircraft on the screen of a Cathode-ray Tube (CRT). This allows pilotage to go beyond some of the limitations of visual observations. Radar transponders are devices that do not operate until interrogated or triggered into action by a suitable signal from another radar transmitter. Then, they transmit their own signal, which the interrogating radar receives. These are used both for fixed navigational aids, such as radar beacon stations and for airborne Identification Friend or Foe (IFF) systems. Doppler radar can detect and show actual ground speed and drift of an aircraft, regardless of wind speed or direction. Radar altimeters give the actual distance from the aircraft to the surface below. The surface below can be a body of water or land masses far above sea level. Radio Navigation Radio navigational aids vary from a fairly simple direction-finding receiver to complex systems using special transmitting stations. These special stations make it possible to fix the position of an aircraft with considerable accuracy. The usable range varies according to its intended use and also with weather and ionospheric conditions. Beacon stations associated with an Instrument Landing System (ILS) are usually of low power. Long-range Air Navigation (LORAN ) stations have a range extending to 1,400 miles under favorable conditions. Aviation Electronics Technicians (ATs) maintain the airborne portions of radio and radar systems. Celestial Navigation Celestial navigation is the method of fixing the position of the aircraft relative to celestial bodies. Since the earth is constantly revolving, an accurate time device is necessary. You may use a sextant to measure the angle of the celestial bodies with respect to the horizon. In marine navigation, the visible horizon is the reference. In air navigation, you use an artificial horizon as the reference point. Also, the navigator needs an almanac to determine the celestial equator system coordinates at the time of observation. The usual method to show a line of position from celestial observation consists of (1) observation, (2) coordinate conversion, and (3) plotting. The navigator tries, whenever possible, to select three bodies about 1200 apart in azimuth. This not only results in lines of position that cross cleanly, it also minimizes the effects of a constant error on the observations. 7-11

Inertial Navigation INS are dead reckoning devices that are completely self-contained, which include a computer and platform or module containing accelerometers, gyroscopes, or other motion sensing devices. They are independent of their operating environment, such as wind, visibility, or aircraft attitude. They do not radiate or receive radio frequency RF energy; therefore, they are impervious to countermeasures such as jamming. Gyros measure the angular velocity of the system. When the gyros use the original orientation of the system as the initial condition and integrate the angular velocity, the system’s current orientation is known at all times. Accelerometers measure the linear acceleration of inertial reference that only is measured relative to the moving system. This can be thought of as a passenger of a train feeling pressed back as the train accelerates (whether forward, backward, left, right, up or down. Inertial navigation systems make use of the physical laws of motion that Newton described three centuries ago. One advantage of an INS is it requires no external references in order to determine its position or orientation. Of course, you must provide and enter the starting position into the system. When known positions are available, you may correct or update the system if an error exists. Inputs to the system are from acceleration detectors that measure the rate of change in the motion of the aircraft. The first integral of acceleration is velocity. Velocity results when acceleration is integrated with respect to time. For example, a body starts from rest and constantly accelerates at 8 feet per second for 11 seconds. The velocity at the end of this time would be 88 feet per second (60 miles per hour). However, in actual practice, acceleration is not always this constant. The integration of acceleration is the process of summing all minute acceleration-time increments over a given amount of time. By integrating velocity with respect to time, the result is displacement (distance). Therefore, the second integral of acceleration is displacement. The inertial navigator’s purpose is to keep track of position and not the total distance traveled. This causes the system to integrate all values of acceleration (positive and negative) detected over the time involved. If the earth were flat and vehicles traveled only on the earth’s surface, a two-axis inertial navigation system could plot the position using two accelerometers. One accelerometer would be sensitive along the x-axis (E-W) and the other sensitive along the y-axis (N-S). The important point to note about detecting acceleration of a body is that each accelerometer detects only the component of the resultant acceleration along its sensitive axis. They have no way of telling whether the detected velocity change is due to a speed change or a direction change or both. It does not matter what forces cause the velocity change. Neither can the accelerometer distinguish between the acceleration of the vehicle and the pull of gravity. Therefore, if the accelerometer tilts off its level, its output will include a component of gravity as well as vehicle acceleration. To get the correct vehicle acceleration in the horizontal plane, the sensitive axis of the accelerometer must be perpendicular to the gravitational field. However, the earth is not flat and not exactly round. Its radius at the poles is less than its radius at the equator. It also spins about its polar axis. A spinning gyro in gimbals tries to maintain a fixed direction in relation to space rather than to any point on Earth. Consider a gyro at the equator with its spin vector direction east, toward the morning sun. After 6 hours of Earth rotation, the spin vector would be up in relation to the earth’s surface. After 12 hours it would be west. After 18 hours it would be down; and after 24 hours it would be east again. Now consider a spinning gyro with its spin axis parallel 7-12

with the earth’s axis of rotation. At the equator it is parallel with the earth’s surface. However, as it moves to the North Pole, it becomes vertical to the earth’s surface. You must take all of these items into account and correct for them. So, to navigate on the earth requires a highly complex inertial system, each component of which is capable of extreme accuracy. A four-gimbal system allows the platform to retain the original orientation regardless of what maneuvers the aircraft makes. This allows the platform to serve as a level mount for the accelerometers. The stable platform contains two identical floated, two-degree- of-freedom gyros. They mount with their spin axes horizontal and at right angles to each other. Using the gyroscopic principle of precession, it is possible to apply a continuous torque to the appropriate axes. This action reorients the gyros to maintain the stable platform horizontal to the earth’s surface and pointed north. An electronic analog computer develops the signals necessary to properly torque the gyros. The correct ions for Earth rate depend on the aircraft’s position on the earth’s surface. Some systems use as many as three accelerometers. Two are horizontal with one sensitive to north-south acceleration and the other sensitive to east-west acceleration. The third accelerometer mounts to determine vertical acceleration. A computer subtracts the gravity component from the output of the vertical accelerometer. A more detailed description of an INS follows later in the chapter. AIRCRAFT COMPASS SYSTEMS Countless navigational devices and methods have been invented and devised. In the present era, with its supersonic speeds, accurate determination of direction has become increasingly important. An error of only a few degrees in a space of minutes will carry the modern aviator many miles off course. During the early days of aviation, direction of flight was determined within the aircraft chiefly by direct-reading magnetic compasses. Today the direct-reading magnetic compass still finds use as a standby compass should the more sophisticated compass systems fail. Compass System Sensors and Indicators In chapter 6, the heading indicator is mentioned as a flight instrument. This instrument is part of the primary heading reference system. The heading indicator receives electrical/electronic signals from various components in the system and shows the pilot aircraft heading in degrees. Sophisticated navigation systems and weapons delivery systems require aircraft heading information in electrical/electronic signal form. In this form, the information goes to computers, indicators, and other components. Also, by using these signals, indicators can include aircraft heading along with other information in a single instrument. Compass Transmitter The compass transmitter, commonly called a flux valve, detects the horizontal direction of the flux lines of the earth’s magnetic field. It is usually mounted within the wing or tail of an aircraft as this area has the lowest aircraft-induced magnetic fields (Figure 7-11). It consists of a hermetically sealed hemispherical bowl containing a sensing element in a damping fluid. The bowl permits the sensing element up to 30 degrees of freedom in the aircraft’s yaw and roll axis, while prohibiting rotation about the pitch axis.

7-13

Figure 7-12 — Flux valve heading changes. The core of the flux valve is a good conductor of magnetic lines of force as it has a high permeability. It is shaped like a three-spoke wheel, as shown in Figure 7- 12. Because the flux valve’s core is magnetically sensitive and allowed free movement within the mount, the earth’s magnetic lines of force will cause the core to move into alignment with the earth’s magnetic field. The damping fluid is used to “dampen” the inertial effects of aircraft/mount movement as the aircraft maneuvers. The mount that is attached to the aircraft is in constant alignment with the aircraft’s heading so in effect the mount moves around the core with aircraft heading changes. The core’s legs are connected in a wye configuration, 120° apart, and have copper windings through which an induced current will flow. The amount of flux and current flow in any one leg is proportional to the angular position of the leg relative to the earth’s magnetic lines of flux.

Figure 7-11 — Compass transmitter. 7-14

Figure 7-13 — Compass transmitter and compensator. There is an exciter coil which is wound around the hub of the core, corresponding to the axle of a wheel. This coil receives 400 Hz ac power. This coil is the primary, and the signal pickup coils in the legs are the secondary. The design of the core and windings prevents transformer action between the coils but the purpose of the primary winding and its applied voltage is to produce a magnetic field. This magnetic field changes the reluctance of the core which is driven to saturation at the peak of each positive and each negative portion of the 400 Hz cycle. The core’s inductance properties are also affected by the legs’ position relative the earth’s magnetic lines of flux which further affects the induced current flow through each of the core’s legs. The applied control flux prevents the earth’s magnetic field from entering the core during the saturation highs and lows. When the driving flux is reduced during the transition the earth’s magnetic field enters the core producing an induced signal output. Above the core the mount contains pickup windings. These pickup windings receive an induced voltage from each of the core’s legs. As the mount moves in relation to the core, which tries to maintain alignment with Earth’s magnetic field, the voltage induced into the mount winding provides voltage that is indicative of mount position from the magnetic north relationship of the core. This proportional voltage is used to drive indicators and provide reference for gyroscopic instruments so that heading can be determined relative to magnetic north. All compass systems must be periodically calibrated. Attached to the top of the compass transmitter is a compensator assembly, see Figure 7-13. It consists primarily of two sets of two small permanent bar magnets. You can change the relative azimuth position of each set by rotating a screw on the outside of the unit. These screws position the magnets by a gear train. One adjusting screw adjusts for north-south compensation, and the other screw adjusts for east-west compensation.

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Figure 7-14 — Compass transmitter schematic and functional symbols. Figure 7-15 — Displacement gyroscope. Two wiring symbols for the flux valve are shown in Figure 7-14. To distinguish them from synchro units, the words “compass transmitter” or “flux valve” are usually included in the drawing.

Displacement Gyroscope Assembly The displacement gyroscope (Figure 7-15) is a hermetically sealed, two-gyroscope platform providing pitch, roll, and azimuth signals to the system indicating instruments. The vertical gyroscope provides a source of pitch and roll information, while the directional gyroscope provides azimuth (yaw, heading) information. Control transmitters (synchros) convert attitude changes into electrical signals that represent pitch, roll, and yaw. Erection circuits maintain the spin axis of the vertical gyroscope in a gravity vertical position. These circuits consist of a roll electrolytic switch, with associated roll torquer, and a pitch electrolytic switch, with associated pitch torquer. A servo loop maintains the spin axis of the directional gyroscope level. The loop consists of a leveling pickoff, an external leveling amplifier, and a leveling torquer. The vertical gyroscope spin motor mounts in the inner roll gimbal, which is free to move in a roll direction. The freedom of movement in the roll direction is limited to ±82° by mechanical stops to prevent gyroscope gimbal lock. Gimbal lock would 7-16

occur if the vertical gyroscope spin axis were to become aligned with the vertical gyroscope pitch gimbal. The directional spin motor mounts in the directional gyroscope leveling gimbal. This gimbal is free to move in a roll direction. Mechanical stops also restrict the freedom of the leveling gimbal to ±82°, preventing gimbal lock. These stops also prevent the inversion of the leveling and inner roll gimbal during gyroscope rotor coastdown after system shutdown. The displacement gyroscope assembly consists of a Vertical Gyro (VG) and a Directional Gyro (DG). These gyros mount in a common outer roll gimbal. The vertical gyro provides pitch and roll signals and the directional gyro provides heading (azimuth) signals. Erection and leveling servo loops erect and maintain the spin axis of the vertical gyro gravity vertical and the directional gyro spin axis parallel to the earth. Roll, pitch, and azimuth control transmitters convert aircraft attitude and heading into electrical signals. The outer roll gimbal is the outermost gimbal for both the vertical and directional gyroscopes. Pitch control transmitters detect pitch movement of the outer roll gimbal about the vertical gyroscope pitch gimbal. The outer roll control transmitter and directional gyroscope control transmitter mount between the outer roll gimbal and gyroscope case (frame). However, they mount at opposite ends. They sense roll movement of the aircraft (and frame) about the outer roll gimbal. The azimuth gimbal may settle at any random heading during power application. Therefore, you must slave the directional gyroscope’s initial azimuth signal to the flux valve. Also, you can correct it manually to a known magnetic heading. You must do this so the azimuth signal reflects actual aircraft heading. The two azimuth control transmitters, between the azimuth gimbal and the directional gyroscope pitch gimbal, will then furnish information on any further change in aircraft heading. The displacement gyroscope incorporates snubbers. They maintain the approximate normal position of the outer roll gimbal and the directional gyroscope pitch gimbal when power is removed. Upon application of power, the snubbers energize, removing their snubbing action. The motor-generator and gear assemblies that drive the gyroscope gimbals have enough power to override the snubbing action should a snubber failure occur.

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Figure 7-16 —Typical displacement gyro and servo loops. Vertical Gyroscope Operation Refer to Figure 7-16 as you read this section. The vertical gyro consists of gyro spin motor B101 (which is the inner roll gimbal) and the vertical gyro pitch gimbal. It also includes the outer roll gimbal and the frame. The frame mounts to the assembly case and follows all aircraft maneuvers. The outer roll gimbal mounts in the frame. It may rotate 360° about the roll axis but follows the aircraft in pitch and yaw.

The vertical gyro pitch gimbal mounts in the outer roll gimbal. This gimbal may rotate 360° about the pitch axis but follows the outer roll gimbal movements in roll and yaw. The gyro spin motor may rotate ±85° in roll but follows the vertical gyro pitch gimbal in pitch and yaw. Mechanical stops (not shown) limit inner roll gimbal movement to prevent B101’s spin axis aligning with the vertical gyro pitch gimbal axis. Such an alignment would cause the vertical gyro pitch gimbal to spin about its pitch axis (gimbal lock). Leveling – At power application, a friction brake (snubber) releases the outer roll gimbal from the frame. The gyro spin motor starts, and electrolytic switches sense unlevel conditions in pitch and roll. The output of the electrolytic switches activates the torquers. The gyro reacts to the applied torque and precesses until the electrolytic switches are level. The inner roll control transmitter is mounted between the vertical gyro pitch gimbal and the inner roll gimbal. This transmitter applies signals to the roll servo amplifier to 7-18

drive the roll motor-generator. The roll motor-generator, in turn, drives the outer roll gimbal to the level of the inner roll gimbal. Pitch Sensing – As the aircraft pitches, the outer roll gimbal follows, but the vertical gyro pitch gimbal remains level. The pitch servo control transmitter detects the pitch attitude and applies pitch signals to the indicators and other aircraft systems. The pitch control transmitter applies pitch signals to the Automatic Flight Control System (AFCS) control amplifier. Roll Sensing – As the aircraft rolls, B101 remains level, but the vertical gyro pitch gimbal rolls (with the outer roll gimbal). The inner roll control transmitter senses the difference. It then causes the roll amplifier to drive the roll motor-generator until the outer roll gimbal is level with B101. The outer roll control transmitter (on front end of frame) detects and applies roll signals to indicators and other systems. The roll control transmitter (on aft end of frame) applies roll signals to the AFCS control amplifier. Directional Gyroscope Operation The directional gyro consists of gyro spin motor B201 (including a leveling gimbal), an azimuth gimbal, and a directional gyro pitch gimbal. The directional gyro pitch gimbal mounts in the outer roll gimbal. The pitch gimbal may move 360° about the pitch axis, but it follows the outer roll gimbal in roll and yaw. The azimuth gimbal mounts in the directional gyro pitch gimbal. It may move 360° about the yaw axis; however, it follows the directional gyro pitch gimbal in pitch and roll. B201 mounts in the azimuth gimbal. B201 is limited to ±85° by mechanical stops (not shown) to prevent gimbal lock. Leveling – The leveling control transmitter output goes to the leveling amplifier, which drives the leveling torquer. When the leveling torque moves the azimuth gimbal, B201 precesses until the leveling control transmitter senses a level condition. The directional gyro pitch gimbal is servoed to the vertical gyro pitch gimbal and maintained perpendicular to the surface of the earth. The pitch servo control transmitter output, through the pitch servo control transformer, is amplified and drives the pitch follow-up motor-generator. The motor-generator positions the directional gyro pitch gimbal. Azimuth Sensing – The azimuth gimbal may settle at any random position in yaw. The only forces acting on the gimbal are gyro rigidity, apparent (Earth rate) precession, and the leveling torquer. Azimuth sensing in the directional gyro operating mode is reliable only after setting the correct heading into the system with the SET HDG control. Two azimuth control transmitters sense any movement of the directional gyro pitch gimbal about the azimuth gimbal. The yaw signal of one azimuth control transmitter goes to the attitude indicator. The yaw signal of the other azimuth control transmitter is processed in the compass adapter-compensator and applied to other aircraft systems.

7-19

Figure 7-17 — Horizontal indicator. Figure 7-18 — Horizontal situation indicator.

Aircraft Horizontal Flight Indicators Horizontal Indicator (HI), shown in Figure 7-17, sometimes called Horizontal Situation Indicator (HSI), shown in Figure 7-18, is used primarily to provide visual information concerning aircraft attitude, steering, and navigation on a CRT display or mechanical dial indicator. The symbols to make up the display are generated by a separate computer or dedicated Electronics Control Amplifier (ECA). Depending on the aircraft this indicator can also provide steering and navigation information with a Tactical Aircraft Moving Map Capability (TAMMAC) superimposed over the display. The internal Digital Map Computer (DMC) controls the map positioning through map data loaded into a Memory Unit (MU) with Personal Computer Memory Card International Association (PCMCIA) cards. Communication between the computer, MU and HSI occurs over the High Speed Interface Bus (HSIB) or Multiplex Bus (MuxBus).

The HSI system functions as a selectable display for the navigational systems on an aircraft. Selectable between INS, Tactical Air Navigation- TACAN or ILS systems, the corresponding system will provide radial bearing and bearing information, NAV flag, course deviation, and To-From signals to the HSI. The bearing signal will position the bearing pointer 1 or bearing pointer 2, as selected. Magnetic and true heading information is supplied to the HSI compass card by the INS, AHRS, or a compass transmitter, as selected. Magnetic heading is normally used to position the compass 7-20

Figure 7-20 — Amplifier power supply. Figure 7-19 — Bearing-distance-heading indicator. card. When a tactical mode is selected, magnetic heading information is switched out of the circuit, and the compass card will be driven by the true heading information. The Bearing-Distance-Heading Indicator (BDHI) is similar to HI and HSIs and may be used with various navigational systems, and it provides information according to the mode selected. Some aircraft may have more than one BDHI, with separate select switches for each instrument. The distance counter numerals may be in a vertical row or horizontal, as shown in Figure 7-18. The lubber index is a fixed reference mark that allows the operator to read the heading from the BDHI. For detailed operation of the HI, HSI and BDHI systems, refer to the applicable Maintenance Instructions Manual (MIM).

Amplifier Power Supply The amplifier power supply (Figure 7-20) contains the following components:  An ac and a dc power filter  Two power supply modules  A roll driver amplifier  Roll, pitch, and leveling servo amplifiers  A leveling modulator  A servo failure monitor module  Two thermal relays  Ten control relays

7-21

The amplifier power supply provides the timing, switching, and voltages required for the start cycle; erection voltage control during system operation; monitoring roll and pitch signals and the neutral power lead of the vertical gyroscope motor; and the roll, pitch, and leveling amplifiers required for control of the displacement gyroscope gimbals. Aircraft three-phase power goes to the amplifier power supply. It then routes through the ac power filter. The filtered three-phase power then goes to other system components. A power supply connected to all three legs of the three-phase power provides 28-volt dc, which goes to a filter. System components use both filtered and unfiltered dc. The second power supply is a three-section supply. Each section is independent of the other s and supplies 95 volts dc. The dc outputs connect to various control relays within the amplifier power supply. The output from the displacement gyroscope photoelectric pickoff is a dc voltage. A leveling modulator converts it to a 400-hertz ac voltage. The amplitude and phase of the modulator output depends on the amplitude and polarity of the dc voltage from the photoelectric pickoff. The 400-hertz ac voltage goes to the leveling amplifier; then the amplified signal goes to the directional gyroscope leveling torquer control winding. Roll and pitch error signals from the displacement gyroscope are amplified by the roll and pitch amplifiers. After amplification, the signals go to the displacement gyroscope for application to the roll and pitch motor-generator control windings. The roll and pitch motor-generators drive the displacement gyroscope roll and pitch gimbals. Two thermal relays and nine control relays perform the timing and switching required for the start cycle. The start cycle is of 60 seconds duration. However, certain conditions change after the first 12 seconds. High pitch erection voltage goes to the pitch torquer for the complete start cycle. High roll erection voltage goes to the roll torquer after 12 seconds and continues until the completion of the start cycle. After the first 12 seconds, motor excitation voltage increases, and the gyroscope motors attain operating speed. After completion of the start cycle, two additional relays provide roll and pitch erection cutout during specific aircraft maneuvers. Compass Adapter Compensator The compass adapter compensator (Figure 7-21) receives heading information from the flux valve and the displacement gyroscope. It processes this information according to the azimuth mode selected by the compass controller. It then provides corrected heading signals to the heading indicator and other aircraft systems. The compass adapter compensator can operate in the following modes: free, compass, and slaved.

7-22

Figure 7-21 — Compass adapter compensator.

In the free mode of operation, you engage the PUSH TO TURN control on the compass c ontroller to set actual aircraft heading. This establishes an initial azimuth reference. As aircraft heading changes, an azimuth synchro on the directional gyro measures the relative change between the aircraft and the directional gyroscope. The signal goes to the compass adapter, which makes corrections for real and apparent drift. The compensating signal for apparent drift is derived from a resolver in the compass controller. The EARTH RATE CAL variable resistor in the compass adapter adjusts this signal. Real drift is caused by mechanical imperfections in construction of the directional gyroscope. The compensating signal for real drift develops across the compass adapter GYRO DRIFT COMPENSATION POT variable resistor. This resistor has a dial calibrated in degrees per hour. The corrected information goes to external aircraft system components through five heading repeater synchros. When operating in the compass mode, the 24-point compensation network corrects the flux valve signal for deviations. This signal goes as an error signal to the azimuth servo loop, resulting in corrected azimuth information (angle data shaft). Again, this information goes to external aircraft system components through five heading repeater synchros. In the slaved mode, the directional gyroscope azimuth information and a compensated flux valve heading correction signal go to a differential synchro in the compass adapter. This slaves or synchronizes the gyroscope azimuth output to the flux valve heading, providing a heading output rather than a displacement output. Fast synchronization starts when the slaved mode is selected and is maintained until close alignment is achieved. Slow synchronization is applied continuously while operating in the slaved mode. 7-23

Figure 7-22 — Compass controller. Compass Controller The compass controller (Figure 7-22) provides switching functions, latitude compensation signals, and slew signals to the system compass adapter. The compass controller also monitors and provides a visual display of the synchronization between the azimuth heading output and the flux valve. The compass controller can operate in three modes—slaved, free, and compass. The compass controller contains a PUSH TO SYNC switch and SYNC IND meter. It also has a PUSH TO TURN control, mode switch (with COMP, SLAVE and FREE positions), and LATITUDE DEGREES control (counter assembly). The mode switch selects the FREE (free), SLAVE (slaved), and COMP (compass) modes of operation. To accomplish this, it controls the mode- selecting relays in the system compass adapter. The mode switch also activates the SYNC IND meter during compass and slaved modes. The PUSH TO TURN control (set heading) provides switching to decouple the autopilot. Also, it controls the direction and rate of slewing for alignment of the system to an azimuth heading. This switching action occurs when using the PUSH TO TURN control to reference the system output to the aircraft heading in the free mode. The PUSH TO SYNC switch provides switching to synchronize azimuth heading output to the flux valve when operating in the slaved mode. The LATITUDE DEGREES control, working with a resolver, provides a compensation signal for apparent drift of the directional gyroscope. Apparent drift results from Earth rotation. The hemisphere switch (N, S) selects the latitude correction signal for the Northern or Southern Hemisphere. The SYNC IND meter shows the synchronization between the azimuth heading output and the flux valve. The slaved mode is the normal operating mode except when in an area where the earth’s magnetic field is distorted. The slaved mode synchronizes the directional gyroscope output to the flux valve heading. When initiating the slaved mode, fast synchronization occurs until close alignment with the flux valve heading is achieved. Free mode is an alternate mode of operation. It is used in areas where the earth’s magnetic field is distorted. When operating in the free mode, only the directional gyroscope output drives the system’s azimuth indicators. 7-24

Figure 7-23 — Switching rate gyroscope. The compass mode is an emergency mode for use when the directional gyroscope fails. Only the flux valve output (compensated) provides heading information. Rate Gyroscope The switching rate gyroscope (Figure 7-23) provides a means of interrupting the roll erection and slaving voltage. It is a single-degree-of-freedom gyroscope. It provides gyroscope sensitivity to rates of rotation about the yaw axis of the aircraft. When the aircraft turns at rates of 150° per minute or greater, the gyroscope precesses away from th e normal condition. This causes the contacts of a magnetic reed switch to close, energizing a single-stage amplifier. Transistor switching action pulls in a relay, completing the 28-volt dc path to the turn cutout relay coil in the compass adapter.

Attitude Indicator The attitude indicator (Figure 7-24) is a three-axis, servo-driven sphere that shows heading and relative roll and pitch attitude of aircraft. Vertical and horizontal pointers provide the pilot with aircraft deviation information from a desired flight path. Signals for operating servo systems of hermetically sealed units are from the displacement gyroscope. Aircraft rate of turn information and vertical displacement deviations from a desired glide path are displayed by a rate of turn pointer and displacement pointer, respectively. Loss of ac power is indicated by a display of a power failure warning flag. Inadequate current to vertical, horizontal, and displacement pointers results in display of respective warning flags. A pitch trim knob lets you adjust the sphere to varying aircraft configurations and reduce parallax error.

7-25

Figure 7-24 — Attitude direction indicator.

INERTIAL NAVIGATION SYSTEM The following paragraphs describe a heading reference system found on the latest model high-speed aircraft and some patrol aircraft. The INS is sometimes maintained by the Aviation Electronics Technician (AT) rating. Some squadrons use a concept called an Integrated Weapons Team (IWT). It is composed of the three Avionics/Armament Division (Work Center 200) ratings— AT, AE, and Aviation Ordnanceman (AO). Regardless of who maintains the INS, you, must be familiar with the theory and operating principles of such a system. Basically, navigation can be divided into two categories: (1) position fixing and (2) dead reckoning. In the first category, you determine your position relative to positions of known objects such as stars and landmarks. The most common example of navigation by position fixing is celestial navigation. Use of loran is another example of navigation by periodic position fixes. Dead reckoning, the second category, is the process of estimating your position from the following known information:  Previous position  Course  Speed  Time elapsed

Two examples of navigation by dead reckoning are Doppler radar and Inertial Navigation Systems. All navigation systems, except the inertial type, rely on some bit of information external to the vehicle to solve its navigational problem. In this respect, the inertial navigation system stands alone; it is completely self-contained within the vehicle. It is independent 7-26

NOTE The word nearly is emphasized because the object will deviate slightly from its straight-line motion. The cause of this deviation is the earth’s rotation about its axis. of its operating environment, such as wind, visibility, or aircraft attitude. It does not radiate RF energy; therefore, it is impervious to countermeasures. It does not depend on ground transmission or any other outside source to determine its instantaneous position. Basic Principles The operating principle of the INS is Newton’s first law of motion. This law states “Every body continues in its state of rest, or of uniform motion in a straight line, unless it is compelled to change that state by forces impressed on it.” In layman’s terms, this law says that a body at rest tends to remain at rest. It also says a body in motion tends to remain in motion, unless acted upon by an outside force. The full meaning of Newton’s first law is not easy to visualize in the earth’s reference frame because Newton’s laws apply to an inertial reference system. You may define an inertial reference system as a nonrotating coordinate frame. It can be either stationary or moving linearly at a uniform speed, in which there are no inherent forces such as gravity. You can make a simple test of whether you are in a true inertial system by releasing an object and observing its motion. If you release the object without imparting any acceleration to it, the object remains in its position relative to you. If you throw the object, it continues on an undeviating path at a constant speed. Such a system can exist only in empty space, far from any mass, for all masses contain gravitational forces. A reference system attached to the earth can closely approximate an inertial system. For this system to work, you must balance the gravitational force on a body by a second force. For example, an object sliding on a flat, frictionless plane on the earth’s surface would move in a NEARLY straight line. The object will have a NEARLY constant speed, as you saw in the earth’s coordinate system.

Newton’s second law of motion shares importance with his first law in the inertial navigation system because the inertial navigation system works on Newton’s second law. Newton’s second law of motion states “Acceleration is proportional to the resultant force and is in the same direction as this force.” Thus, the second law is written F = ma where; F = force m = mass a = acceleration

The physical quality in the above equation that pertains to the inertial navigation system is acceleration. You can derive velocity and displacement from acceleration. For 7-27

example, consider this fact: Before an object can change its state of rest or state of motion, it must first experience acceleration. Since acceleration is a change in velocity and velocity is a change in position, acceleration is a change in the change of position. However, before any change can have meaning, it must include the unit of time. Therefore, you can define a change per unit of time as a rate of change. Thus, a rate of change of displacement is velocity. A rate of change of velocity is acceleration. A rate of change of a rate of change of displacement is acceleration. Differentiation is the process of investigating or comparing how one physical property varies with respect to another. Integration, the reverse of differentiation, is the process of summing all rate of changes that occur within the limits under investigation. The inertial navigation system is not a differentiating system; it is an integrating system. However, before integration can be done, it must first have a rate of change. Therefore, the inertial navigation system, when stripped to its barest essentials, is a detector and an integrator. It first detects changes of motion. It then integrates these changes of motion with time to arrive at velocity, and again with time to arrive at displacement. Fundamentals of Integration Since an INS performs integration, the following is a review of integrating principles. The equations for the integrals of acceleration and velocity are: ∫a dt = v ∫v dt = s ∫∫a dt dt = s where, s = displacement v = velocity a = acceleration ∫ = integration symbol dt = time differential

When acceleration (a) is integrated (∫) over a specific period of time (dt), the result is velocity (∫a dt = v). When velocity (v) is integrated (∫) over a specific period of time (dt), the result is velocity displacement (s). Therefore, when acceleration (a) is integrated twice (∫) over a specific period of time (dt²), the result is displacement (s). Remember from elementary physics that acceleration, whose units are ft/sec², multiplied by time in seconds is velocity in ft/sec. Also, that velocity (ft/sec) multiplied by time (sec) is displacement (ft). The integration of acceleration, for example, is the mathematical process of summing all minute acceleration-time increments over a given period. The result of the integration of acceleration is velocity over the same period. The same integration process performed on velocity gives displacement or distance traveled over the same period.

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Figure 7-25 — Simple single-axis INS block diagram. Simple Single-Axis Inertial Navigation System An example of how a simple single-axis INS operates is illustrated as follows: Assume a person is on an INS-equipped train on railroad tracks at the equator. The tracks run in a straight line east and west only. The INS consists of an acceleration detecting device (accelerometer), an integrating device, and a displacement readout device. The accelerometer can sense movement in only one direction, along its sensitive axis. The sensitive axis is an imaginary line parallel to the movement of the mass within the detecting device. The acceleration detecting device is oriented in the train so that it detects accelerations when the train is movi ng forward or backward. Figure 7-25 is a block diagram of such a device. If the train starts moving at point A, you will note a specific reading on the displacement readout device. When the train reaches point B and stops, the readout device will show the new position. The distance traveled from point A to point B added to the reference value noted at point A will show on the displacement indicator. The train returns to point A by traveling backwards. Thus, the simple inertial device is not disoriented. At point A the readout device shows the value that was chosen as a reference. This is the displacement at point B minus the distance traveled from point B to point A. Figure 7-26, view A, is a graph of the detected acceleration. View B is the velocity curve obtained by integrating the acceleration curve shown in view A. View C is the displacement curve obtained by integrating the velocity curve shown in view B. All three curves are plotted as a function of time. The acceleration curve (Figure 7-26, view A) begins at time as the train begins to travel from point A. Look at view C. The acceleration at time has a value of , and it remains at that value until . At the train ceases to accelerate. Therefore, acceleration goes to zero. At this point, the train reaches a steady velocity. The train continues traveling at a constant velocity until time where the train begins to stop. The acceleration detector detects an acceleration equal in value to , but its direction is opposite. This acceleration is constant from time to time . At the acceleration goes to zero. The train is now stationary and standing at its destination—point B.

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Figure 7-26 — Integration of acceleration and velocity: (A) acceleration, (B) velocity, (C) displacement. Look at the velocity curve for the time interval to (Figure 7-26, view B). It is the result obtained when acceleration is integrated over the same interval. The velocity curve is the output of the first integrator from to . During the interval to , velocity is changing in an increasing or positive direction. This means that a positive acceleration is taking place. Velocity is constant during interval to , which means acceleration is zero. At time , velocity begins to decrease. This says that acceleration is again taking place. In this case the acceleration is negative. At time , both acceleration and velocity are zero. The purpose of an INS is to keep track of position and not total distance traveled. To do this it integrates all values of acceleration (positive and negative) detected over the interval. Therefore, it is the net value of acceleration that interests the INS. For instance, in the interval to , all accelerations that occur over the interval are summed, giving a net value at time . In this case, integration of acceleration (Figure 7-26, view A) is the process of summing the area bounded by the acceleration curve and the time axis. The area above the time axis is positive, and the area below the time axis is negative. Since the areas above and below the time axis are equal, the net value for interval to to acceleration is zero. The integral of acceleration for the interval to to is therefore zero. This means that the velocity at time is equal to the velocity at time to, in this case zero. Integrating velocity from time to is the job of the second integrator. It gives B units of displacement on the displacement axis at time . The displacement readout device changes continuously as long as the second integrator produces an output. The second integrator ceases to produce an output when the first integrator (velocity) ceases to produce an output. The velocity integrator continues to produce until receiving an acceleration that balances out the initial acceleration. At this point, it produces a net acceleration of zero. The readout device stops at the point where the net acceleration is zero. Until reaching this condition, the readout device shows a continuous change in displacement. 7-30

Figure 7-27 — Two-axis inertial navigation system, block diagram.

The return trip is described as follows. The train is at point B during time interval to ; it begins traveling backwards to point A at time . The acceleration detector senses an acceleration – , which is negative and slightly less than the previous acceleration, - . At time it reaches a steady velocity, and acceleration goes to zero at this point. Note that velocity is now negative since the direction of travel is reversed. Since the size of acceleration – is less than that of , maximum velocity on the return trip is less. Therefore, the time required to return to point A is greater. Interval to , greater than time interval to , reflects this fact. The train begins to stop within a short distance of point A. This happens at time , producing an acceleration of as sensed by the acceleration detector. The train comes to a full stop at time . Here the detector senses zero acceleration. Since the net acceleration over the interval is again zero, the output of the first integrator (velocity) is zero. The second integrator (displacement) output stops with the displacement readout device showing the reference value. This value is the same originally noted at reference point A. The simple single-axis INS just described will detect and compute all changes in displacement. However, the acceleration detector (accelerometer) must retain its straight-line orientation. Also, all motion must be along a straight line passing through the reference or initial point. Obviously, using this simple INS, a person must navigate along a straight line. Two-Axis Inertial Navigation System Suppose, for example, that the earth is flat. If so, you can determine position by using a system of coordinate axes. This system of coordinates uses two sets of parallel lines (x and y). One set of lines is perpendicular to the other set of lines. These lines form a grid network over the earth’s surface. If you use two single-axis inertial navigation systems, you can determine position on the plane (flat surface). You simply maintain proper orientation of each accelerometer’s sensitive axis relative to the coordinate system. One accelerometer mounts on a platform so that its sensitive axis lies along the x-axis. The other accelerometer mounts on the same platform so its sensitive axis lies along the y-axis. This will maintain their axes mutually perpendicular. The accelerometers will then sense any rate of change of velocity along the coordinate axes. Figure 7-27 is a block diagram of a simplified two- axis inertial navigation system. 7-31

Figure 7-28 — Two-axis inertial platform in a plane coordinate system.

Figure 7-28 is an illustration of the inertial platform mounted on a vehicle moving over a plane coordinate system. Note that the platform and accelerometers remain oriented with the coordinate axes regardless of vehicle heading. The vehicle’s ground track represents the vehicle’s displacement over the grid system. You can locate the vehicle at any given time by the x and y coordinates. You plot the x-displacement left to right, and the y-displacement is top to bottom on the page. You reference time to the x-axis. Figure 7-29 is an illustration of a typical set of acceleration and velocity curves from the INS shown in Figure 7-28.

Referring to Figures 7-27 and 7-28, the operation of the plane inertial navigation system is explained as follows: The vehicle aligns (initializes) on the coordinate system with a displacement of 3 on the x and y-axes. That is, both x and y displacement indicators read 3. At time , the vehicle experiences an acceleration, A, in a direction of 45° from the x-axis. The accelerometers detect only that portion of the acceleration that lies along its sensitive axis. This means the x accelerometer detects the component of acceleration along the x-axis. This is A cos a. The y accelerometer detects the component of A along the y-axis, which is A sin a. The vehicle continues in a direction of 45° until time . At this time it begins a turn to the right. Since sine and cosine are 7-32

Figure 7-29 — Acceleration and velocity curves. equal at an angle of , the displacements along x and y are equal at time . This says x = 15 and y = 15. Also, it means the acceleration and velocity along the x-axis is equal to the acceleration and velocity along the y-axis. At time , the vehicle begins a right turn; it completes the turn at time . The new direction is parallel to the x-coordinate and perpendicular to the y- coordinate. By looking at Figure 7-28, you see interval to shows the x-accelerometer detecting a positive acceleration. It also shows the y accelerometer detecting a negative acceleration during this interval. If the vehicle maintains a constant speed throughout the turn, the detected acceleration results from a velocity change. This change is due to a change in direction rather than a change in speed. This acceleration is radial (centripetal) acceleration ( ). The direction of the radial acceleration is toward the center of the turn and perpendicular to tangential velocity ( ). Find coordinates (17.5, 17) in Figure 7-28. If the speed hadn’t been constant during the turn, a tangential acceleration ( ) would have occurred. This acceleration would parallel the tangential velocity ( ) vector and be normal (at a right angle) to the radial acceleration vector ( ). The direction of the tangential acceleration would depend upon whether the speed was increasing or decreasing, positive or negative. If the turning vehicle’s acceleration is due to changes in speed and direction, the accelerometers detect the x and y components of the resultant of the two accelerations. Remember, when detecting acceleration of an accelerating body, accelerometers detect only the component of the resultant acceleration along their sensitive axis. Accelerometers can’t tell if the detected velocity change is due to a speed change or a direction change or both, nor does it matter what forces cause the velocity changes. The result is the same, provided the accelerometers maintain correspondence with the coordinate axes. Refer to the acceleration and velocity curves in Figure 7-28. Notice the integration of the x-component of acceleration for the interval to . It shows an increase in the x- component of velocity and, therefore, a corresponding increase in displacement along the x-axis. Integration of the y-component of acceleration over the same interval shows that the velocity goes to zero at time . Therefore, the displacement along the y-axis ceases to change. Hence at time , the displacement is (15, 15). At time , the displacement is (20, 15) and at time , the displacement is (25, 15), etc. The INS just described navigates very well on a flat surface. However, navigation on the earth requires a highly complex inertial system. The earth, of course, is not flat, neither is it exactly round. Its radius at the poles is less than its radius at the equator. It also 7-33

spins about its polar axis and orbits around the sun. You must take all of these factors into account and correct them (except the earth’s motion in orbit around the sun). Once you accomplish these corrections, you may navigate on the earth by inertial means. The earth’s motion about the sun does not affect an Earth inertial navigation system. This motion is translational, and it is equal at all points on the earth. Basic System Components The inertial navigation system continuously measures aircraft accelerations to compute aircraft velocity and change in present position. These measurements are made by precision inertial devices mounted on a three-axis stable element, which is part of a four-gimbal structure. The four-gimbal structure allows the stable element to move with 360 degrees of freedom about the three axes. Two gyros provide gimbal stabilization signals to maintain the stable element level with the earth’s surface and aligned to true north. Also, the system uses these signals to measure aircraft pitch-and-roll attitudes. The inertial characteristics of the gyroscopes used in the system define and maintain the reference axes for relatively long periods with great accuracy. With a gyrostabilized platform as a reference, it is possible to accurately detect components of motion in any direction. To do this, precision accelerometers and analog or digital computers are used in an INS. Accelerometers The primary data source for the inertial navigation system is the accelerometer. Three accelerometers are mounted on the stable element between the gyros. They provide output signals proportional to total accelerations experienced along the three axes of the stable element. The system uses these accelerations to produce aircraft velocities and changes in position. An accelerometer consists of a pendulous mass that is free to rotate about a pivot axis in the instrument. Figure 7-30 shows one form of this device. It has an electrical pickoff that converts the rotation of the mass about the pivot axis to an output signal. An acceleration of the device to the right causes the pendulum to swing to the left. This provides an electrical pickoff signal, which causes a torquer to restrain the pendulum. The pickoff signal goes to a high gain amplifier. The output of the amplifier connects to the torquer on the accelerometer. During an acceleration, this feedback loop sends a voltage to the torquer. This voltage holds the pickoff signal at a null under the influence of the measured acceleration. This voltage is proportional to the measured acceleration. It also provides the electrical output acceleration signal that goes to the computer. Figure 7-30 — Typical torque-balanced accelerometer. 7-34

Figure 7-31 — Principle of an accelerometer: (A) accelerometer at null, (B) true acceleration, (C) spurious acceleration due to gravity.

The accelerometer (Figure 7-31, view A) cannot distinguish between the acceleration of the vehicle and gravitational acceleration. Therefore, if the accelerometer tilts off level, its output includes a component of gravitational acceleration as well as vehicle acceleration. Look at Figure 7-31, view C. To get the correct vehicle acceleration in the horizontal plane, hold the sensitive axis of the accelerometer normal to the gravitational field. Refer to Figure 7-31, view B. The accelerometer mounts on a platform (stable element) in a way that it is always level. In this position the accelerometer measures true aircraft acceleration in a horizontal direction along its sensitive axis. Mounting another level accelerometer perpendicular to the first one gives you the x and y-axes. The system can now determine total true acceleration in a horizontal plane for any movement in any direction.

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Figure 7-32 — Analog integrating device. Figure 7-33 — Basic inertial navigation system. Integrators To convert the measured acceleration to aircraft position information, the system processes the acceleration signals to produce velocity information. It must then process the velocity information to derive distance traveled. Figure 7-32 shows an analog type integrator. It is an electromechanical device that receives electrical input (acceleration or velocity) and produces a shaft speed proportional to the input. The shaft angle is the output of the integrator, and it is the mathematical integral of the input. If the input is acceleration, the output is velocity; if the input is velocity, the output is distance.

If one of the horizontal accelerometers points north, the other one will always point east. By connecting the accelerometer outputs to integrators (Figure 7-33), the system can determine distance traveled in the north-south and east-west directions. It is important to maintain the proper accelerometer pointed north and maintain both accelerometers horizontal to the earth’s surface. If the accelerometers tilt off level, it measures gravitational components, which results in navigation errors. A third accelerometer sometimes mounts on the stable element in the vertical plane to determine vertical acceleration. The computer subtracts the gravity component from the output of the accelerometer. The resulting signal represents actual aircraft vertical acceleration. A vertical acceleration signal goes to an integrator in the attitude computer. This computer computes vertical velocity.

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Figure 7-34 — Simplified platform stable element. Platform Stable Element To maintain the proper orientation of the accelerometers, they mount on a stable element together with gyroscopes. The gyroscopes are the sensing elements for controlling the orientation of the stable element. The stable element (Figure 7-34) mounts on gimbals, which isolate it from angular motions of the aircraft. GYROSCOPES – The stable element contains two identical, floated, two-degree-of- freedom gyroscopes. They mount one on top of the other in a dumbbell configuration (Figure 7-34). The gyroscopes have their spin axes horizontal and at right angles to each other. The wheels in these gyroscopes, which spin at high speed, resist any effort to change the orientation of their spin axes.

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Figure 7-35 — Single-axis, gyrostablized platform.

Figure 7-35 shows a two- degree-of-freedom gyro and a single-axis stable platform. The pickoffs on the gimbals within the gyro produce electrical signals. These signals occur when the gyro case moves from its null position with respect to the gyro motor. The electrical pickoffs will sense any displacement of the stable element from the frame of reference. The signals thus created drive the platform gimbals to realign the stable element.

PLATFORM GIMBAL STRUCTURE – Figure 7-34 shows the four-gimbal platform configuration actually used in an inertial navigation system. The stable element mounts in the gimbal structure so that, regardless of aircraft maneuvers, it retains the original orientation. The stable element serves as a level mount for the accelerometers. An azimuth gimbal lets the aircraft change heading without affecting the orientation of the stable element. A pitch gimbal removes the effect of aircraft pitch, and a roll gimbal stops the effects of roll. An extra roll gimbal prevents the occurrence of gimbal lock during certain aircraft maneuvers and makes the system truly all-attitude. Look at Figure 7-36. Note the inner roll gimbal that prevents gimbal lock, which would cause the stable element to tumble. Gimbal lock occurs when two of the gimbal axes become aligned parallel to each other. This causes the stable element to lose one of its degrees of freedom. When the aircraft exceeds 90° in pitch, the outer roll gimbal rotates through 180°. The gimbals are oriented so the system may sense aircraft attitude and heading by measuring angles between the gimbals. Synchros send this information to the attitude indicator and other systems in the aircraft.

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Figure 7-36 — Gimbal flipping action.

PLATFORM ORIENTATION – Figure 7-37, view A, shows the apparent rotation of a stabilized platform located at the equator. As shown, the platform will remain fixed with respect to inertial space. However, it appears to rotate about the surface of the earth as the earth spins about its polar axis. This is undesirable for navigation since the accelerometers will not remain horizontal to the earth’s surface. Consequently, this produces gravitational components of acceleration in the outputs of the accelerometers.

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Figure 7-37 — Earth rate torquing: (A) without gyro torquing; (B) with gyro torquing. Figure 7-38 — Aircraft rate torquing: (Frames 1 and 2) without torquing; (Frames 3 and 4) with gyro torquing.

Consider what happens to a stable element as the aircraft flies over the surface of the earth. As the aircraft flies straight north from the equator to the North Pole, the aircraft sees a continuing pitch maneuver. Look at Figure 7-38, frames 1 and 2. At the pole, instead of the platform being level with the surface of the earth, it is now 90° off level.

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GYRO TORQUING COMPUTATIONS – To overcome the problems that arise from platform tilt, the system uses the gyroscopic principle of precession. By using this principle as the aircraft flies over the rotating Earth, it is possible to apply a continuous torque to the proper gyro axis. This reorients the gyros to maintain the stable element horizontal to the earth’s surface and pointed north. Figure 7-37, view B, and Figure 7- 38, frames 3 and 4, show platform operation with proper Earth-rate and aircraft-rate torquing corrections. An analog or a digital computer develops the signals necessary to properly torque the gyros. The corrections for Earth rate depend on the aircraft’s position on the earth’s surface. The analog corrections come from highly accurate potentiometers that produce trigonometric functions of aircraft position. Position integrator shafts drive the potentiometers. To maintain the stable element oriented to the north reference, torquing corrections rotate the platform. The rotation is about the vertical axis compensating for vehicle velocity. Schuler Pendulum A pendulum is any suspended mass that is free to rotate about at least one axis. However, its center of gravity is NOT on the axis of rotation. Therefore, any pivoted mass that is not perfectly balanced is, by definition, a pendulum. The inertial platform is a pendulous device and, therefore, behaves as all pendulums behave. It aligns to the dynamic vertical when at rest. The pivot axis and the center of gravity align with the gravity vector. The center of gravity will be on the bottom. Also, it tends to break into its natural period of oscillation whenever the aircraft accelerates. Pendulous oscillation is periodic angular motion with the gravity vector as its midpoint. Periodic motion around the local vertical produces obvious errors from an inertial platform. This happens because misalignment about the horizontal plane introduces gravity components on accelerometer inputs. The system will interpret gravity accelerations as horizontal acceleration of the aircraft. The Schuler pendulum is a specially constructed pendulum without the unwanted oscillatory motions of non-Schuler pendulums. It is a special case of both the simple and the compound pendulums, which are discussed in the following paragraphs. SIMPLE PENDULUM – The simple pendulum consists of a small body suspended by a weightless string. The motion of the simple pendulum is both periodic and oscillatory. The period of the simple pendulum is given by the mathematical formula √

where, T = time of one oscillation in seconds, L = length of the string, and g = local gravity. The formula shows that the period of a simple pendulum is proportional to the square root of the length of the suspending string; the longer the string, the longer the period. One property of the simple pendulum that is very useful in the construction of an inertial stable element is shown in Figure 7-39. Two pendulums are suspended by strings of different lengths. Equal forces horizontally accelerate the suspension point of each 7-41

pendulum. The inertia of the bob resists the change in its state of motion. This action causes the bob to lag the point of suspension. It also produces an angular motion of the pendulum about the local gravity vector. Figure 7-39 shows that the length of pendulum (B) is longer than pendulum (A). It also shows angular motion of pendulum (B) is less than pendulum (A) for a corresponding linear motion of the suspension point. Therefore, the longer the suspending string, the less the angular motion of the pendulum for a given linear motion of the suspension point.

Consider what would happen in the following case. The suspending string is long enough to maintain the bob at the center of the earth. The suspension point is transported horizontally along the earth’s surface (Figure 7-39, view B). The bob is hypothetically at the center of the earth, the seat of the earth’s gravity field. Accelerating the suspension point along the earth’s surface merely realigns the suspending string with the new local gravity vector. Therefore, the angular motion of the pendulum about the gravity vector for any horizontal acceleration of the suspension point is zero. This particular pendulum is the Schuler pendulum, shown in Figure 7-39, view B. This pendulum gets its name from the German engineer, Maximilian Schuler. Schuler solved the problem of oscillating shipboard gyrocompasses in the early 1900s. Of course, Schuler couldn’t use the simple pendulum itself to solve this oscillating Figure 7-39 — Simple pendulum. 7-42

Figure 7-40 — Compound pendulum. problem. He used the principle of the simple pendulum to construct a pendulum that reacted like a simple pendulum. The length of this pendulum equals the radius of the earth, which is about 3,440 nautical miles long. The period of oscillation for this pendulum is about 84.4 minutes. Remember, the period of oscillation of a pendulum is proportional to the square root of its length. Therefore, any pendulum constructed to oscillate with a period of 84.4 minutes would have an equivalent length of about 3,440 nautical miles. Such a pendulum is the Schuler pendulum, a special case of the compound or physical pendulum. Figure 7-40 shows three examples of compound pendulums. COMPOUND PENDULUM – In Figure 7-40, view A, the pivot point, P, is farthest away from the center of gravity, represented by distance d. In view B, the pivot point is closer to the center of gravity than in view A. However, it is farther away than the one shown in view C, which pivots at the center of gravity.

The pivot point of each pendulum in Figure 7-40 is given the same acceleration. Therefore, each pendulum has the same linear motion at its pivot point. Yet, each pendulum has a different angular motion. As distance d decreases, the angular motion of the pendulum about the local vertical (gravity vector) decreases and distance L increases. Distance L is the distance from pivot point P to the center of oscillation, point O. Also, the pivot point and the center of gravity come closer together, and equivalent length L of the pendulum becomes longer. Figure 7-40, view C, shows the pendulum pivoted at the center of gravity. In this case there is no angular motion of the pendulum and the equivalent length L is infinite. Therefore, it is not a pendulum; it is a perfectly balanced mass that has an infinite period of oscillation. Thus, it is possible to construct a pendulum of infinite equivalent length and period. It is also possible to construct one that has an equivalent length of 3,440 nautical miles. Such a pendulum would be pivoted at some distance d from the center of gravity. This distance would be greater than the one in Figure 7-40, view C, but less than the one in Figure 7-40, view B. When 7-43

Figure 7-41 — Frame of reference. pivoted at a point where the period of oscillation is found to be 84.4 minutes, it becomes a Schuler pendulum. The stable element is essentially a Schuler pendulum. However, it is not entirely mechanical because the earth’s radius varies with latitude. The earth’s radius is greater at the equator than it is at the poles. For this reason the stable element uses the process of Schuler tuning. Schuler tuning torques the platform to a position normal to the gravity vector by signals received from a computing loop. Frame of Reference The frame of reference about which the INS defines the instantaneous position of the aircraft is the conventional latitude-longitude coordinate system (Figure 7-41). The local vertical, established and maintained by the inertial navigation system, is the gravity vertical and is coincident with the geographic vertical. The inertial navigation system orients to the true north reference by sensing the motion of the earth rotating on the polar axis. The frame of reference defined is horizontally aligned in a plane parallel to the surface of the earth and oriented to true north.

ESTABLISHING THE REFERENCE – Refer to Figure 7-34 and Figure 7-41. By establishing the frame of reference, three perpendicular axes of the stable element will align to the horizontal coordinates of the latitude-longitude navigational system. That is, the stable element z-axis aligns with the local vertical and the y-axis aligns north-south. Therefore, the z-axis is coincident with lines of longitude, and the x-axis aligns east- west coincident with lines of latitude. In all calculations, the x-axis is positive east and the y-axis is positive north. The z-axis is positive away from the center of the earth. A pair of two-degree-of-freedom gyroscopes establishes and maintains the stable element to the frame of reference. Since a two-degree-of-freedom gyroscope has two sensitive axes, it is necessary to use two such gyroscopes (Figure 7-34). The upper gyroscope z-axis is not in use. They physically mount on the stable element so their spin axes are exactly perpendicular in the horizontal plane. With this arrangement, alignment of the upper gyroscope spin axis north-south will automatically align the lower gyroscope spin axis east-west. The stable element containing the gyroscopes is supported by the platform gimbal system. Thus, the gyroscopes control the stable element. However, if a free gyroscope 7-44

initially orients so the spin axis aligns east-west in a horizontal plane, the gyro will precess. The precession will be about the earth’s surface because of the earth’s rotation on its polar axis. To maintain an Earth reference, the system must torque the gyro opposite and equal to the apparent precession. The earth’s rotation affects the upper (XZ2) and lower (YZ1) gyros. Corrections for Earth rotation go to the Y and Z1 torquing coils (Z1 and Z2 are caged together and both respond accordingly). The system does not use the X torquing coil for Earth rate corrections. To establish an Earth frame of reference, the gyroscopes are controlled by continuously computed signals. These signals introduce forces (torque) to cause the gyro spin axes to precess in the desired direction. This torque is in the form of direct current signals applied to torquing coils mounted on the gyro float assembly. It creates a magnetic field that aids or opposes the magnetic fields of the permanent magnets mounted on the end bells. This circuit effectively torques the gyro, causing the spin axis to precess to the desired orientation. The first step in establishing a frame of reference is leveling the stable element. To level the stable element, you align it to the local vertical (gravity vector). This is done by torquing the XZ2 and YZ1 gyros. This moves the stable element until the x and y accelerometers cease to sense any acceleration caused by gravity. This says that the outputs from the accelerometers provide the torquing signals for the gyros. During this time, and while operating, computed Earth rate torquing signals continuously go to the y and z axes torquing coils of the lower gyro. The size of these Earth rate torquing signals is resolved by computing the vertical and horizontal components of Earth rate as a function of latitude. After establishing the stable element in a rough level position, the x-axis torquing signal drives the stable element in azimuth to null this signal (fine alignment). This signal consists of Earth rate acceleration only, which is a measure of stable element unlevelness. At this time, the X gyro’s spin axis is aligned to true north establishing the frame of reference. This alignment condition will remain until you manually sequence the inertial navigation system to the navigate position. As previously shown, the earth’s rotation does not affect the upper X-axis gyro. Therefore, no compensating Earth rate torquing signal goes to this gyro. MAINTAINING THE HORIZONTAL REFERENCE – When the aircraft remains stationary or moves at a constant velocity, the accelerometer outputs are zero. If the aircraft attitude changes while maintaining a constant speed, the accelerometers on an unstabilized platform sense an acceleration due to gravity. Since the accelerometers cannot distinguish gravitational accelerations from horizontal accelerations, the integrators develop a fictitious velocity with a corresponding distance error. It is essential, therefore, that the system maintains the accelerometers in a truly horizontal reference plane. This plane must be independent of aircraft attitude at all times. This is a basic requirement of the inertial navigation system. The accuracy with which the horizontal reference is maintained determines the overall performance capabilities of the system. A gyrostabilized platform in a gimbal structure serves as an inertial reference. Also, it accurately defines directional reference for the coordinate system. With the platform installed in the aircraft and aligned along the y-axis, it will remain level regardless of aircraft attitude. The instant the aircraft begins to change pitch attitude (Figure 7-42), the platform gyro senses this angular movement. The gyros then begin to precess at a rate proportional to the pitching rate. A pickoff coil on the gyro axis senses this movement and changes it to a voltage. After amplification, the voltage goes to the pitch gimbal servo drive motor. The motor rotates the stable element exactly equal and 7-45

Figure 7-42 — Gyro maintaining inertial platform level. opposite to the aircraft angular motion. As a result, it continuously precesses the gyro to its neutral or level position. This action maintains the gyro output signal at null. Regardless of any new pitch attitude the aircraft assumes, the gyro will keep the stable element level. This is the only position that allows the gyro’s output signal to be at null. In actual practice, the stable element is maintained level. It uses a similar method to maintain its azimuth alignment in yaw and roll. This is necessary so the sensitive axis of the north-south accelerometer aligns true north-south. It is also necessary for the east-west accelerometer to align true east- west. The stable element is then accurately aligned to the three coordinates—north (y-axis), east (x-axis), and up or true vertical (z-axis). This arrangement allows the accelerometers to accurately detect aircraft motion. MAINTAINING THE VERTICAL REFERENCE – The stable element must remain perfectly level or the accelerometers will sense a false acceleration due to the earth’s gravity. The gyros try to maintain their inertial position in space and not with respect to the local vertical. This causes the stable element to drift off level as the aircraft moves over the curvature of the earth (Figure 7-42). If this situation were allowed to build up, very large errors in velocity and distance would occur. This condition develops whether or not the aircraft is moving over the earth’s surface. The earth’s rotation alone will develop the same type of errors. The stable element must always be perpendicular to the local vertical. Therefore, the system must make the gyros precess in such a manner as to maintain the stable element level. With the stable element level, the sensitive axes of the accelerometers are maintained horizontal to the earth at all times. Now they respond only to the horizontal component of acceleration. MAINTAINING THE FRAME OF REFERENCE – Accuracy in maintaining the stable element to the frame of reference determines the overall performance capabilities of the system. Gyro torquing rate signals are continuously computed to maintain the frame of reference. After alignment, the inertial navigation system is manually sequenced to its operating condition (navigate). If the aircraft were to remain stationary, the gyro torquing rate signals would consist of Earth rate only. However, as the aircraft moves over the curvature of the earth, the stable element Earth reference would be lost. This happens because vehicle movement causes the gyros to precess. Refer to Figures 7-41 and 7- 42. Therefore, the system continuously computes additional gyro torquing signals to compensate for the vehicle’s movement. These signals are aircraft rate torquing signals. They depend on the velocity of the aircraft with respect to the frame of reference; that is, 7-46

east-west velocity and north-south velocity. The angular rate is directly proportional to the velocity of the aircraft along the circumference of the earth. The system applies aircraft rate torquing to both gyros so they will precess about all three axes (x, y, and z). Thus, the system continuously maintains the frame of reference. Deriving Velocity and Distance The inertial navigation system can accurately detect aircraft acceleration, and using precision integrators, determine aircraft velocity and measure distance traveled. Accelerations are measured in units of ft/sec² by the accelerometers. However, for analog computation, the accelerations are developed in volts per g of accelerating force. The velocity integrator integrates the accelerating force to obtain velocity. The distance integrator integrates velocity to obtain the distance traveled. There are two velocity integrators in the system to obtain and along the two horizontal axes. There are two distance integrators to obtain distance traveled along the two horizontal axes. In addition, some inertial navigation systems employ one other velocity integrator to obtain along the vertical axis. Accelerometer Output Corrections The arrangement of the accelerometers is perfectly suited to navigation over a stationary plane or over flat terrain moving at uniform speed in a straight line. The earth, however, is a rotating sphere. As far as the inertial platform is concerned, only points along the equator can be considered to possess uniform linear motion. Here, and only here, the accelerometer signals can translate directly into position information. For this reason, it is necessary to provide an automatic device to alter the accelerometer signals. The automatic device allows the system to report meaningful information. The corrective device is purely electrical. All or part of the circuitry is active whenever a velocity signal voltage is present anywhere north or south of the equator. These circuits use the velocity signals, modified according to the latitude of the aircraft position, They insert artificial acceleration signals to those already in the accelerometer output circuits. The circuits are divided logically— some are for centripetal effect, some for Coriolis. Note, however, that the corrections are complementary. Although it is convenient to assign a separate purpose to the circuits, as in the following discussions, the functions overlap, and it is not accurate for you to consider them separately. CENTRIPETAL CORRECTION – Centripetal errors are false accelerations sensed when the platform is torqued to maintain its plane of reference. Centripetal correction differs from that of the Coriolis correction in that it has no relationship to Earth dynamics. Even if the earth were stationary, it would still be necessary to insert centripetal correction voltages to the accelerometer signal. Correction voltages ensure an accurate plot of any course that does not coincide with one of the earth’s coordinate great circle routes; that is, an exact polar or an exact equatorial orbit. On a great circle route, a route that circles the earth’s center, every linear acceleration initiates motion. However, unless the route is due north-south or directly along the equator, the system can’t plot the route accurately from raw accelerometer signals. The orbit resulting from linear acceleration, and the logic of applying centripetal corrections to get an accurate plot of track can be understood if you consider any simple circumstance involving a single acceleration and the inertial reaction. For example, assume a perfect bowling lane surrounds the earth at the equator. Theoretically, a bowler could stand behind the pins and roll the ball in the lane in the opposite direction 7-47

Figure 7-43 — Centripetal correction along latitude. (away from the pins). After a few years’ time, the bowler gets a perfect strike on the head pin. However, if the lane is on latitude 10°N, the ball would invariably veer south. After a few thousand yards, the ball would fall into the right-hand gutter if rolled east. It would fall into the left-hand gutter if rolled west. In both cases, the ball would roll into the south gutter. (The reason for this phenomenon is clear if you consider a lane built on a latitude in the Arctic only a few yards from the pole. Here, the curve of the lane is obvious.) It is apparent that, if the ball accelerates due east, it will follow, in inertial space, a straight course intersecting the outside gutter. For the ball to roll at a uniform speed and remain in the alley, a uniform north acceleration force must be exerted on the ball in transit. In this case, the north acceleration force is exerted on the ball during transit. Also, remember that north acceleration is a corrective force and does not produce north velocity with respect to the alley. As shown in Figure 7-43, a north-south accelerometer aboard an aircraft circling the pole finds the same phenomenon. In a steep-banked turn, the centrifugal force deflects the north-south accelerometer. It blindly reports a steady north acceleration. The system will show a growing north velocity when, in fact, only an east velocity exists.

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Figure 7-44 — Platform on great circle route. The fault is not in the accelerometers but in the geographic coordinate system’s nonlinear pattern to which the platform is slaved. The coincidence of Earth axis and pole implies Earth dynamics in the phenomenon, but actually this is not a factor. Assume we shift the coordinate systems to place the pole at New York. Now use this city as the focal point of one accelerometer axis of an inertial navigator circling the city. The accelerometer sensitive to this axis will report acceleration toward the city. In a spherical coordinate system, any linear vehicle acceleration initially affecting both accelerometers will result in the vehicle NEVER reaching the pole. The vehicle follows a great circle track that first approaches one of the poles. It will fly due east or west for a brief period, and then depart the pole, as shown in Figure 7-44.

Although the velocity resulting from any given acceleration is initially computed correctly with respect to space, the direction of the speed must be constantly altered. This is necessary if the navigational system is to accurately report a great circle course that crosses both latitude and longitude. On such a course, the aircraft does not turn as it does following a line of latitude. Therefore, it does not generate uniformly false north acceleration signals. The centripetal correction circuit does, however, continue to plant signals of acceleration toward the equator. This has the effect of altering the reported speed (the result of velocity along both axes). No actual acceleration has taken place to 7-49

cause a speed change. Thus, the centripetal correction circuit must simultaneously plant positive accelerations in one axis if it plants negative accelerations in the other. Therefore, when canceling some of the north or south velocity, the system must add a sufficient increment of east or west acceleration. This allows the reported total speed to remain unchanged, while the reported direction of flight bends toward the equator. As shown in Figure 7-44, after a single north-east acceleration has occurred, the north vector becomes progressively shorter. Note that the resultant speed vector is always of the same size. In summary, you can see that an east or west acceleration in either hemisphere contains a hidden element of acceleration toward the equator. Also, you can say centripetal correction simply acts to reveal this element. When aircraft velocity is not due north or due south, in the Northern Hemisphere the centripetal correction manufactures a south velocity component. In the Southern Hemisphere, it manufactures a north velocity component. CORIOLIS CORRECTION – As mentioned before, the scope of the centripetal correction makes no allowance for Earth dynamics. Since the earth rotates toward the east, all points on the surface have a constant tangential velocity. This velocity is maximum along the equator and lessens at higher latitudes. Tangential velocity is a linear quantity. It refers to the speed and direction an object would travel in a straight line if freed from the earth’s gravity. An object near the equator travels about 1,000 miles per hour in a circular path. Assume the object is free from Earth’s gravity and atmosphere. Now it will travel at that speed in a straight line away from the earth (on a tangent to the earth). Its tangential velocity is 1,000 miles per hour. If the object moves toward the North Pole, its speed in circular travel will decrease as it approaches the pole. The speed will be zero when placed exactly over the pole. Although the earth has a trajectory in space, this motion is not important to the inertial navigation system because every point on the sphere shares this trajectory. The only variable involved is the variation in the earth’s tangential velocity at different latitudes. While the accelerometers don’t automatically make allowance for this variation, the system must take it into account. The need for such an allowance is illustrated in the following discussion. Put an aligned INS, devoid of any Coriolis corrective mechanism, aboard a train in the Northern Hemisphere, and transport it north. The earth’s tangential velocity at the latitude where the INS aligned is 800 knots east. It is obvious the train’s eastward velocity must be constantly reduced as it progresses north. This progressive reduction in velocity represents an acceleration. Since the tracks constrain the train, a force from the east will be exerted on the wheel flanges. This force is sensed by the east-west accelerometer as acceleration to the west. The system begins computing west velocity that will grow. Refer to Figure 7-45. An aircraft flying north directly along a moving longitude meridian subjects the east-west accelerometer to this same force. This happens as its course in space alters to the left, compensating for the decreasing eastward velocity of the earth’s surface.

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Figure 7-45 — Coriolis effects.

The train traveling north stops at a latitude where the earth’s tangential velocity is 600 knots. This is 200 knots less than the tangential velocity at the point of its initial alignment. The velocity computer will continue to report that the vehicle is traveling west at 200 knots even though it is perfectly static. The Coriolis correction prevents this from happening. It creates enough east acceleration signal to offset the continual west acceleration generated by the east-west accelerometer as it travels north. The east and west accelerations cancel, thus, there is no change in longitude to report. Of course, on the return trip, the platform is under the delusion that 600 knots east is zero velocity. So as it travels south, the increasing Earth tangential velocity causes it to constantly report east acceleration. In this case, the Coriolis correction also reverses. It now manufactures a west acceleration that exactly voids the east acceleration. In both of the above cases, the Coriolis correction supplies an artificial signal of acceleration to the right side of the actual track. This is the nature of the Coriolis correction, regardless of the direction of travel in the Northern Hemisphere. Look at Figure 7-46. When the track is east along a latitude line, centripetal correction offsets the increment of north acceleration generated by aircraft speed. The Coriolis correction accounts for the additional force generated by the earth’s rotation. For example, if the earth’s tangential velocity at latitude L is 700 knots and aircraft speed 350 knots, total speed is 1,050 knots. Centripetal correction offsets the acceleration caused by the 7-51

Figure 7-46 — Coriolis and centripetal corrections. centrifugal force of a 350-knot turn of this radius. Coriolis correction offsets the force of a 700-knot turn. In this case, the two corrections are summed.

If the course is reversed and the plane flies west, the correct ions become opposite in polarity. The centripetal signal is still south, but Coriolis is north (to the right of the track). Therefore, only a part of the larger correction is effective. Schuler-Tuned Loop The Schuler-tuned loop is a closed loop circuit between the accelerometer, velocity integrator, and stable element. It prevents large velocity and distance errors caused by misalignment of the stable element. Figure 7-47 shows a simplified Schuler-tuned loop with the platform aligned. The output of the accelerometer is integrated to provide a velocity signal. The velocity signal is multiplied by 1/R where R equals the earth’s radius. This operation derives an angular velocity about the earth’s surface, V/R. The system uses this angular velocity to torque an integrating gyro. The torque causes the platform to precess about the earth‘s surface. This precession equals the rate the platform is being transported over the surface. This maintains the platform normal to the local vertical.

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Figure 7-48 — Simplified Schuler- tuned loop, platform unlevel. There are two such loops in an inertial navigation system, one for the north and the other for the east. The accelerometer in the north loop senses north-south accelerations, yet the gyro in the north loop senses east-west angular rates. That is, the vehicle’s angular movements about the east-west axis. By convention, we name accelerometers and gyros according to the direction of their sensitive or input axis. The inertial or Schuler loop takes the name of its accelerometer. The north loop contains the north accelerometer and the east gyro. The east loop contains the east accelerometer and the north gyro.

With the platform initially unlevel, as shown in Figure 7-48, the accelerometer senses a component of gravity, g sin θ. This signal is integrated, resulting in the velocity signal Vθ. The velocity signal then causes the gyro to precess in a clockwise direction. When the accelerometer is positioned to sense zero gravity, the velocity output continues to torque the platform in a clockwise direction. This causes the accelerometer to now sense a gravity component of the opposite polarity. This signal causes the velocity signal to decrease to zero. The velocity signal now builds up in the opposite direction and precesses the platform in a counterclockwise direction. The oscillation set up by this mechanization has a period of 84 minutes, equal to that of the Schuler pendulum.

Figure 7-47 — Simplified Schuler- tuned loop, platform level. 7-53

Figure 7-49 — Schuler tuning – acceleration errors versus velocity errors.

Figure 7-49 shows the buildup and decay of acceleration and velocity errors as a result of such errors as just described.

ALIGNMENT Inertial navigation depends on the integration of acceleration to obtain velocity and position. In any integration process, the system must first know the initial conditions. In this case the initial conditions are velocity and position. The accuracy in solving the navigation problem depends greatly upon the accuracy of the initial conditions. Therefore, system alignment is of paramount importance. System alignment consists of creating a coincidence between the platform axes and the c omputer axes. This can be done by rotating either or both systems. There are two general methods of accomplishing this condition. 1. The system is slaved to an external reference source. 2. The system may have the built-in capability to sense misalignment and correct itself. External references take three basic forms—terrestrial, celestial, and inertial. The terrestrial system uses surveyed lines, bench marks, plumb bobs, and bubble levels. These methods result in level accuracies of about 10 seconds of arc and heading accuracies to 3 minutes of arc. Celestial information from star trackers and radio sextants has accuracies to 10 seconds of arc. When using an inertial system as a source, the accuracies depend on its initial source and length of time since last aligned. Such a method is usually for mobile alignment where primary sources cannot be used. 7-54

Figure 7-50 — Typical leveling loop. The use of an external reference system requires transfer devices to transmit the reference information to the system. The transfer devices are either optical or electromechanical devices. Optical devices include theodolites and autocollimators; the electromechanical devices are synchro-resolver or digital type. Optical methods can produce accuracies of a few seconds of arc. The electromechanical methods are accurate to about 30 seconds of arc. In self-alignment, the inertial sensing instruments mounted on the platform sense the deviation from the desired position. To determine the orientation of a three-axis orthogonal (perpendicular) coordinate system, you must have at least two noncollinear (not parallel) reference vectors. For self-alignment, the earth’s spin vector and the mass attraction gravity vector serve this purpose. The self-alignment puts fewer requirements on the computer. In self-alignment, the accelerometer outputs don’t have to be resolved into components of gravity and vehicle acceleration. When accelerometers and gyros mount on the same element, their relative position is fixed and does not have to be computed. Self-alignment often divides into two modes— rough or course alignment (sometimes called caging) and fine alignment. Fine alignment itself divides into two modes— leveling and gyrocompassing. Rough Alignment Rough alignment provides a convenient starting point for the later phases of alignment. In most cases, the gimbals slave to their own synchro outputs or to some external source. This external source will have a particular orientation with respect to the vehicle. A timing network controls the duration of rough alignment, which is usually a short period. Fine Alignment (Leveling) To accomplish fine alignment or leveling, the system rotates the platform axes to the computer axes. For a locally level system, this is done by placing the x and y accelerometer axes mutually perpendicular to the gravity vector. Since accelerometers mount at right angles, a motion about one axis causes the other to go through an angle with respect to the gravity vector. Therefore, you connect the accelerometer output in a way to allow it to torque about its perpendicular axis. The accelerometer can now slew itself to a null position, where it senses no component of gravity. As pointed out earlier, the device the system can torque about an axis is a gyro. In this case, the gyro being torqued is the one sensitive about the axis perpendicular to the accelerometer being leveled; that is, the y gyro (north) torques the x accelerometer (east) to level, etc. The accelerometer will provide a dc voltage that is proportional to the sine of the off-level angle. In the leveling mode, after amplification this voltage goes to the gyro, whose sensitive axis is perpendicular to the sensitive axis of the accelerometer. See Figure 7-50. In other words, the output of the x accelerometer goes to the y gyro. The output of the y accelerometer goes to the x gyro. 7-55

Figure 7-51 — Typical heading alignment loop. Gyrocompassing As mentioned, any gyro not having its spin axis parallel to the earth’s spin axis will apparently precess about the earth. The rate at which it precesses is proportional to the angle between its spin axis and the earth’s spin axis. This angle can be resolved into two components. Since the gyro spin axis lies in the level plane already established, one component can be found in the plane itself. Now, find the angle between the spin axis and a vector in the level plane that intersects the earth’s spin vector. Once you find these two angles, you know the exact position of the platform axis. You can see the rate at which the gyro in question appears to precess is equal to Ω cos λ sin a where Ω = the earth’s rate of rotation λ = latitude α = the angle the platform makes with true north

The spin axis of this gyro can be torqued to a place where this term goes to zero (α = 0). Also the computer can develop a torquing term equal to this and apply it to the gyro. In either case, the position of the platform is known. Figure 7-51 shows a typical heading alignment loop. In a north-seeking platform, the system uses the earth’s rotation to align the platform to true north. It accomplishes this by using the output of the y accelerometer. The system applies this output to the torquing coils of the z (azimuth) and the x (east) gyros. At the beginning of the gyrocompass phase (after the platform is leveled), the stable element is torqued in azimuth. This nulls out the residual east gyro torquing rate. If the stable element is not aligned to true north now, it begins to tilt due to precession of the gyros. The y accelerometer senses deviation of the stable element from level because of gravity. The output of the accelerometer then torques the x gyro until the stable element is level. At the same time, the output signal also torques the z gyro in azimuth. The process continues until the stable element aligns to true north. Once the platform is aligned, the operator switches the system from the alignment phase to the navigation phase of operation. In the navigation phase, the stable element would maintain an orientation about free space if not for corrections supplied by the computer. The computer maintains the stable element level with respect to the earth and oriented to true north. If not, the accelerometers sense gravity in addition to movement of the aircraft. Coriolis, centripetal, and Earth rate corrections are computed and used to hold the stable element level and aligned to true north. In the navigation phase of operation, the orientation to true north is dependent on the original aligned position and the computed corrections. 7-56

Figure 7-52 — Wander angle. To this point, a north-pointing inertial system has been discussed. A disadvantage of the north-pointing system is that it cannot operate in the Polar Regions; it must always be physically pointed north. If the system flies directly over the pole, it must rotate 180° to a gain be pointing north. This rotation would not be physically possible because of the extremely high torquing rates necessary. Most north-pointing inertial systems cannot operate within several hundred miles of the poles due to stress on the system components. Wander Azimuth The wander-azimuth inertial system solves the problems of operating an inertial system at the poles. The fundamentals of a wander-azimuth system are the same as a north- pointing system. During the gyrocompassing mode, the system allows the platform to take an arbitrary angle (wander angle) with respect to true north. As previously mentioned, the platform is leveled; but, the accelerometer outputs are now supplying torquing signals to both gyros. This action compensates for the earth’s rotation (this signal was sent to x gyro only in the north-pointing system). Eventually, the correct Earth rate torquing signals maintain the platform level. The computer then uses the ratio of Earth rate compensation to compute the wander angle (Figure 7-52). As the wander-azimuth system navigates around the earth, the wander angle (with respect to true north) changes as a function of longitude. The system operation is the same as a northpointing system. However, the wander angle is taken into account by the computer with the north and east sensored accelerations. Alignment at Sea Problems that arise in aligning an aircraft INS on aircraft carriers at sea are more complex than aligning the INS ashore. This situation exists even though our carrier- based inertial reference is another INS of very high accuracy. The inertial navigation reference system aboard an aircraft carrier is the Ship’s Inertial Navigation System (SINS) and the Relative Velocity Computer (RVC). Outlets on the flight deck make it convenient to pipe the SINS reference information into the aircraft inertial navigation system. However, one major problem still remains that makes proper alignment difficult: the accelerations experienced by the ship’s inertial navigation accelerometers located below deck. These accelerometers are remote from the aircraft accelerometers, and therefore do not sense the same accelerations. 7-57

COARSE SEA ALIGNMENT – During coarse sea alignment, the best available true heading is from the SINS and the RVC. Aircraft carrier true heading goes to the RVC. Here a manually-selected aircraft heading angle, with respect to the aircraft carrier, is inserted. The combined signals then go to the heading computer in the aircraft’s inertial navigation system, thus concluding the coarse sea alignment. FINE SEA ALIGNMENT – During fine sea alignment, the accelerometers sense the aircraft carrier movement in addition to gravity. Only the gravity component is used in the leveling. This allows accelerometer output caused by aircraft carrier movement to be canceled. The SINS and RVC accomplish this task by supplying continuously computed corrections. The accelerometer error signals are integrated to supply a velocity. The reference velocity supplied by the RVC during sea alignment, actual aircraft velocity, is subtracted from the accelerometer derived velocity. The difference corresponds to the gravity component sensed by the accelerometer. After amplification it is applied to the torquing coils of the gyros. The processing of the gyro pickoff signals cause the gimbals to rotate and cancel the pickoff error signals. The stable element is torqued until the accelerometers show a null or level condition. TYPES OF INERTIAL NAVIGATION SYSTEMS You can classify inertial navigation systems under two broad types— pure and hybrid. The types of pure INS are analytic, semianalytic, geometric, and strapdown. The types of hybrid INS are radio inertial, Doppler inertial, and stellar inertial. Pure Inertial Navigation Systems As the name implies, a pure INS is not combined with other equipment to improve its operating performance. Analytic Inertial Navigation System The analytic INS uses a platform with a fixed angular reference to some point in inertial space. The system makes no attempt to force the accelerometer input axes into a preferred alignment with respect to the earth. This method does not require gyro torquing. As a result, this platform is subjected to errors of gyro drift only. Because the platform remains rigid in space and rotates about the earth, the output accelerations become complex. They essentially consist of two major accelerations— the actual acceleration of the vehicle and the gravitational acceleration of the earth. For navigation purposes only aircraft accelerations are required and wanted; therefore, the gravitational accelerations must be canceled. Yet this cancellation is difficult to obtain because the earth’s gravitational acceleration is not uniform. Therefore, the computer must store an enormous amount of data to effect this cancellation. The significant disadvantage of analytic inertial navigation is the result of maintaining the accelerometer referenced to a fixed point in inertial space. As the stable element navigates about the earth, the accelerometers must sense aircraft acceleration and the earth’s gravitational field component. The accelerometers for this system must have a wide dynamic range as well as a high overall accuracy. The most serious problem, however, is the cancellation of the gravitational accelerations. Irregularities in the earth’s shape and mass cause variations in the gravitational field. Therefore, the cancellation of these variations requires a complex computer with a very large storage capacity. 7-58

Figure 7-53 — Typical semianalytic inertial navigation system, block diagram. Semianalytic Inertial Navigation System The semianalytic system is the INS most commonly in use today. All naval aircraft that use INS have this type of system. It may either be a pure system or work with another navigation system as a hybrid system. This system’s chief advantages are the simple platform gimbal structure and computer functions that are easily attained by either analog or digital means. The semianalytic system always maintains the stable element normal to the earth’s gravitational vector just as in other systems already discussed in this chapter. In this system, the computer converts the output accelerations of the stable element to angular velocities. These angular velocity signals then torque the platform gyros to maintain the platform normal to the earth’s gravity vector. The computer also develops signals to prevent the platform from processing off level due to the earth’s rotation about its polar axis. These signals are equal to the angular velocity of the earth resolved into the system axes. The system then applies these signals to the gyro torquers. A typical simplified block diagram of a semianalytic INS is shown in Figure 7-53.

In a semianalytic inertial system, the platform aligns normal to the gravity vector. It may or may not align to true north. The output of the north accelerometer goes to an integrator. Here, the output is summed with acceleration correction terms to derive a true vehicular acceleration over the earth’s surface. This acceleration signal is then integrated with respect to time, deriving the north velocity component of the vehicle’s track. Through scaling, the INS converts the velocity term to an angular velocity. It then integrates the angular velocity to provide a position readout in the form of latitude. In addition, the latitude function generator uses the north angular velocity signal to develop accelerometer correction terms. It also uses this signal to develop a gyro torquing signal for the east gyro. 7-59

Figure 7-54 — Transport of the geometric system’s stable element. The east accelerometer output is summed with accelerometer correction terms and integrated to provide an east component of the vehicle’s track. Through scaling, the INS converts the velocity signal to an angular velocity. This angular velocity goes to the latitude function generator. The function generator uses it to develop accelerometer correction terms and torquing signals for the north gyro and the azimuth gyro. In addition, the INS integrates the angular velocity to develop a position readout in the form of longitude. Geometric Inertial Navigation System The geometric INS uses a gyro system that, like the analytic system, is referenced to inertial space in a nonrotating plane. The accelerometers, however, mount on the gimbal structure in a manner as to remain normal to the earth’s gravitational field. Figure 7-54 shows the relationship of the accelerometers and gyros as the platform moves over the surface of the earth. When the platform is aligned at the equator and then moves north, the gyros maintain their position in inertial space. The accelerometers remain in a plane tangent to the earth’s surface at all times. The main advantage of this system is that the gyros are not torqued. Therefore, scaling of the gyros is not critical. The major disadvantage is economy. The system requires a high degree of accuracy to position the latitude and longitude gimbals. The semianalytic system requires much less precision to achieve similar accuracy; therefore, it costs less.

Strap-Down Inertial Navigation System In the strap-down system, the gyros and accelerometers mount directly to the frame of the vehicle. Its principal use is in ballistic missiles and spacecraft. This type of system can be mechanized for use in aircraft. However, the present state of technology makes it more feasible to use one of the other types of systems for aircraft use. T he strap-down system requires complex digital computers; analog computers are not accurate enough for use in this system. The computer in the strap-down system replaces the gimbal structure as the gimbal structure replaced the physical length of the Schuler pendulum. Figure 7-55 shows a simplified block diagram of a strap-down inertial navigation system. 7-60

Figure 7-55 — Strap-down inertial navigation system, block diagram.

In the strap-down system, the gyros provide angular rates, which the system converts to directional cosines (for example, space vectors). The strap-down uses these signals to determine vehicle attitude about an inertial frame of reference. The coordinate converter, using inputs from the accelerometers and the directional cosine converter, determines accelerations along the inertial reference axes. The position converter accepts inertial acceleration and altitude information to develop Cartesian coordinates representing the vehicle’s position in inertial space. These vectors then go to the vector solver, where they are summed to provide readouts of latitude and longitude. To accomplish strap-down system alignment, you must supply the directional cosines of the vehicle frame to the computer. The vehicle requires no physical orientation. Hybrid Inertial Navigation Systems The hybrid system is a combination of INS and some other type of navigation system. The other navigation system is for updating or improving the accuracy of the inertial navigation system. In other words, the hybrid inertial system combines two navigation systems so that the good characteristics of both are maintained. There are two types of updating processes used in hybrid systems. One type is the damping effect, which compares the inertial ground velocities with the ground velocities of some other system. The system uses the error, or difference between the two velocities, to damp out platform errors. The other type is the reset method. This method ignores the orientation of the platform and merely resets the position of the velocity shafts periodically. CALIBRATION As you learned earlier in this chapter, variation and deviation affect the accuracy of a magnetic compass. Variation is a natural phenomenon whose magnetic strength varies in intensity throughout the world. Variation is marked on navigation maps and is corrected for by the pilot. You can consider deviation as man-made. The magnetic fields of aircraft components cause deviation. These components are engines, electric equipment, landing gear struts, and flight control surfaces and their control cables. You can keep the effects of deviation to the very minimum by a process called compass swinging. 7-61

Figure 7-56 — Compass rose. As an AE, you must be completely familiar with the two methods of compass swinging. These two methods are the MC-2 compass calibration set and the use of a compass rose. For a gyrostabilized compass or inertial navigation system, the MC-2 is considered the primary means of swinging. Deviation in a wet or standby compass is corrected for on a compass rose. The following paragraphs describe the compass rose and the MC-2 calibration set. Compass Rose Aircraft magnetic compasses (wet or standby) have devices called compensators, which provide a means for correcting deviation errors. You cannot eliminate all errors, but you can reduce them to a minimum. Swinging the compass, you first compensate the N-S and E-W headings. Then set the aircraft on every 15- or 30-degree heading on the compass rose. Here, you note the difference between the aircraft heading and the indicated heading. You then adjust the compensators to reduce this difference or deviation to a minimum. There are two types of compensators. One type is the universal screw type. It consists of an assembly having a group of small compensating magnets permanently installed in it. To change the compensating effect of the assembly, you use two adjusting screws. One screw is for north-south compensation, the other for east-west. The other type of compensator has small, loose magnets that you place in special chambers on the compass as needed. The chamber positions allow one to make east-west corrections. The other (at right angles to the east-west chamber) corrects north-south deviation. Compensation is done only on the cardinal headings on standby compasses. However, on all other compass systems in naval aircraft, compensation is at 15-degree increments. Before starting the swinging operation, you should make sure all magnetic equipment is in the position it occupies in normal flight. Also, be sure that no one near the aircraft compasses during swinging operations has any magnetic materials on their person. Magnetic materials include tools, pocketknives, mechanical pencils, wristwatches, dog tags, bracelets, eyeglasses, jewelry, officer caps, badges, etc. Remember, too, that you use a nonmagnetic screwdriver in adjusting universal compensator screws. You actually swing a compass in one of several ways. However, as an AE, your chief interest is in ground swings. You usually accomplish a ground swing with the aircraft at rest on a compass rose. Look at Figure 7-56. Most air stations have a compass rose. The compass rose looks much 7-62

like an oversized card from a navigation compass. The directions shown by it are magnetic directions, and the north arrow points toward the earth’s north magnetic pole. A compass rose may also have a line showing true north. Jacks, lifts, hoists, or any dolly needed to perform the ground swinging job should preferably be of nonmagnetic material. However, this is not always possible. Devices used in the swinging process must be tested for their effects on the compasses. You do this by moving them about the aircraft in a circle with normal separation distance between the device and the instruments. Do not use devices that cause more than one- quarter degree change in the compass reading. Trucks, automobiles, railroad cars, and other aircraft containing magnetic metals should not be within the swinging area. These items could have a magnetic effect on the compasses of the aircraft being adjusted. You should also be sure that the compass is in good condition. Examine the compass for clear liquid and proper level. Check to see that the card assembly is level. Also check that it turns freely when the aircraft’s tail is in a level flying position. Set the compensator so it has no effect on the main compass magnets. Using a loose- magnet compensator, remove all loose magnets from their chambers. Set universal screw-type compensators for zero effect by turning both adjusting screws until the dots on the screws match with the dots on the compensator case. Then, place the aircraft on a south magnetic heading over the compass rose, with the tail in a level flying position. The aircraft engine(s) should be turning, and as many pieces of avionics equipment as possible turned on. This will create as many stray magnetic fields as possible and simulate the condition of the aircraft in flight. Note the compass reading and record it. From this reading, it is simply a matter of algebraic subtraction (or subtraction of numbers having plus and minus signs) to determine the deviation on the south heading. The deviation is the algebraic difference between the magnetic heading and the compass reading. Deviation is the error in a magnetic c ompass caused by electromagnetic disturbances in the aircraft. After doing this, place the aircraft on a west heading. Again, note the compass reading and determine the deviation or difference between the magnetic heading and what the compass reads. Next, turn the aircraft heading to magnetic north. Take the compass reading on this heading and determine the deviation. Now subtract, algebraically, the south heading deviation from the north heading deviation and divide the remainder by two. For example, if the compass reads 175 1/2° while on the south heading (180°0), record this as a deviation of +4 1/2°(180°– 175 1/2°). If the compass reading is too low, the deviation is plus; if the reading is too high, the deviation is minus. Suppose that on the north (000°), heading, the compass reads 006 1/2°. Such a reading is 6 1/2° too high. You would record this as a deviation of –6 1/2° (000° – 006 1/2°). The next job is to determine the coefficient of north-south deviation. You accomplish this by subtracting, algebraically, the deviation on the south heading from the deviation on the north heading. You then divide the remainder by two. (–6 1/2°) – (4 1/4°) = –11° = – 5 1/2° 2 2 The aircraft is still on the north heading and the compass reads 006 1/2°. Since the coefficient of the north-south deviation is – 5 1/2°, you must adjust the north-south 7-63

compensator by this amount. The compass reading on the north heading will now be 001°. This adjustment also corrects the south deviation by the same amount (but in the opposite sense). The south heading on the compass will now read 181°. The coefficient of north-south deviation, which is – 5 1/2° in this case, is called coefficient C. On the loose-magnet type compensator, you adjust north-south deviation by inserting the necessary number of magnets into the lateral (athwartship) chamber of the compensator. If the compass has a universal compensator, you make the adjustment by turning the north-south (N-S) compensator screw. The next step is to determine the east-west deviation. Turn the aircraft heading to magnetic east, according to the compass rose. Record the compass reading on that heading. Now determine the coefficient of east-west deviation, otherwise known as coefficient B. Assume, for example, that the compass reads 276° when the aircraft was on the west (270°, heading. Also assume it reads exactly 90° on the east (90°) heading. You find coefficient B by algebraically subtracting the deviation on west (– 6°) from the deviation on east (0°) and dividing by two. (0°) – (–6°) = + 6 = + 3 2 2 While the aircraft is on the east heading, adjust the east-west (E-W) compensator to add 3° to the compass reading. This reading becomes 93° on the east heading, and the compass would read 273° on the west heading. Make this adjustment by turning the E- W screw on a universal compensator. On the loose magnet type compensator, add the necessary magnets in the longitudinal (fore-and-aft) chamber. Leaving the aircraft on an east magnetic heading, next compute an overall deviation correction based on coefficient A. This coefficient is equal to the algebraic sum of the compass deviations on all four cardinal headings (north, east, south, and west) divided by four. (–6 1/2°) + (–0°) + (4 1/2°) + (–6) = (– 8) = – 2° 4 2 You must compensate instrument panel compasses for coefficient A if it amounts to 2° or more in either direction. When making this correction, leave the magnetic compensators alone. To compensate for coefficient A, move the instrument in its mounting. Compensate panel-mounted compasses for coefficient A by slightly realigning the whole instrument panel. You can also turn the compass a little with relation to the front of the panel and placing washers or spacers under its mounting screws. After completing this swing, swing the aircraft again on at least eight equally spaced headings (for example, 0°, 45°, 90°, 135°, 180°, 225°, 270°, and 315°). Record the compass readings for each heading on a compass correction card. Figure 7-57 shows an illustration of a compass correction card.

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COMPENSATING SWING RESIDUAL SWING ACTUAL HEAD (M) AIRCRAFT COMP. DEV'N ACTUAL HEAD (M) AIRCRAFT COMP. N 000 000 006 ½ –6 ½ 000 001 045 045 E 090 090 090 0 090 093 135 135 S 180 180 175 ½ +4 ½ 180 181 225 225 W 270 270 276 –6 270 273 315 315 (1) (2) (1) – (2) (3) (4) IF SWINGING COMPASS USED AHEAD OF AIRCRAFT ADD OR SUBTRACT 180 DEGREES COEFF C = N - S = (–6 1/2°) – (4 1/2°) = –11° = – 5 1/2° 2 2 2 COEFF B = E - W = (0°) – (6°) = + 6° = + 3° 2 2 2 COEFF A = N+E+S+W = (–6 1/2°) + (0°) + (4 ½ °) + (-6°) = (- 8°) = - 2° 4 4 4 Figure 7-57 — Compass correction card. BU# 166983 SER# 9548-563 SWUNG 2-28-12 AIRCRAFT COMPASS BY ____________ TO FLY STEER TO FLY STEER N 001 0180 181 015 016 195 196 030 030 ½ 210 210 045 045 225 225 060 061 240 240 ½ 075 078 255 256 ½ 090 093 270 273 105 107 285 286 ½ 120 121 300 300 ½ 135 135 315 315 150 150 330 330 165 165 ½ 345 346 7-65

Figure 7-58 — MC-2 magnetic compass calibrator set.

Detach the small right-hand portion of the compass correction card and mount with the compass. It is thus available for ready reference, telling the pilot or navigator the comparative compass headings and magnetic headings. Turn the larger portion of the card into maintenance control for insertion into the aircraft logbook. MC-2 Magnetic Compass Calibrator Set The MC-2 provides a controlled magnetic field (simulated Earth’s magnetic field) about the aircraft flux valve to accurately calibrate the compass system. Use of the MC-2 only requires the aircraft be accurately placed on a north-south line, thus eliminating the need for rotating the aircraft on a compass rose. The compass calibrator provides electrical heading inputs from 0° to 345° in 15-degree increments with an accuracy of 0.1°. The compass calibrator can survey an area for magnetic uniformity. It also provides the necessary data for layout and marking of a compass swing site. The compass calibrator (Figure 7-58, frame 1) consists of four major components—the control console, magnetic field monitor, remote transmitter turntable, and field tester. The set also includes various cable assemblies, reels, racks, tripods (Figure 7-58, frame 2), and some special alignment equipment.

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Figure 7-59 — Alignment equipment. The control console contains the controls, indicators, and electronic components that allow the compass calibrator set to operate. It uses 115 volts, 400 Hz at a maximum of 1 ampere. MC-2 changes this electrical power into ac/dc voltages that it requires. The magnetic field monitor is an engineer’s transit that has been modified to operate as a part of the compass calibrator set. The modification consists of installing a magnetic sensing element in place of the magnetic compass. The monitor is of nonferrous and nonmagnetic materials. It has a telescope, a horizontal circular scale with an adjustable vernier azimuth scale, levels, and leveling adjustment screws. The telescope is 22- power with an interior focusing optical system, and rotates 180 ° in a vertical plane. The remote transmitter turntable is also an engineer-type transit with the compass, vertical circle, and telescope removed. Also included with the turntable is a transmitter mounting bracket and a rain hood. The field tester is a portable metal-encased tester. It consists of a test panel, a shield can assembly, and a magnetic azimuth reference detector. All connectors, controls, switches, and electronic parts mount on the test panel. The shield can assembly contains a valve assembly within two magnetic shield cans. The magnetic azimuth reference detector consists of a 6-power telescope with azimuth adjustment and a flux valve assembly on a triangular support plate. The valve assembly has an attaching cable assembly. The alignment equipment (Figure 7-59) consists of a telescope, two plate assemblies, shaft coupling, quick connector, plumb bob and adapter, screwdrivers, magnifier, wrenches, and sunshade. The parts you use depend on the aircraft and transmitter under calibration. The telescope is a fixed-focus type, 8-power, with 360-degree azimuth rotation. A drum dial fine-adjusts azimuth, and an azimuth lock prevents unwanted rotation. The compass calibrator set is used to conduct an area magnetic survey. The survey determines the size and direction of the earth’s magnetic field at a proposed aircraft swing site. You will also use the compass calibrator to conduct the actual compass swing. The control console provides controlled dc currents for the transmitter. The monitor detects the size and direction of the earth’s magnetic field and supplies this information to the control console. You use the alignment equipment with the turntable to optically align the compass system transmitter (flux valve). A review of the flux valve will be helpful in the following discussion. In an electrical compass swing, a dc magnetic field is generated in the transmitter and varied in size and direction. This allows the MC-2, in combination with the horizontal 7-67

Figure 7-60 — Electrical swing and manual swing at a 90-degree heading. component of the earth’s field, to simulate an equivalent Earth’s field in the transmitter at a desired heading. Errors in the compass system are measured as the difference between the aircraft magnetic heading and the simulated Earth’s field magnetic heading. The aircraft heading shows on the aircraft compass indicator. The simulated Earth’s field heading shows by selecting the HEADING SELECTOR switch on the control console. Controlled dc currents to the secondary coils the transmitter generate an electromagnetic field (electrical swing). By applying a current to coil leg A of the transmitter (Figure 7-60), you generate a field that aligns to leg A. In the electrical swing, this field provides the north-south component of the simulated Earth’s field. A dc current also goes through coil legs B and C to generate two fields, each aligned to its respective coil. These fields are so oriented that north-south components of these two fields cancel, leaving one east-west component. By reversing the direction of the current flow, you can rotate the east-west component 180°.

The procedures for an electrical compass swing using the magnetic compass calibrator set are as listed below. 1. Set up the turntable over the spot where the remote compass transmitter will be when the aircraft is on the north line. 2. Remove the remote compass transmitter (flux valve) from the aircraft and mount it on the turntable. 3. Determine the alignment of the transmitter to magnetic north and its electrical calibration to the ambient magnetic field. Calibrate the N-S and E-W adjustments on the transmitter. 4. Mount the necessary optical alignment equipment to the remote compass transmitter. Align the telescope to a predetermined target one-half mile or more away. 7-68

5. Tow the aircraft into position exactly on the north line by using plumb bobs or some other accurate method. 6. Compute the optical alignment correction. Insert the correction into the optical alignment scope. Replace the compass transmitter in the aircraft (sighting on the same target used in step 4 above). 7. With the transmitter fastened down, reconnect the leads. 8. Using the appropriate adapter cables, connect the compass calibrator set into the compass system. The aircraft magnetic headings are set in with the heading selector on the control console. Record the errors as the difference between the indicated heading and that set in with the heading selector. Calibrate the compass system components to within 0.10°of the heading selector position. For more detailed information on compass swinging, refer to Military Standards, MIL- STD -765A. Consult this specification for additional information in connection with swinging, compensating, and calibrating compasses. GLOBAL POSITIONING SYSTEM FUNDAMENTALS Global Positioning System (GPS) is funded by and controlled by the U. S. Department of Defense (DOD). While there are many thousands of civilian users of GPS world-wide, the system was designed for and is operated by the U. S. military. GPS provides specially coded satellite signals that can be processed in a GPS receiver, enabling the receiver to compute position, velocity, and time. Four GPS satellite signals are used to compute positions in three dimensions (Figure 7-61, frames 1 and 2) and the time offset in the receiver clock. GPS is a one-way (listen only) system, in which the satellites transmit signals but are unaware who is using the signal (no receiving function). The user (or listener) does not transmit a signal, and therefore cannot be detected by the enemy (military concern), and cannot be charged for using the system (civilian concern). As GPS is a multi-satellite system, there are always a number of satellites visible simultaneously anywhere on the globe and at any time. The Satellite Constellation is made up of 24 operational satellites in 10,898 mile high, semi-synchronous orbits. A minimum of 5 satellites are observable from anywhere on Earth with four satellites required to produce the most accurate position solution. 7-69

Figure 7-61 — GPS basic components.

GPS can support a number of positioning and measurement modes in order to satisfy simultaneously a variety of users, from those requiring only navigation (decameter) accuracies to those demanding very high (millimeter - centimeter) accuracies for military use which are dependent on cryptographic codes. GPS Today The evolution of airborne navigation from compasses and inertial gyros of the 1950s to the GPS receivers of today has produced a dramatic increase in the speed and accuracy with which an aircraft’s position on the earth can be determined. GPS was rapidly adapted for aviation, as it can give a position (latitude, longitude and height) directly, without the need to measure angles and distances between intermediate points. Position can now be established almost anywhere as it is only necessary to have a clear view of the sky so the signal from the GPS satellites can be received clearly. Today’s GPS satellites transmit two carrier frequencies that are commonly referred to as L1 and L2, both of which contain codes that provide positioning, timing, and navigation information. Utilizing these frequencies and codes allows GPS receivers to track several satellite signals at the same time, so that precise positioning can be calculated anywhere on Earth. 7-70

As shown in Table 7-1, the L1 carrier contains Coarse/Acquisition (CA) code, which is commercially available. The L2 carrier contains only the P/Y, which is an encrypted code reserved for military use. Crypto keys enable GPS to receive highly accurate P- code (precise) navigation signals.

Table 7-1 — GPS Carriers Carrier Frequency Code L1 1575.42 MHZ C/A and P/Y L2 1227.6 MHZ L2C and P/Y

Each satellite transmits two RF signals. Each signal is modulated with a unique code sequence and navigation data message. The code sequence allows the GPS to identify each satellite. The navigation data message provides the GPS with ephemeris and almanac data. Almanac data represents current satellite positions while ephemeris data represents satellite clock and position errors calculated by dedicated ground stations. The GPS receives, tracks, and processes L1 and L2 frequency band RF GPS signals from the antenna and provides Position, Velocity, and Time (PVT) information to aircraft interfaces. The GPS is made up of hardware and software to do the GPS signal navigation tasks and to do Built-in Test (BIT) on the GPS receiver and the batteries installed. Signal processing involves reception and amplification of the satellite signal, sequential code and carrier tracking to measure pseudo-range, delta range, and data demodulation to verify correct reception. GPS aircraft position data is not susceptible to local atmospheric pressure variations or other environmental effects but it is affected by two types of atmospheric delays which can affect the accuracy of satellite signal measurements. Tropospheric delay can be predicted and included in the satellite almanac data. After collecting the almanac data, the GPS removes the predicted delay from the satellite signal measurements. Ionospheric delay occurs because the ionosphere is thicker in some areas than in others (Figure 7-62). This delay causes a greater phase shift of the L2 RF signal than in the L1 RF signal. The delay is measured by the difference in phase shift between the two signals.

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NOTE Nonvolatile memory can include: 1. Satellite tracking information 2. Last computed position and velocity 3. Selective availability and antispoofing data 4. Waypoint data 5. GPS Crypto key data

GPS in Naval Aircraft The typical GPS you will encounter on naval aircraft are comprised of the following components: GPS receiver – Receives modulated navigation signals from satellites to determine aircraft PVT and then provides PVT information to aircraft interfaces for use by other navigation and sensor systems. The receivers have batteries to maintain nonvolatile memory when aircraft power is removed.

The GPS receiver determines distance to a satellite by measuring the time difference between when the satellite transmits the signal and the time GPS receives the signal. The time the GPS receives the signal is determined by the GPS clock. When the GPS clock is not perfectly synchronized with the satellite clock, the time measurement is inaccurate. Figure 7-62 — GPS atmospheric delay. 7-72

The inaccurate time measurement used in the distance calculations prevents accurate GPS position to be found. The satellite clock error is measured by a dedicated ground station and the correct data is included in the data sent to the GPS receiver. The sole purpose of one of the four satellites is to provide additional measurement data needed for GPS to calculate clock error common to all distance measurements. Once the clock error is found, it is removed from the distance measurements. GPS antenna – Usually flush mounted on the aircraft’s upper fuselage surface and provides RF navigation signals to GPS Receiver in L1 and L2 range. RF cable assembly – A coaxial cable that carries the RF signals from the antenna to the GPS receiver. The cable has a frequency range of 1 to 1.6GHz and usually has maximum attenuation values over this frequency range that must be periodically verified. Keyfill panel – Can be located in the nose landing gear wheelwell, avionics bays, cockpit or crew section of the aircraft and contains the following components: data fill connector, data fill indicator and data fill switch.

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End of Chapter 7 COMPASS AND INERTIAL NAVIGATION SYSTEMS Review Questions 7-1. A point that is defined by stated or implied coordinates is known as a ________.

A. direction B. reference C. position D. destination

7-2. The intended horizontal direction of travel is known as ________.

A. course B. direction C. position D. heading

7-3. In what two reference directions can you express bearings?

A. Radio bearing and visual bearing B. Magnetic heading and true heading C. True north and the direction in which you are pointing D. Relative and magnetic

7-4. The east/west geographical coordinate is known as ________.

A. latitude B. longitude C. meridian D. parallel

7-5. You measure longitude 180° east or west from what point?

A. The equator B. The pole C. The agonic line D. The prime meridian

7-6. The angle between true north and the direction of the earth’s magnetic field is known as ________.

A. lines of magnetic force B. direction C. variation D. agonic line

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7-7. How do you label variation?

A. Angle of declination B. East or west, respectively, of true north C. Difference in direction of heading and true north D. Comparison from north and true north

7-8. Magnetic influences cause what type of error in magnetic compasses?

A. Deviation B. Variation C. Swinging D. Deflection

7-9. The net result of both variation and deviation is known as ________.

A. compensating B. compass swinging C. compass error D. deviation plus

7-10. What should you do when variation and deviation have the same name to obtain compass error?

A. Add to get compass error B. Subtract to get compass error C. Subtract the smaller from the larger D. Divide the larger from the smaller

7-11. You can determine a position from the record of a previously known position, course, speed, and time traveled by what process?

A. Pilotage B. Magnetic dip C. Compass error D. Dead reckoning

7-12. What navigation system makes use of the physical laws of motion that Newton described three centuries ago?

A. Magnetic B. Inertial C. Observation D. Direction

7-13. What is the purpose of the four-gimbal system in Inertial Navigation Systems?

A. Provides pitch and roll for the aircraft B. Allows the platform to retain the original orientation regardless of aircraft maneuvers C. Shows the amounts of deflections in degrees D. Measures the rate of change in the motion of the aircraft 7-75

7-14. What is used primarily to provide visual information concerning aircraft attitude, steering, and navigation on a CRT display or mechanical dial indicator?

A. Azimuth Sensing B. Control Amplifier C. Horizontal Indicator D. Memory Unit

7-15. When operating the AN/ASN-50 attitude heading reference system, when do you use the compass mode?

A. When the displacement gyroscope is malfunctioning B. When operating at latitudes greater than 700 feet C. While in areas where the earth’s magnetic field has appreciable distortion D. When the flux valve fails

7-16. What is the operating principle of the Inertial Navigation System (INS)?

A. Acceleration is velocity B. Velocity with respect to time C. Newton’s first law of motion D. Acceleration and the pull of gravity

7-17. What must an object first experience before its state of rest or state of motion can change?

A. Acceleration B. Change of displacement C. Physical property change D. Change of motion

7-18. What is the result of acceleration being integrated over a specific period of time?

A. Direction B. Displacement C. Time differential D. Velocity

7-19. What is the purpose of an INS?

A. To keep track of acceleration B. To keep track of position C. To track the total distance traveled D. To detect and compute changes in displacement

7-20. What is the primary data source for the INS?

A. The digital computer B. The gyros C. The integrator D. The accelerometer 7-76

7-21. What allows the INS to determine total true accelerations in a horizontal plane for any movement in any direction?

A. Mounting on a stable element together with gyroscopes B. Mounting another accelerometer perpendicular to the first one C. Holding the sensitive axis of the accelerometer normal to the gravitational field D. Using integrators to convert the measured acceleration to aircraft position information

7-22. What prevents gimbal lock making the INS a true all-attitude system?

A. The stable element mount B. The gimbal axes C. The inner roll gimbal D. The synchros

7-23. How long is the period of oscillation for the Schuler pendulum?

A. 34.4 minutes B. 44 minutes C. 84.4 minutes D. 94 minutes

7-24. False accelerations sensed when the platform is torqued to maintain its plane of reference are known as ________.

A. centripetal corrections B. Coriolis corrections C. centripetal errors D. linear accelerations

7-25. Which are three basic external references that may be used to align an INS?

A. Surveyed lines, bench marks and bubble levels B. Terrestrial, celestial, and inertial C. Optical devices, electromechanical devices, and inertial sensing instruments D. Star trackers, radio sextants and transfer devices

7-26. The rate at which a gyro precesses is proportional to the angle between its spin axis and the spin axis of the ________.

A. platform

B. gyros C. accelerometer D. earth

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7-27. What INS solved the problem of operating an inertial system at the poles?

A. Wander-azimuth B. Gyrocompassing C. Centripetal D. Northpointing

7-28. Having to maintain the accelerometer referenced to a fixed point in inertial space is a significant disadvantage of the ________.

A. semianalytic system B. pure inertial navigation system C. analytic inertial navigation system D. strap-down system

7-29. What is the main advantage of the geometric inertial navigation system?

A. It is referenced to inertial space. B. The gyros are not torqued. C. The accelerometers remain in a plane tangent to the earth’s surface. D. The platform is aligned at the equator and then moves north.

7-30. What two types of compensators are used on flux valves?

A. Magnetic and nonmagnetic B. Universal screw and the loose magnet type C. Swinging and compensating D. Standby and magnetic

7-31. GPS provides specially coded satellite signals that can be processed in a GPS receiver, enabling the receiver to compute position, velocity, and ________.

A. size B. distance C. time D. frequency

7-32. What is the minimum number of observable satellites that is required to produce the most accurate position solution?

A. Two B. Three C. Four D. Five

7-33. What type of signal is transmitted from each satellite?

A. AC B. CA C. DC

D. RF

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7-34. GPS can give a position (latitude, longitude and height) directly, without the need to ________.

A. transmit signals to the user even when not using the signal B. compute positions in three dimensions C. measure angles and distances between intermediate points D. be observable from anywhere on Earth

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CHAPTER 8 AUTOMATIC FLIGHT CONTROL AND STABILIZATION SYSTEMS

Aircraft fly under many conditions. External conditions can alter the desired flight characteristics of the aircraft. To maintain the desired characteristics of the aircraft, the pilot moves the control surfaces either manually or automatically. You have already learned about indicating systems and instruments that supply the pilot with information on the performance of the aircraft. The pilot must be able to see and interpret each of these indicators and then react to get the desired performance. In high- performance aircraft, especially single-piloted aircraft, other flight duties require much of the pilot’s time. Navigation, communication, radar, and other special equipment are severely limited if the pilot has to work continually on the physical manipulation of the controls. In high-performance aircraft capable of supersonic flight, aircraft speed is so great that the pilot’s normal response time is far too slow. For example, by the time the pilot reacts to an indicator to position a control surface, the aircraft may already be out of control. Automatic flight control and stabilization systems ease the pilot’s workload and provide aircraft stability at all speeds. The information now flows directly to a flight control computer rather than to an indicator. This action lessens the time required to start a control movement to nearly zero. The result is increased stability. These systems also provide command controls by which the computer can control the aircraft in nearly any desired flight condition. Some automatic flight control systems are capable of flying the aircraft by radio navigation aids, correcting for wind, and making pilot-unaided landings. The term Automatic Flight Control System (AFCS) is used instead of the older term, automatic pilot, or the shortened version, autopilot. A reliable AFCS is necessary because pilots have duties other than moving the flight controls. However, regardless of how sophisticated the AFCS computer may be, the reasoning power of the pilot cannot be duplicated. LEARNING OBJECTIVES When you have completed this chapter, you will be able to do the following: 1. Explain the principles of flight for both fixed- and rotary-wing aircraft. 2. Recognize functions, operating principles and modes, including air data, flap position information and coordination inputs. 3. Identify AFCS components. PRINCIPLES OF FLIGHT To understand automatic flight control and stabilization systems, you must study the effects that the various controls have on the aircraft. Airman, NAVEDTRA 14014, contains a basic introduction to the principles of flight and flight controls. You should review this text before proceeding with this chapter. 8-1

Figure 8-1 — An airfoil.

An airfoil is any part of an aircraft designed to produce lift. Obviously, a wing is the primary airfoil on an aircraft, but propeller blades, tail surfaces, and even the fuselage itself are important airfoils. The design of a specific airfoil is determined by its job. All a irfoils have the basic elements shown in Figure 8-1. An airfoil consists of two nearly parallel surfaces, with one surface being more rounded than the other. As air passes over these two surfaces, the air passing over the rounded surface has farther to travel than the air passing over the flat surface. However, two particles of air leaving the airfoil’s leading edge at the same instant, one going over the rounded surface and one over the flat, arrive at the trailing edge at the same time. Therefore, you can infer that air passing over the rounded surface travels at a higher velocity than air passing over the flat surface. Bernoulli’s theory concerning the behavior of fluids, E = VxP, explains how pressure is changed and lift is produced. Here, E is the total energy produced by the airfoil passing through the air, V is velocity energy, and P is pressure energy. An airfoil that passes through air at a velocity of 50 feet per second and exerts a pressure of 10 pounds per square inch on the flat surface produces a total energy of 500 foot-pounds per square inch per second. E = VxP = 50x10 = 500 foot-pounds inch2/sec If the airflow velocity over the rounded surface is increased to 60 feet per second, and the total energy is unchanged, it exerts a pressure of 8.33 foot-pounds per square inch per second on the rounded surface. P =

=

= 8.33 foot-pounds inch2/sec The difference in the pressure between the rounded surface and the flat surface of the airfoil is called lift. In actual practice, the flat surface is not perfectly flat and causes some decreased pressure. The decreased pressure is negative lift. Negative lift is compensated for by the creation of high pressure on the flat surface. Air packed beneath the airfoil (dynamic lift) causes the high pressure. The true measure of lift remains the difference in pressure between the rounded and flat portions of the airfoil. Increased lift is the result of a larger pressure difference between the surfaces. The difference can be produced in two ways— by increasing the forward movement of the airfoil through the air (Figure 8-2), or by changing the angle of attack. Angle of attack is the acute angle between the chord line of an airfoil and its direction of motion relative to the air.

8-2

Figure 8-2 — Lift increases as velocity increases.

The chord of an airfoil is an imaginary straight line drawn from the leading edge to the trailing edge of the airfoil (Figure 8-3). As the angle of attack increases, the air strikes the leading edge closer to the flat portion of the airfoil. The distance air must flow over the rounded portion becomes even greater in relation to that flowing over the flat portion. This action causes a larger pressure difference and develops more lift. If the angle of attack increases too much, airflow over the airfoil’s rounded portion separates from the surface and becomes turbulent. Turbulence causes the pressure on both surfaces to become nearly equal, and the airfoil is said to stall.

8-3

Figure 8-3 — Constant velocity versus increasing angle of attack. Figure 8-4 — Induced drag. When producing lift, a secondary effect called drag is also produced. Drag produced by a lifting surface or airfoil is called induced drag (Figure 8-4). Induced drag develops in direct proportion to lift—when lift increases, induced drag also increases. 8-4

Figure 8-5 — Lift and drag change proportionately with the shape of the airfoil. At a given speed and angle of attack, a thick airfoil produces more lift and drag than does a thin airfoil. It follows that large, subsonic aircraft have thick wings to produce a great amount of lift at slow speeds. Supersonic aircraft must have very thin wings to decrease drag at high speeds. Many airfoils have devices attached to them to increase or decrease lift in various flight conditions or attitudes. These devices may mount on the leading edge, trailing edge, rounded surface, or flat surface. If the device attaches to the trailing edge by a hinge and has controls to move the trailing edge, you control lift by changing the angle of attack (Figure 8-5). When the trailing edge moves into the higher pressure air on the airfoil’s flat side, the angle of attack is effectively increased. This angle increase causes more lift and drag. Conversely, if the trailing edge moves into the airfoil’s low-pressure side, the angle of attack decreases. Lift and drag decrease accordingly.

In flight, each aircraft has certain forces acting upon it (Figure 8-6). To sustain flight at a constant altitude, the total lift of all airfoils must equal the aircraft weight. To change altitude you must change the total lift. If the aircraft weighs 10,000 pounds, 10,001 pounds of lift causes the aircraft to climb; 9,999 pounds of lift causes the aircraft to descend. 8-5

Figure 8-7 — Forces in a turn. Figure 8-6 — Forces on an aircraft. To fly at a constant airspeed, the forces of thrust and drag must be equal. When one force is greater than the other, the aircraft accelerates or decelerates.

To turn, place the aircraft in a bank angle (Figure 8-7). The lift developed by the airfoils ca n then be broken down into components of horizontal and vertical lift. The horizontal component of lift pulls the aircraft around in the turn. The vertical component of lift must be equal and opposite to gravity for the aircraft to remain at a constant altitude. (Note that total lift must be increased to prevent a loss in altitude.)

8-6

Figure 8-8 — Fixed-wing aircraft controls. When centrifugal force equals horizontal lift, the aircraft is in a constant-rate turn. For a faster rate of turn, increase horizontal lift by increasing the bank angle. When all lift is vertical to gravity, any turning motion is called a skid. FIXED-WING AIRCRAFT A fixed-wing aircraft is one in which the main lifting surface remains stationary with respect to the rest of the aircraft. Today’s fixed-wing aircraft are mostly electronically operated and have certain fixed surfaces or airfoils— wings, vertical stabilizers and Leading Edge Extensions (LEX)—that provide stability (Figure 8-8). In addition, these aircraft have movable control surfaces, one each per side—ailerons, Leading Edge Flaps (LEF), Trailing Edge Flaps (TEF), horizontal stabilizers and rudders. These surfaces permit the pilot to control the aircraft through a sophisticated integrated electronic flight control system.

Movement about the lateral axis of the aircraft (the axis that extends from wing to wing through the center of gravity) is pitch. To control pitch you use the horizontal stabilizers. Pitch is controlled by symmetrical deflection of the horizontal stabilizers. If the operator desires a nose-up attitude, apply an aft motion on the aircraft controller grip assembly (control stick). This causes the horizontal stabilizer’s leading edge to move down. Rotation about the lateral axis then causes the nose to lift, due to air on top of the wing to move faster than air on the bottom of the wing. Conversely, if you want to lower the 8-7

aircraft nose, apply a forward motion on the aircraft controller grip assembly. This causes the horizontal stabilizers leading edge up. Rotation about the lateral axis then causes the nose to lower. Movement of the aircraft about the longitudinal axis (from nose to tail) is known as bank or roll. Roll is controlled by differential deflection of the LEFs, TEFs, ailerons, and horizontal stabilizers working simultaneously. Per side; the TEFs, ailerons and horizontal stabilizers deflect in the same direction and the LEF deflect in the opposite direction. Moving the aircraft controller grip assembly left or right produces the required differential deflections for the present angle of attack, altitude, and airspeed. For the aircraft to enter a left bank, the angle of attack of a portion of the right wing must increase. You accomplish this by symmetrically lowering the right aileron, right TEF, and right horizontal stabilizer to increase the lift on that wing. The left aileron, TEF and left horizontal stabilizers deflect asymmetrically from the right side and ris e to decrease the lift on that wing. The aircraft then rotates about its longitudinal axis until the ailerons and horizontal stabilizers are neutralized in some angle of bank. The rudders are symmetrically deflected to minimize sideslip during roll maneuvers. The aircraft remains in that bank angle until you again move the ailerons. Refer to Figure 8-7. Whenever the aircraft is in a bank, lift developed by the wings is displaced from the vertical position. If you do not increase lift, its vertical component is insufficient to maintain the aircraft at a constant altitude. A change in pitch attitude to increase the angle of attack of the wings is used to prevent a loss in altitude. A few degrees of bank angle require an imperceptibly small pitch change. A 90-degree bank in level flight (no altitude change) is theoretically impossible because of the absence of vertical lift. As the bank angle changes, coordination between ailerons, horizontal stabilizers and TEFs are necessary to prevent a loss in altitude. Notice in Figure 8-7, the LEFs are canted downward; not level with the fuselage and the left aileron is full down providing lift and the right is streamlined as the aircraft prepares to land on the deck. To return to level flight, increase lift on the left wing by lowering its aileron into the higher pressure area beneath the airfoil. Reduce lift on the right wing by raising its aileron into the lower pressure area at the top of the airfoil. As the wings become level, neutralize the ailerons. Movement about the vertical axis is yaw. Usually this movement is undesirable in an aircraft. Use the rudder to correct any tendency of the aircraft to yaw. The rudder is NOT used to turn the aircraft (change heading). When placing the ailerons into the airstream, the aircraft has a tendency to yaw. When banking to the right, the aircraft produces more lift and drag on the left wing, and less lift and drag on the right wing. Even though the intention is to turn to the right by going into a right bank, the initial tendency is for the nose of the aircraft to go to the left. This movement happens because of the increased drag on the left wing and decreased drag on the right wing. This effect is adverse yaw; you compensate for it by displacing the rudder in the same direction as the intended turn. If an aircraft in a turn tends to slip into the inside of the turn or skid to the outside of the turn, this is also yaw. You also compensate for it by using the rudder. Many other things may cause yaw, such as the engines on one wing of a multiengine aircraft producing more power than the engines on the other wing.

8-8

Figure 8-9 — Flight controls: (A) Horizontal stabilizer and aileron, (B) Rudder. You can see, then, when you place a fixed wing aircraft in a bank angle, coordination between all three controls—ailerons, horizontal stabilizers, and rudder—is necessary. The pilot accomplishes control of horizontal stabilizers, ailerons, and rudder through the use of a control stick and rudder pedals (Figure 8-9). To operate the ailerons, move the control stick right or left in the direction of the intended turn (Figure 8-9, view A). Aft force on the control stick raises the trailing edge the horizontal stabilizers and causes the nose to pitch up. Forward pressure on the control stick lowers the trailing edge of the horizontal stabilizers and causes the nose to pitch down. You use your feet to operate the rudder pedals (Figure 8-9, view B). Pressure on either rudder pedal causes rudder deflection in that direction.

Weight distribution in an aircraft varies for many reasons. For example, fuel may be used faster from one wing tank than from the other, allowing that wing to become lighter. In large aircraft where crew members or passengers walk around, the balance point, called the Center of Gravity (CG), shifts whenever someone changes position in the aircraft. As fuel is used, the aircraft gross weight reduces. The pilot must reduce the angle of attack of the wings to lessen lift and prevent a gain in altitude. The pilot must use control pressures to compensate for these unbalanced flight conditions. Several methods are used to reduce these control pressures and to ease the pilot’s workload. In newer aircraft, the most common method is the setting autopilot. Look at Figure 8-8. The figure shows aircraft control surfaces. Setting autopilot is desirable on this particular aircraft because flight computers and the air data system compensate for unbalanced conditions and automatically adjust flight surfaces. When pilots must exert a force on the control stick, they can use the trim control to relieve that force. For instance, when they must hold left rudder pressure to prevent yaw movement to the right, they can move the rudder trim control to the right. The airflow on the vertical stabilizer strikes the rudder and moves the complete rudder a little to the left. Since the rudder trim control now supplies the required rudder pressure, the pilot no longer has to hold pressure on the rudder pedals. 8-9

ROTARY-WING AIRCRAFT An aircraft that derives its main lifting force from a horizontally driven propeller device (rotor) is a rotary-wing aircraft. The most common rotary-wing aircraft is the helicopter. For lift to happen there must be relative motion between an airfoil and an air mass. Therefore, the major advantage of a rotary-wing aircraft is its ability to maintain zero or very low airspeed while the wings (rotors) are still creating lift. Forces acting on a rotary-wing aircraft are identical to those acting on a fixed-wing aircraft (Figure 8-6). You must also control the rotary-wing aircraft about the vertical, longitudinal and lateral axes, as shown in Figure 8-8. In the conventional helicopter, the main and tail rotors are engine driven. Remember the e arlier discussion on airfoils. You increase lift either by increasing the speed of the airfoil through the air or by increasing the angle of attack of the airfoil. In helicopters, the airfoil’s angle-of-attack is known as blade pitch. When the rotor speed is constant, the pilot maintains complete control of the aircraft by varying the pitch of the rotor blades. Figure 8-10 shows helicopter flight controls. The pilot operates collective control with the left hand, cyclic control with the right hand, and rudder control with the feet. The collective and cyclic controls command the main rotor. Operation of the rudder control changes the blade angle of the tail rotor.

In helicopter flight (except hovering flight), the main rotor provides altitude, bank, and directional control through use of the collective and cyclic controls. The tail rotor prevents the main body of the helicopter from spinning (yawing) with the torque of the main rotor. It prevents yaw in a similar way as the fixed-wing aircraft rudder. Collective control maintains or changes altitude. Moving the collective control causes an equal change in pitch (angle of attack) of all main rotor blades. Also, through a Figure 8-10 — Helicopter flight controls. 8-10

Figure 8-11 — Coning angle increases as load increases. mechanical mixer, collective control automatically changes tail rotor pitch to compensate for increases or decreases in main rotor torque. Since the rotor blades are somewhat flexible, the more collective control applied, the more an action called coning takes place. As the blades rotate, they take the shape of a cone (Figure 8-11). The speed and pitch of the blade tips determine the coning angle. With a constant pitch, the faster the rotor blades turn, the more horizontal the blades become because of centrifugal force. As the blade pitch increases, lift also increases, and the coning angle increases because of the load on the blades.

In hovering flight, cyclic controls contain pitch and roll, which create forward and sideward motion, respectively. The collective controls maintain altitude and the rudder pedals control heading. 8-11

Figure 8-12 — Flapping angle creates horizontal lift: (Frame 1) Hovering flight; (Frame 2) Forward flight. The cyclic stick provides pitch and directional control of the helicopter. When the pilot applies pressure to the cyclic stick, each blade moves to a specific pitch angle as it passes a certain point in its rotation (Figure 8-12, frame 1). During forward flight, the blade pitch is greatest as it passes the 90-degree position. The blade pitch is least at the 270-degree position, and equal at the 0-degree and 180- d egree positions. To turn the aircraft, lateral motion of the cyclic stick causes blade pitch to be greatest at 0 and 180 degrees, and least at 90 and 270 degrees. Since the blades form a spinning mass, the gyroscopic principle of precession occurs 90 degrees in the direction of rotation from where the lifting force is applied. The coning angle remains the same. However, the cone tilts in the direction of the desired flight path, creating what is called the flapping angle. You can again break down lift into its vertical and horizontal components. Look at Figure 8-12, frame 2. To maintain altitude, the vertical lift is increased until it is equal to gravity. With the cone at a flapping angle, the helicopter accelerates in the desired direction until drag is equal to horizontal lift. To accelerate the helicopter in a forward direction, move the cyclic control stick forward. You also must make a corresponding increase in collective control to maintain altitude. As the collective control increases, torque on the main rotor blade increases. This action makes the helicopter tend to rotate in the direction opposite to the rotor blade rotation 8-12

(nose right). A mechanical mixer automatically changes the pitch of the tail rotor to overcome the right turning tendency (skid). To turn the helicopter (change heading), place the cyclic control stick to the right or left. Flapping action of the main rotor blades causes the cone to tilt in the direction of the desired turn (Figure 8-12, frame 3). As with the fixed-wing aircraft, the pilot must maintain coordination in a turn by using the rudder pedals to prevent skid or slip. Also, the pilot must adjust collective control to prevent a loss in altitude. In hovering flight, the pilot uses the rudder pedals only to turn the helicopter, thus producing a skid. AUTOMATIC FLIGHT CONTROL SYSTEMS (AFCSs) In the human body, signals to move us from place to place start with our five senses as they reference outside conditions. The brain processes these signals and sends them through the nerves to the muscles. The body then does its required movement by muscle power. Similarly, most AFCSs have their component parts divided into three major groups—sensors (information inputs), amplifier/ computer, and output units. The sensors originate the signals as they are acted upon by outside references. They only sense changes and do not have sufficient power to make corrections. The amplifier and computer are the brains for the AFCS. They receive the weak signals from the sensors, which in most cases are synchros, and determine how much and in which direction correction is necessary. The synchro signals are usually in millivolts, but the correct strength needed is in volts. Therefore, the amplifier increases the weak signal to a workable voltage. The value of the synchro signal depends on the amount of rotor displacement with respect to the stator from the null position. The direction of rotor displacement from the stator determines the direction of the correction. Most amplifiers have at least two stages of voltage amplification—one stage of phase discrimination, and another stage where power amplification takes place. Other types of amplifiers control the voltage to control valves in hydraulic servos. The output unit is the muscle of the AFCS. It consists of an electro/hydraulic booster package. There is a booster package for each control surface— rudder, aileron, and elevator. The boosters also assist the pilot in manual control of the aircraft. Summing up the major groups, the sensors send a small signal to the amplifier/computer when a displacement occurs. The amplifier/computer amplifies the weak signal to a workable voltage and sends it to the output unit. The output unit changes the electrical energy to mechanical displacement. It then moves the control surfaces by an amount commanded by the sensor signal. AFCS COMPONENTS The AFCS consists of many controls, sensors, and electromechanical components. To understand the entire system, you need to know what each component of the system does. In this section, you will read about the components that make up the AFCS. Electrical/Electronic Components The electrical/electronic components make AFCSs work. The control panel is used to program any pilot-desired maneuver that is within the capability of the system. Originally, AFCS systems were very limited. They supplied only one-channel operation to the ailerons to keep the wings level. Newer aircraft receive signals from other aircraft systems. Radar and barometric altimeter signals couple with the AFCS to maintain the aircraft at a constant altitude. Some aircraft use signals from data-link systems to fly the 8-13

Figure 8-13 — AFCS control panel. aircraft during approaches and landings. Some fighters have the fire control system tied in so the aircraft can fly automatically to an enemy aircraft. Fighter bombers with a weapons control system tie-in can fly automatically to the target and release their weapons at the proper time. Long-range patrol aircraft have their Anti-submarine Warfare (ASW) systems tied into their AFCS. The AFCS operates the rudder, the elevator, and the ailerons by using various sensors and electrically controlled hydraulic servos. Before engagement, the AFCS is synchronized with the flight control surfaces to prevent sudden or violent maneuvers. The system senses deviation from the reference flight condition and causes the aileron control to maintain either a reference bank angle or a heading. It also causes the elevator control to maintain either a reference pitch angle or an altitude. Additionally it makes the rudder control coordinate turns and provides automatic yaw damping. CONTROL PANEL— The AFCS control panel contains all the switches and controls necessary for the pilot to select/control the autopilot modes. Control panels are designed for the particular type and mission of the aircraft. Some control panels are simple, while others are complex. Figure 8-13 is an example of an AFCS control panel. Here, the switches serve as manually operated interlocks in setting up the circuitry to engage the various AFCS modes of operation. This control panel has six switches. They are labeled as follows: ACL/OFF/PCD, ALT/OFF/MACH, HDG OFF/NORM/ROLL CMD, AUTO/STAB-AUG, ON/OFF, and ATTITUDE REF. The ACL/OFF/PCD, ALT/OFF/MACH, and HDG OFF/NORM/ROLL CMD switches are solenoid-held toggle switches. Each has a lever-lock toggle feature that prevents accidental engagement in the operate position. The AUTO/STAB-AUG and ON/OFF switches are solenoid-held, spring-loaded switches. When not engaged or when no power applied, the switches return to the STAB-AUG and OFF positions, respectively. The ATTITUDE REF switch is a miniature, positive-break, aircraft-type toggle switch. AIR NAVIGATION COMPUTER (ANC)— All operating functions of the AFCS channel through the air navigation computer. It is sometimes called the amplifier computer, and it is the heart of the entire system. The ANC modifies the combined signals supplied by various sensors and command controls to develop output signals. The output signals control the aircraft’s ailerons, rudder, and elevators. By use of these flight controls, the aircraft automatically maintains a reference attitude, heading, and altitude. Also, it can maneuver in a coordinated manner in response to turn and pitch control settings on the control panel. 8-14

Figure 8-14 — (A) Air navigation computer; (B) One-channel amplifier/computer. The physical appearance of the ANC depends on the type and mission of the aircraft. Figure 8-14 shows two types of ANC.

Figure 8-14, view A, shows an ANC that consists of an equipment rack and seven amplifier modules, which are listed below. 1. Roll servo amplifier 2. Pitch servo amplifier 3. Yaw servo amplifier 4. Roll computer amplifier 5. Pitch computer amplifier 6. Heading computer amplifier 7. Command coupler Each of the seven modules contains subassemblies and sub-subassemblies. Some of these are interchangeable between modules. The roll, pitch, and yaw servo amplifiers are identical. The other modules have individual differences. The computer, through an interlocking relay arrangement in conjunction with the control panel mode selection 8-15

Figure 8-15 — Control stick steering components. switches, controls signal switching operations. A calibration board on the front of the ANC provides gain adjustments of the major system parameters. Figure 8-14, view B shows a one-channel amplifier/computer. Normally, this particular type of autopilot computer consists of three individual amplifier/computer modules (one for each control surface)—aileron channel (roll), rudder channel (yaw), elevator channel (pitch). This one-channel computer accomplishes analog computations by using servomechanisms. These servomechanisms consist of electromechanical computer cards and electronic amplifier cards mounted in the amplifier/computer. In addition, a transformer board and a resistor board provide summing networks. The networks combine the various signals supplied to and generated within the unit. An interlocking relay arrangement is included to perform most of the switching control in the AFCS. CONTROL STICK— Control stick or control wheel steering is used on some aircraft to control the aircraft electronically through the AFCS. On fighters, the signals are generated in a unit such as the one labeled “motional pickup transducer” in Figure 8-15. When the AFCS is on and the control stick is moved left or right, pressure on the roll force switch momentarily disengages the roll channel of the AFCS. The pilot then controls the roll attitude of the aircraft through regular stick control. When the pilot releases stick pressure, the force switch opens, allowing the AFCS to reengage roll. If the bank angle is above a given angle (for example, 5 degrees), the AFCS will maintain the bank angle. If the bank angle is below the given angle, the AFCS automatically returns to wings level. The pitch force switches close when a fore or aft pressure is on the control stick. This action momentarily disengages the AFCS. The stick pressure also couples a signal through the E pickoff transformer that is labeled “force sensor,” as shown in Figure 8-15. The signal couples with the AFCS pitch channel. Depending on the direction of the stick pressure (fore or aft), the aircraft either climbs or dives. 8-16

Figure 8-16 — Signal generator pickoff operation. Electrical/Electronic Sensors Many electrical and electronic sensors provide input to the AFCS. This section of the Rate Training Manual (RTM) includes a review of the sensors already discussed, and it introduces you to other sensors in the AFCS. SIGNAL GENERATOR PICKOFF (SYNCHRO)—Figure 8-16 illustrates the principal of operation for a signal generator pickoff. The pickoff consists of a stator and rotor. The stator is ring-shaped and has four poles. Each pole has a primary and secondary winding. The rotor has no windings. It serves to change the reluctance of the magnetic flux path between the stator poles. The primary and secondary windings are connected so the voltages induced into the secondaries are of opposite polarity on adjacent poles. However, opposite poles have the same polarity.

The voltage output of the secondary is zero if the rotor is in its neutral position (Figure 8- 16, view A). Repositioning the rotor (Figu re 8-16, views B and C) makes a stronger magnetic field on a single pair of poles. This results in a voltage output on the secondary winding. The amplitude of the output voltage is proportional to the amount of rotor displacement—the greater the displacement, the greater the amplitude. The direction of rotor movement determines the polarity of the output voltage. The polarity will be either in phase with the input voltage or 180 degrees out of phase with it. 8-17

Figure 8-17 — Vertical gyro components. Synchro construction allows for more accurate voltage production than angle signal, effective through 360 degrees of rotation. For more information about synchros, servos, and gyros, you should refer to Navy Electricity and Electronics Training Series (NEETS), Module 15, Principles of Synchros, Servos, and Gyros. GYROS— There are several different gyros used with the AFCS. You have already learned about most of them earlier in the RTM. The following paragraphs provide a review of gyros and how they specifically affect the AFCS. Vertical Gyro—The vertical gyroscope (Figure 8-17) is an electrically driven gyro that provides pitch and bank attitude references for the AFCS. It can also provide pitch and bank attitude references for servo indicators and other systems of the aircraft. It has enough signal load capacity to sustain several systems at the same time.

The vertical gyro is a two-degree-of-freedom gyro. This gyro is termed the vertical gyro because it is continuously erect with its spin axis vertical to the surface of the earth. The spin axis provides a vertical reference for measurement of aircraft bank angle and pitch angle. Pitch and roll gimbals isolate the gyro from its housing and the aircraft. Thus, the aircraft can bank or pitch while the gyro remains vertical because of gyroscopic action (Figure 8-18). A pitch synchro, mounted on the vertical gyro’s pitch pivot, continuously senses the relative pitch angle between gyro and aircraft. Similarly, a bank synchro, mounted on the vertical gyro’s bank pivot, continuously senses the relative bank angle between gyro and aircraft. 8-18

Figure 8-18 — Vertical pitch and roll reference. The gyro motor rotates at a speed of about 20,000 RPM. A solenoid-operated friction brake prevents tilting and tumbling when the gyro motor is idle and during the initial starting torque. Synchros mounted on the gyro detect motion between the gyro, the gyro gimbal, and the gyro case. A synchro generates a weak signal when its rotor is displaced from its stator. The pitch synchro mounts with its rotor on the gyro pivot and its stator on the gyro gimbal. Thus, the synchro measures the displacement angle between the gyro and the gimbal. This is the pitch displacement angle from the vertical reference. The bank synchro mounts with its rotor on the gyro gimbal pivot and its stator on the gyro case. Thus, the synchro measures the displacement angle between the gimbal and the case. This is the bank displacement angle from the vertical reference. As the pitch attitude of the aircraft changes, the gyro case and gimbal turn about the gyro rotor. The pitch synchro rotor is held rigid in space by the gyro. The gyro motor generates voltages in the pitch synchro proportional to the aircraft’s pitch angle with respect to the surface of the earth. During changes in pitch attitude, the bank synchro remains at null since the gyro gimbal is not free to rotate with respect to the case in pitch. Therefore, it tilts with the case. As the bank attitude of the aircraft changes, the bank synchro stator turns about the bank synchro rotor. The gimbal is held rigid in space by the gyro because it is not free to rotate with respect to the gyro in bank. Thus, the bank synchro generates voltages proportional to the bank angle between the aircraft and the surface of the earth. 8-19

Figure 8-19 — Three-axis rate gyro orientation diagram. When the aircraft yaws, the case, the gimbal, and the gyro stator turn about the rotating gyro motor. This has no significant effect upon the relative positions between the gyro and the case. As a result, there are no pitch and bank synchro output voltages in response to changes in yaw. Three-Axis Rate Gyro—Rate gyros sense the rate of movement of an aircraft about its vertical, lateral, or longitudinal axis. They provide synchro signal outputs representing yaw rate, pitch rate, or roll rate to the air navigation computer. These units are sometimes very similar in appearance to the rate switching gyro, but they provide entirely different information to the system. Physically, rate gyroscopes are the same. They differ only in respect to calibration, alignment, range, sensitivity, and natural frequency. Each gyro measures angular rate. It uses the proportional precessional torque generated by the rate of movement about the gyro-sensitive axis (Figure 8-19) to make these measurements.

Internally, each rate gyro consists of a small viscous-damped, single-degree-of-freedom gyro with a differential transformer pickoff (Figure 8-20). The gyroscopic element of each gyro is the rotor of a synchronous motor. The rotor mounts in a gimbal frame and spins at high speed about its spin axis. The gimbal is flexible and free to rotate about an output axis. This axis is perpendicular to both the spin axis and the input axis. A torsion- restoring spring that couples the gimbal to the case limits the rotational freedom about 8-20

Figure 8-20 — Rate gyro axis orientation. the output axis. The gyro gimbal carries the pickoff rotor on an extension along its output axis. The pickoff rotor senses the relative angular displacement of the gimbal and case. With the pickoff rotor in its zero or neutral position, the mutual inductance is zero. The current flowing in the pickoff primary causes essentially no voltage in the secondary (output) winding. As the pickoff rotor is turned one way or the other about its output axis by gyro gimbal deflection, a proportional mutual inductance is introduced. The polarity of the inductance (positive or negative) depends upon the direction of deflection from the neutral position. Hence, the current flowing in the primary produces a voltage proportional to this mutual inductance in the pickoff secondary. The output voltage is proportional to the aircraft’s angular velocity input to the gyro in the particular axis. COMPASS INFORMATION— Normally, compass information for the AFCS is supplied by the aircraft compass system or the Inertial Navigation System (INS). However, some compass information is developed for the AFCS. The compass system/INS incorporates a gear train to drive several synchros. The gear train is driven by a motor generator unit that aligns to aircraft heading. Of the several synchros, one provides heading information to the pilot’s compass indicator. Another synchro, which is attached to the gear train through a clutch, provides a clutched heading. When the AFCS is not engaged, the clutch remains deenergized, with its rotor spring loaded to an electrical null condition. When the pilot engages the AFCS, the clutch engages the engaged heading to establish a reference heading for the system. If the aircraft drifts off heading, the gear train drives against the spring tension on the rotor, generating an electrical signal. This signal goes to the aileron channel, much like the signal generator pickoff operation (Figure 8-16). Limiters in the AFCS prevent the bank angle from becoming excessive when large heading errors are detected. Another type of compass information is derived from the heading indicator in the flight station. The pilot selects a desired heading on the face of the indicator. The difference between the selected heading and the actual heading becomes an error signal to the AFCS. This error signal causes the aircraft to turn to the desired heading. 8-21

A radio navigation aid can also supply heading information to the AFCS. If the pilot desires to fly to a selected ground station, the radio receiver develops a signal to produce the desired ground track directly to the station. AIR DATA INFORMATION— Changes in the speed of an aircraft also affect the effectiveness of the control surfaces. At a given altitude, slow speeds require more control surface movement than high speeds require to accomplish the same maneuver. The pilot maintains (or changes) the altitude by referencing the altimeter. The AFCS can also maintain a constant altitude. To accomplish this task the AFCS uses altitude data supplied by an air data sensor or Air Data Computer (ADC) as the reference altitude. Airspeed— Control surface signals are modified by a gain control unit to compensate for changes in airspeed. This unit uses the difference between ram pressure and static pressure. A mechanical schematic of the gain control unit is shown in Figure 8-21. Here, you can see that as airspeed increases (ram air pressure increases), the bellows cause the spring to become more compressed. This compression allows the armature to move each potentiometer’s sliding arm to modify the control surface signals an amount representative of the change in airspeed of the aircraft. When airspeed decreases, the armature moves to the right, selecting a different amplifier gain. The opposite occurs for increases in airspeed.

Altitude— The AFCS includes an altitude control feature to maintain the aircraft at a fixed altitude. The altitude controller consists of an aneroid, a mechanism for transmitting and magnifying the motion of the aneroid, and a solenoid-operated clutch. It also includes a synchro transmitter and a centering device for returning the synchro transmitter rotor to the null or no-signal position. Some aircraft do not use an altitude controller. In place of an altitude controller, the AFCS uses signals from the ADC. Figure 8-22, view A, shows a three-quarter view of a barometric altitude control. The outside appearance of controls of this type varies, depending upon the manufacturer; however, the working parts are similar. Figure 8-22, view B, shows the internal parts of a barometric altitude control, and view C shows a simplified mechanical schematic.

Figure 8-21 — Mechanical schematic of a gain control unit. 8-22

Figure 8-22 — (A) Barometric altitude control; (B) Internal parts; (C) Simplified schematic.

The aneroid consists of two diaphragms sealed internally at standard (sea level) barometric pressure. The two diaphragms connect in tandem to a single pushrod and are mounted in an airtight case (Figure 8-22, view B). A tube connects the case to a source of static air pressure. The diaphragm pushrod mechanically links to one of the clutch plates. This linkage consists of a lever, a pivoted shaft to which a sector gear is a ttached, and a pinion gear. When the aircraft deviates from the barometric pressure altitude to which the altitude control switch is set, the aneroid diaphragms move. This motion is transmitted through the linkage to displace the rotor of the synchro transmitter. This displacement generates a signal in the synchro transmitter stator. The signal is applied to the elevator channel to return the aircraft to the pressure altitude indicated by the aneroid. When the aircraft reaches the correct altitude, the synchro transmitter signal becomes zero, and normal AFCS operation resumes. When the altitude control switch is off, the magnetic clutch and the centering device actuating coil deenergizes. The deenergizing of the coils opens the clutch to disengage the synchro transmitter rotor from the aneroid mechanism. The centering yoke, by spring action, closes on the synchro transmitter rotor shaft lever to return the rotor to the null or no-signal position. In this way, the synchro transmitter rotor is always at the no- 8-23

Figure 8-23 — Flap position transmitter and schematic. signal position when altitude control is not selected. Since the clutch is disengaged, the aneroid is free to move. This allows the pilot to engage the altitude control at any time. Regardless of the aneroid position, the altitude that it senses is the one used as the reference altitude. It is not necessary to wait for synchronization or alignment. FLAP POSITION INFORMATION—When the flaps are lowered on some aircraft, the increased lift causes the aircraft to gain altitude (normally called ballooning). Ballooning is undesirable, and it is counteracted by using nosedown pressure on the flight control. When the AFCS is engaged and the flaps are lowered, automatic nosedown force is applied to the elevator. Flap position is detected by the use of a flap position transmitter. The flap position transmitter consists of two synchro transmitters with a single input shaft (Figure 8-23). The synchro transmitters supply flap position information to the AFCS elevator channel and the external flap position indicator.

ACCELEROMETER TRANSMITTER— The normal accelerometer (Figure 8-24, view A) generates a signal proportional to normal vertical acceleration. This signal is used for altitude or Mach hold vertical path damping and as the g-command reference. The unit consists of a cast housing assembly, a sensitive element assembly, bellows, and c alibration resistors (R51, R52, and R53). The sensitive element assembly has an E pickoff, an armature and armature support, flexure springs, and a backplate. When assembled, the sensitive element assembly and bellows are sealed inside the housing, which is filled with damping fluid. The metal bellows allow the volume of the damping fluid to change with temperature and pressure variations. The damping fluid provides viscous damping during motion of the armature. The sensitive element is mechanically biased to produce zero output when mounted in the correct position and subjected to the normal gravity force of 1 g. 8-24

While you read this section, refer to Figure 8-24. As the aircraft accelerates in the sensitive vertical direction, the suspended armature tends to remain behind due to its inertia. This reaction varies the reluctance of the magnetic circuit set up by the E pickoff windings and armature (view B). The armature completes the magnetic circuit through a small air gap. The relative motion between the armature and E pickoff varies the reluctance through the signal output windings. This variation results in a signal that is proportional to acceleration. When operating, the output voltage is either in phase or 180 degrees out of phase with the excitation, depending on the direction of acceleration. COORDINATION INPUT— In some aircraft, a dynamic vertical sensor detects lateral accelerations (slip or skid) of the aircraft. The sensor supplies a signal to position the rudder to correct the slip or skid, coordinating the turn. The signal is proportional to the amount of the aircraft deviation from the vertical axis of the aircraft. In other aircraft, a horizontally mounted accelerometer aligned with the lateral axis of the aircraft provides the same information. Only the dynamic vertical sensor is covered in this RTM.

Figure 8-24 — Accelerometer transmitter. 8-25

Figure 8-25 — Cutaway view of a dynamic vertical sensor.

The dynamic vertical sensor consists of a viscous-damped pendulum mechanically connected to the rotor shaft of a transmitter synchro. The cutaway view of the sensor (Figure 8-25) shows the mechanism assembly, which includes a synchro transmitter with a pendulum and vane assembly attached to the rotor. The vane moves in an oil-filled chamber. The damping effect of the fluid gives a long-term sensing characteristic that makes the unit relatively insensitive to transient oscillations. The damping chamber also limits displacement of the pendulum to 10 degrees either side of the center position.

A pin in the housing fits into a slot in the mechanism shell. This positions the mechanism to align the synchro rotor with the longitudinal axis of the aircraft when installed. The sensor functions in the same manner as the ball in a turn-and-bank indicator. The ball gives a visual indication of slip or skid resulting from lateral acceleration. The sensor provides a signal output of this condition. Figure 8-26 shows a diagram of the forces acting on the aircraft in a turn. Refer to this figure as you read this section. In a coordinated turn, the vertical and lateral forces resolve into a vector perpendicular to the span of the aircraft. When the aircraft is turning with the two forces in balance, the ball is centered; the pendulum in the sensor gives a null output. When the aircraft bank angle is too large for the turn rate, the balance is upset (Figure 8-26, view B). The ball moves away from the center toward the inside of the turn, and the pendulum moves the synchro rotor from the center null position. The rotor displacement produces a signal with magnitude proportional to the displacement angle and signal polarity corresponding to the direction of displacement. The unbalanced condition results from a sideways accelerating force, causing the aircraft to slip toward the inside of the turn.

8-26

Figure 8-26 — Dynamic sensor pendulum positions: (A) Coordinated turn; (B) Slip; (C) Skid. When the aircraft is insufficiently banked for the turn, an acceleration acts toward the outside of the turn (Figure 8-26, view C). The ball in the turn-and-bank indicator moves f rom the center, and the pendulum in the dynamic vertical sensor is displaced from null in the direction corresponding to the ball. This gives a signal whose polarity is opposite to that of the signal when the aircraft was in a slip. The signal is fed to the rudder channel for the right or left rudder to coordinate the turn. Since the pendulum is unaffected by transients, the rudder adjustment is on a comparatively long-term basis. Hydraulic Components Hydraulic systems provide the physical power to move the flight control surfaces. All electrohydraulic (AFCS) actuators work in the same manner. A brief discussion of electrohydraulics is presented in the following paragraphs. 8-27

Figure 8-27 — Hydraulic booster block diagram. ELECTROHYDRAULIC SERVO ACTUATORS — Electrohydraulic servo actuators (hydraulic booster packages) are discussed in chapter 4 of this RTM. Modern aircraft use several types of actuators, with each booster actuator having at least two modes of operation—manual mode and electrical signals. The first mode is the manual mode. Its primary purpose is to aid the pilot in manually positioning the control surfaces. Control surfaces on large aircraft are much too large to move unaided. On smaller, high-speed aircraft, the high air pressure makes it nearly impossible to move the controls unaided. In the manual mode of operation, the hydraulic booster package is connected between the pilot’s control stick and the control surface. It provides hydraulic assistance to the pilot in a similar manner that power steering aids the driver of a car or truck. The second mode of the hydraulic booster package uses electrical signals from the AFCS to move the flight control surfaces. In this mode, the booster package connects the AFCS to the control surface and provides the power to move the surface. Synchro devices on the boost package provide feedback signals to the AFCS (Figure 8-27). 8-28

The surface position transmitter sends the AFCS a signal representing the amount and direction of control surface displacement from the streamline position. This signal acts as a follow up to prevent overshoot of the controls. Also, it returns the control surface to the trimmed condition as the original signal returns to zero. The modulating piston is displaced only when the control surface is in motion. The position of the modulating piston is monitored to sense the control surface rate of movement. This action generates a signal to damp control surface movement. Hydraulic load sensors determine the amount of pressure the AFCS is applying to the control surface so the pilot can properly trim the aircraft before disengaging the AFCS. If the aircraft is not properly trimmed and the control pressure is suddenly relieved, the control surface moves rapidly, causing sudden aircraft movement. Some aircraft use both manual and AFCS modes simultaneously. Aircraft stabilization is provided from the AFCS, while the pilot manually controls the aircraft. All flight control systems have a method of disconnecting the booster package. This method gives the pilot manual control of the flight control surfaces if the booster malfunctions or failure of the hydraulic system occurs. AUTOMATIC TRIM—Some AFCS systems incorporate automatic trimming. When a signal is present in the control channel, there is an unbalance in fluid pressure at the input to the hydraulic booster. The hydraulic load sensor (Figure 8-28) detects this unbalance. Its signal is amplified and drives the trim servomotor. The engagement clutch engages when the AFCS is operating and the automatic trim system is functioning properly. The slip clutch allows the pilot to override the automatic trim in case of a malfunction. Most AFCSs use auto trim in the pitch channel; however, it can be used for yaw and roll as well. 8-29

Figure 8-28 — Automatic pitch trim block diagram. NOTE The AFCS channel shown in Figure 8-29 represents the basic design of the channel and is for instructional use only. For further study of the AFCS, you should refer to the Maintenance Instruction Manual (MIM) for your particular aircraft.

THEORY OF OPERATION The AFCS has three main control channels to control movement of the aircraft about its axis. These channels control the yaw (rudder), roll (aileron), and pitch (elevator). Each of the control channels has similar functional equipment groupings. The groupings include controls, sensors, signal coupling circuits, AFCS servo loops, and aircraft flight controls. Each channel of the AFCS supplies control signals to the flight controls and receives error signals from the sensors. The sensors and the signal coupling circuits in use in each control channel depend on the mode of operation. The AFCS servo loops and the aircraft flight control system provide the final amplifying link to move the flight control surfaces. Look at Figure 8-29. Each AFCS control channel is basically the same; however, its design depends on the type and mission of the aircraft.

8-30

Figure 8-29 — Basic AFCS control channel.

The Air Navigation Computer (ANC) receives various error and control signals. These weak signals are coupled (summed), modified, and amplified to develop the control surface command signal to drive the electrohydraulic actuators. The actuators move a ce rtain direction for a specific distance, depending on the command signal’s polarity and magnitude. The flight control surfaces (rudder, aileron, and elevator) are mechanically linked to the electrohydraulic actuators. As the control surface moves, a position transmitter synchro d evelops a feedback signal having the opposite polarity to the error signal. The magnitude of the feedback signal increases as the control surface displacement increases. When the error signal and feedback signal are equal and opposite in magnitude and polarity, the control surface will no longer move. Movement of the control surface causes the aircraft to displace about its axis (or reference). This movement corrects the original error signal as sensed by the sensors. Without any error signal inputs to the control channel of the amplifier/computer, only the feedback signal is present. The feedback signal is opposite in polarity and magnitude to the original error signal. This drives the electrohydraulic actuator an equal distance in the opposite direction and brings the control surface back to the null or reference position. 8-31

AFCS MODES This part of the chapter contains a brief description of the various modes available in a typical AFCS. The pilot selects one of these modes on the AFCS control panel by moving the control stick or using knobs on some instruments. Because the circuitry is complex and varies among the different weapons systems, no specific mode will be diagramed. Stability Augmentation Mode The stability augmentation (STAB AUG) mode provides improved control of the aircraft by automatically damping oscillations about the pitch, roll, and yaw axes. Signals from rate gyroscopes command control surface movement through electrohydraulic actuators. In some high-speed aircraft, STAB AUG is considered critical to safe flight. For this reason, the STAB AUG engagement switch connects in series with all other modes of the AFCS. This arrangement ensures that STAB AUG is engaged before any other mode of the AFCS. Attitude Hold The attitude hold mode is the basic, hands-off mode of operation. With attitude hold mode engaged, the AFCS maintains aircraft attitude at the time of engagement in pitch and roll. This particular mode is governed by the actual degrees of bank or pitch of the aircraft. For example, a typical attitude hold mode will release when the aircraft exceeds ±60 degrees in pitch or ±70 degrees in roll. Altitude Hold With altitude hold engaged, the aircraft will maintain the altitude at the time of engagement. If the aircraft is climbing or diving at engagement, the aircraft returns to the altitude that existed at engagement. The AFCS receives control signals for this mode from the ADC. Heading Hold With heading hold engaged, the AFCS maintains aircraft heading at the time of engagement. If the pilot is flying a heading of 180 degrees and engages the heading hold, the AFCS maintains the heading of 180 degrees. Control Stick Steering/Control Wheel Steering The control stick steering/control wheel steering mode lets the pilot manually (moving the stick/wheel) change the attitude of the aircraft with the AFCS engaged without disengaging it. After achieving the new attitude, the pilot releases the stick/wheel, and the AFCS resumes control of the aircraft. Heading Select In the heading select mode, the aircraft will automatically turn to a course selected by the pilot. Upon engagement, the aircraft will assume a fixed maximum roll and turn to the selected heading. Mach Hold The Mach hold function maintains the Mach number existing at the time of Mach hold engagement. In this mode, the ADC commands a pitch-up or pitch-down when airspeed is above or below the selected mach number. 8-32

Automatic Carrier Landing System (ACLS) In the Automatic Carrier Landing System (ACLS) mode, the pilot can make a “hands-off” carrier landing. The aircraft follows command signals generated by the data link receiver. Ground Control Bombing Similar to ACLS, the aircraft follows command signals from personnel on the ground. HELICOPTER AFCS In many respects, the helicopter differs radically from conventional fixed-wing aircraft. However, rotary-wing aerodynamics are very similar to fixed-wing aerodynamics. A review of Airman, NAVEDTRA 14014, will help you understand the material in this discussion. The AFCS is an electrohydromechanical system. It provides inputs to the flight control system to aid the pilot in maneuvering and handling the helicopter. The AFCS consists of three major subsystems—the Stability Augmentation System (SAS), the Stabilator System, and the Digital Automatic Flight Control System (DAFCS). All engagement controls for the three subsystems are on the AFCS and stabilator control panels. Each subsystem operates independently of the other two subsystems, and they complement one another. The pilot engages autopilot functions by pushing the AUTO PLT push button on the AFCS CONTROL panel. The AFCS system provides the following features:  Pitch, roll, and yaw stability augmentation  Stabilator control  Cyclic, collective, and pedal trim  Pitch and roll attitude hold  Airspeed hold  Heading hold  Barometric altitude hold  Radar altitude hold  Pitch and roll hover augmentation/gust alleviation  Turn coordination  Maneuvering stability  Automatic approach to hover  Hover coupler  Automatic depart  Crew hover  Longitudinal stick gradient augmentation (pitch bias actuator)  Blade-fold assist  Automatic preflight check 8-33

Figure 8-30 — Helicopter automatic flight control system panel. Figure 8-31 — Stabilator control panel.  Diagnostics (mode failure display) AFCS Control Panels The pilot controls the AFCS from the AFCS CONTROL panel and the stabilator control panel (Figure 8-30 and Figure 8- 31). The stabilator control panel contains all the operating controls for the stabilator. All the other AFCS controls are on the AFCS CONTROL panel. All detectable AFCS mode failures, except the stabilator, illuminate the AFCS DEGRADED light on the caution/advisory panel. They will also illuminate the appropriate mode failure capsule on the failure advisory section of the AFCS CONTROL panel. Stabilator failures illuminate the STABILATOR caution light on the caution/advisory panel and generate an aural warning tone in the pilot’s and copilot’s headsets. The AFCS CONTROL panel has switches to electronically engage the SAS 1, SAS 2, and SAS/BOOST HYD switch. The SAS/BOOST HYD switch controls pressure application to SAS actuators and pitch boost servos.

8-34

AFCS Failure Monitoring During AFCS failures, a trip of the fault circuit depends on airspeed, which disengages automatic mode when a fault occurs. As a result, the driver supply voltage to both amplifiers is removed. This deenergizes the auto engage relays in the stabilator control panel and turns off the AUTO CONTROL-PUSH TO RESET lighted pushbutton switch ON legend. The stabilator control panel applies an automatic mode disengagement signal to the primary and backup computers to turn on the STABILATOR legend on the mission displays. The automatic mode disengagement signal also energizes the audible warning signal unit that generates a beeping warning tone in the pilot’s and copilot’s headsets. The MASTER CAUTION PRESS TO RESET capsules on the pilot’s and copilot’s master warning panels will also go on when automatic mode is disengaged. Pressing either MASTER CAUTION PRESS TO RESET capsule will cause the pilot’s and copilot’s MASTER CAUTION PRESS TO RESET capsules and the audio warning tone to go off. The STABILATOR caution legend will remain on. Reengagement of the automatic mode can be attempted by pressing the AUTO CONTROL-PUSH TO RESET pushbutton switch. This applies an AUTO CONTROL RESET signal to each stabilator amplifier to initiate automatic mode engagement. If the fault has not been corrected, automatic mode will not engage. The Advanced Flight Control Computer (AFCC) continuously monitors AFCS operation by comparing sensor input signals, by testing program functions, by checking output signals, and by checking servo response to output signals. Should a malfunction be detected, the AFCC automatically disables any function affected. The AFCC also provides fail advisory output signals to the FAIL ADVISORY legends on the AFCS control panel and to the primary and backup computers. The AFCC checks SAS servo valve flapper valve coils by comparing the return current flow from the coil to the output drive current. Any difference will indicate an open or shorted valve solenoid. The AFCC will turn on the affected SAS 1 or SAS 2 FAIL ADVISORY legend on the AFCS control panel and in the case of SAS 2, disable the affected SAS 2 channel (pitch, roll, or yaw). The pitch, roll, yaw, and collective trim servos each contain a position sensor that indicates trim position. If the sensor feedback to the AFCC does not follow the trim drive voltage, a malfunction is indicated. The AFCC will disable the affected channel and turn on the TRIM FAIL ADVISORY legend on the AFCS control panel. Each FAIL ADVISORY legend will flash to indicate a malfunction. This, in turn, will flash the AFCS DEGRADED legend on the mission display panel and the MASTER CAUTION PRESS TO RESET capsule on the master warning panel. By pressing the AKNL ADVSY pushbutton on the AFCS control panel, the flashing FAIL ADVISORY legend will stop flashing, and go on steady. This allows a reset of the MASTER CAUTION PRESS TO RESET capsule by pressing the MASTER CAUTION PRESS TO RESET capsule. Pressing any of the MODE RESET switches resets the malfunctioned AFCC mode. If reset has occurred, the affected FAIL ADVISORY legend on the AFCS control panel and the AFCS DEGRADED legend on the Caution Advisory Panel will go off. Table 8-1 provides a listing of AFCS control panel FAIL ADVISORY legends and the conditions they indicate.

8-35

Table 8-1 — AFCS control panel fail advisory legends The AFCC monitors the AFCS while in flight. Malfunctions detected by the AFCC are stored in nonvolatile memory Flight Bite Code (FBC). Predefined malfunctions have preassigned codes to identify that particular problem. FBCs can be stored. If there are more than five problems, the first four and the last occurring FBC is stored. Upon landing, the FBCs are read out one at a time on the AFCC display by actuating the CODE ADV switch. FBCs can be cycled through and read out any number of times. FBCs can be cleared from memory with the CLR BITE switch. The AFCC also has the capability of performing an AFCS test on the ground. A system test is initiated by setting the GND INIT/CODE ADV switch to GND INIT. The AFCC then proceeds to check sensors, exercise servos and actuators, and check internal functions. Malfunctions are recorded as a 3-digit numeric Ground Bite Code (GBC) and an 8-character alpha series of messages. Predefined malfunctions have preassigned c odes to identify that particular problem. GBCs are cycled through by actuating the CODE ADV switch and can only be read out one time. Removing power from the AFCC will erase the GBCs. Legend Condition A/S Airspeed hold is lost. ALT Barometric or radar altitude hold is lost or degraded, dependent upon which is selected. ATT Pitch or roll autopilot attitude hold is lost. If TRIM legend is also on, both pitch and roll attitudes have been lost or are malfunctioning. If pitch attitude failure occurs, the A/S legend will also go on, indicating airspeed hold is also not available. AUG Hover augmentation is lost. A/S, ATT, or SAS 1 and SAS 2 legends may also be on. BIAS Provisional. CH Loss of crew hover mode. CORD Turn coordination is lost. AUG, SAS 1 and SAS 2, or A/S legend may also be on. CPLR Approach/hover coupler capability is lost. ATT, A/S, or ALT legend may also be on. HDG Heading hold is lost or degraded. SAS 1 A malfunction has occurred, causing improper pitch, roll, or yaw SAS 1 operation. SAS 2 SAS 2 pitch and/or roll failure has occurred. Affected axis is automatically disabled. TRIM Cyclic stick pitch or roll trim, pedal yaw trim, or collective stick trim is lost or malfunctioning. 8-36

Stability Augmentation System (SAS) The SAS 1 system provides electrical control signals proportional to sensor inputs to the pitch, roll, and yaw servo valves. The servo valves convert electrical control signals into hydraulic commands for the SAS actuators. The SAS actuators respond to the hydraulic commands and produce mechanical movement of the flight control linkages without moving the cyclic sticks and pedals. The flight control linkages direct changes in main rotor and tail rotor pitch. With both SAS 1 and SAS 2 engaged, the SAS actuators each have 10 percent authority of flight control movement. Each SAS system has 5 percent authority. With only SAS 1 engaged, the gain of the SAS amplifier is double, but the control authority remains at 5 percent. The AFCS control panel provides a signal to the SAS amplifier to indicate SAS 2 engagement. The SAS 1 pitch channel provides dynamic stability for the helicopter’s pitch axis. The No. 1 pitch rate gyro senses changes in pitch rate and applies this rate to the No. 1 stabilator amplifier. The No. 1 stabilator amplifier filters this pitch rate signal and applies it to the SAS amplifier. The SAS amplifier processes the pitch rate signal to remove long-term rate signals and applies a correction signal to the pitch SAS servo valve. The servo valve controls hydraulic pressure to the SAS actuator. The SAS actuator provides mechanical movement of the flight controls, producing rotor head movement opposing the sensed pitch rate. The SAS 1 roll channel provides dynamic stability and limited roll attitude retention for the helicopter roll axis. A roll attitude signal (H-60S model) is provided to the SAS amplifier from the No. 1 Embedded Global Positioning System /Inertial Navigation System, also known as EGI, by the roll attitude signal path, which is provided by an internal roll rate gyro and provides sensed roll rate. In the H-60H model, the roll attitude signal is provided to the SAS amplifier from the copilot’s Attitude Heading Reference System (AHRS) by the roll attitude signal path. The SAS amplifier receives a roll input signal representing any degree of roll attitude, but limits the signal to represent only up to ±2.1 degrees of roll attitude. The roll attitude signal is summed with a resultant rate + lag rate signal from the internal roll rate gyro. This resultant correction signal is applied to the roll SAS servo valve by the SAS 1 roll valve correction signal path. The servo valve controls the hydraulic pressure applied to the SAS actuator. The SAS actuator provides mechanical movement of the flight controls, producing rotor head movements that oppose the rate gyro and EGI inputs for H-60S models and oppose the rate and displacement gyro inputs for H-60H models. The opposing flight control movements result in the reduction of the sensed roll rate and roll attitude signals to provide rate damping for the helicopter to be restored to a level roll attitude. The SAS 1 yaw channel provides dynamic stability for the helicopter’s yaw axis and turn coordination at airspeeds of 50 knots or greater. At airspeeds of less than 50 knots (for H-60H models) and less than 60 knots (for H-60S models), the No. 1 stabilator amplifier provides a +12 to +15 vdc airspeed switch discrete. The discrete inhibits No. 1 lateral acceleration and roll rate signals in the SAS amplifier. The discrete controls the internal yaw rate gyro in the SAS amplifier as the yaw channel sensor. When sensing a yaw rate, the SAS amplifier processes it to remove long-term yaw rate signals. It then develops a short-term correction signal and applies it to the yaw SAS servo valve. The servo valve controls the actuator, which moves the flight controls to oppose the yaw rate. At speeds below 50 knots (for H-60H models) and below 60 knots (for H-60S models), the airspeed transducer is applied to the No. 1 stabilator amplifier and sent to the SAS 1 8-37

amplifier (discrete signal). This discrete signal allows the yaw SAS amplifier and flight controls to use the internal yaw rate gyro. At speeds above 50 knots (for H-60H models) and above 60 knots (for H-60S models), the SAS amplifier receives a –12 to –15 VDC airspeed switch discrete from the No. 1 stabilator amplifier. This enables filtered No. 1 lateral acceleration and roll rate to sum with the yaw rate signal. Airspeed discrete from No. 1 stabilator amplifier is inhibited for this situation above 50 knots (for H-60H models) and above 60 knots (for H-60S models). Airspeed discrete is not removed from the SAS amplifier. The AFCS consists of many components. To understand the entire system, you need to know what each component does. The following discussion describes the components that make up the AFCS. AFCS POWER SWITCHING ASSEMBLY—The AFCS power switching assembly provides power switching for the DAFCScomputer and the SAS and BOOST shutoff valves. NO. 1 LATERAL ACCELEROMETER— The accelerometer provides lateral acceleration to the SAS amplifier via the No. 1 stabilator amplifier for the turn coordination function of SAS 1. NO. 1 PITCH RATE GYRO—The gyro provides pitch rate to the SAS amplifier via the No. 1 stabilator amplifier for the pitch dynamic stability function. The pitch rate gyros are also part of the stabilator control system. SAS ACTUATORS—The SAS actuators, on the pilot-assist servos, convert electrical signals from the analog SAS 1 and digital SAS 2 to mechanical motion to move the flight controls. The actuators are electrohydraulic, operating from pressure supplied by the No. 2 transfer module, applied through the SAS shutoff valve. When the SAS/BOOST HYD switch is pressed, power is removed from the SAS shutoff valve, opening the valve. Opening the valve allows the actuators to move with signals applied from SAS 1, SAS 2, or both. With pressure removed from the actuators, a spring-loaded device locks the actuator piston in center position. Each actuator output piston connects to linkage. When the actuator piston becomes locked, it forms a fixed pivot point on one end of the linkage. This allows the linkage to move with stick, pedal, or trim inputs. SAS AMPLIFIER— The SAS amplifier contains the No. 1 roll and yaw rate gyro power supplies and processing circuitry for SAS 1. The outputs of the internal roll and yaw rate gyros are used by SAS 1 and the DAFCS computer. The amplifier processes rate and proportional signals. Amplifier outputs operate the SAS actuators on the pilot-assist servo assembly. NO. 1 STABILATOR AMPLIFIER— The No. 1 stabilator amplifier provides filtered lateral acceleration and pitch rate signals and an airspeed discrete to the SAS amplifier. The stabilator amplifier is part of the stabilator control system. Stabilator System The stabilator system optimizes trim attitudes for cruise, climb, and autorotation. Also, it provides pitch stability augmentation to complement the SAS system for additional redundancy. The stabilator system is completely independent of the other two AFCS subsystems except for common airspeed sensors, lateral accelerometers, and pitch rate gyros. The stabilator control system is a completely automatic fly-by-wire control system with a manual backup slew control. The primary purpose of the stabilator control system is to stop undesirable noseup attitudes. The noseup attitude is caused by rotor downwash impinging on the horizontal stabilator during low-s peed flight and transition to a hover. 8-38

The stabilator panel contains an Automatic Control (AUTO CONTROL) switch, a TEST pushbutton, and a Manual Slew (MAN SLEW) switch. The AUTO CONTROL switch is used to engage the automatic mode or to reset the stabilator if it should fail. The pilot can manually position the stabilator to any position within the stabilator limits by moving the MAN SLEW switch. The TEST pushbutton is used to check the automatic mode fault detector. Two electric jackscrews, working in series, position the stabilator. Each actuator provides one-half the input to position the stabilator and is controlled by a separate and redundant stabilator amplifier. The stabilator travels from 40 degrees trailing edge down for hover and low-speed flight below 30 knots to 8 degrees trailing edge up for cruise and maneuvering flight. Four inputs are required to position the stabilator—airspeed, collective stick position, lateral acceleration, and pitch rate. Each stabilator amplifier receives these four inputs, but they receive the inputs from independent sensors. The DAFCS computer monitors each of these sensors for malfunctions, and the stabilator control system monitors and compares the position of the two actuators. Any system malfunction caused by a difference between the two stabilator actuator positions results in the stabilator remaining in the last position. A malfunction also causes an automatic power shutdown to both actuators, an aural tone to the pilot, and a STABILATOR caution light illuminating. The shutdown threshold between the two stabilator actuator positions is 10 degrees for airspeeds less than 50 Knots-Indicated Air Speed (KIAS) and 4 degrees for airspeeds greater than 120 KIAS, with a linear variation between 50 KIAS and 120 KIAS. The airspeed input aligns the stabilator with the main rotor downwash during slow- speed flight. The collective stick position input decouples aircraft pitch attitude from collective position. Pitch rate and lateral acceleration inputs improve the dynamic response of the aircraft, especially in gusty air conditions. The pitch rate input supplements the dynamic stability provided by the SAS and DAFCS. The lateral accelerometer input decouples the aircraft pitch response from changes in tail rotor lift caused by changes in airflow on the canted tail induced with sideslip. If a malfunction of the stabilator system occurs, the pilot can manually position the stabilator with the manual slew switch. The manual slew switch bypasses the stabilator amplifier automatic mode, applying power directly to the actuators through relays in the amplifiers. A stabilator position indicator aids the pilot in positioning the stabilator to any position between the stabilator travel limits. However, total travel is restricted if the malfunction is an actuator failure. Stabilator travel is restricted to 35 degrees if an actuator fails in the full-down position and 30 degrees if an actuator fails in the full-up position. The stabilator control rate is limited to ±6 degrees per second. The following is a description of the components of the stabilator system. To understand the system, you will need to know what these components do. AIRSPEED TRANSDUCER— The stabilator position program is a function of four sensor inputs. Each function has dual sensors for fail-safe operation. Airspeed for the No. 1 stabilator system is sensed by the airspeed transducer. It connects into the ATO pitot-static system and produces a dc output voltage that is proportional to airspeed. AIR DATA TRANSDUCER— An air data transducer accomplishes airspeed sensing in the No. 2 stabilator system. The air data transducer connects into the pilot’s pitot-static system. It produces dc output voltages proportional to airspeed, altitude, and altitude rate. Airspeed data is used by the stabilator system and AFCS computer. Altitude and altitude rate are applied to the DAFCS computer only for altitude hold and depart modes of operation. 8-39

LATERAL ACCELEROMETERS—There are two lateral accelerometers—No. 1 and No 2. They each produce a dc output signal proportional to helicopter lateral acceleration. Each accelerometer dc signal goes to its related stabilator amplifier, where it is conditioned. The filtered lateral acceleration signal is used to position the stabilator to counteract tail rotor downwash on its upper wing surface. No. 1 and No. 2 filtered lateral acceleration is used by the DAFCS computer for signal quality comparison and software generation. No. 1 filtered lateral acceleration is also used in analog SAS for slip or skid correction above 50 knots. STABILATOR POSITION INDICATOR— Stabilator position is displayed by an indicator on the center of the crew station instrument panel. This indicator is a synchro-type device driven by a synchro transmitter mounted in the stabilator position transmitter and limit switch assembly in the tail pylon. STABILATOR CONTROL PANEL— Control functions for the system are provided by the stabilator control panel. The panel consists of an AUTO CONTROL PUSH TO RESET pushbutton switch, a TEST pushbutton, and a MAN SLEW UP/DOWN switch. Engagement of the system is automatic upon application of helicopter ac and dc power, provided all interlocks are in their proper condition. TEST Pushbutton—A TEST pushbutton, operational below 50 knots, provides a check of the system fault monitors by inserting an airspeed-derived test signal into the No. 1 system. This signal drives only the No. 1 stabilator actuator, which produces a difference between the two actuators. The fault monitor circuit in either the No. 1 or No. 2 amplifier, or both, should disengage the automatic mode of operation when the programmed threshold trips. MAN SLEW UP/DOWN Switch— If the automatic mode disengages, and cannot be reset due to a malfunction, the MAN SLEW switch is used to manually position the stabilator. Relays in the No. 1 and No. 2 stabilator amplifiers are operated by dc power from the switch when it is placed to UP or DOWN. Using the switch when the automatic mode is engaged will disengage the automatic mode. As a result, the STABILATOR caution light will go on, and a beeping tone will be heard in the Intercommunication System (ICS). COLLECTIVE STICK POSITION SENSORS— Collective stick position also affects the stabilator position schedule. Stick position is sensed by two collective stick position sensors. No. 1 and No. 2 collective stick position sensors each produce a dc output signal proportional to the collective stick position. Both signals are used by the DAFCS computer for signal level comparison and for software generation of collective-to-yaw pedal coupling. STABILATOR AMPLIFIER— Processing of airspeed, lateral acceleration, collective stick position, and pitch rate is accomplished within each stabilator amplifier. The amplifiers contain a power supply, processing and feedback circuits, and a fault monitor circuit. Sensor inputs are processed, summed, and applied to a motor driver circuit. The motor driver circuit output is applied, through contacts of relays, to the respective stabilator actuators. Any difference of actuator position is sensed by the fault monitor circuit in either or both amplifiers, causing an automatic mode disengagement. ACTUATORS— Two actuators position the stabilator. Each actuator contains an electric motor (geared to a jackscrew), limit switches, and a feedback potentiometer. The potentiometer provides actuator position feedback to each amplifier. The actuators extend or retract, as necessary, to position the stabilator. 8-40

PITCH RATE GYROS—There are two pitch rate gyros—No. 1 and No. 2. Each rate gyro produces a dc output signal relative to the pitch rate of the aircraft. Each rate gyro signal goes to its respective stabilator amplifier, where it is conditioned. The stabilator system uses the filtered pitch rate signal to enhance the AFCS system’s ability to correct short-term pitch disturbances. The DAFCS computer uses No. 1 and No. 2 filtered pitch rate signals for signal quality comparison and software generation. The No. 1 filtered pitch rate signal is also used in analog SAS for short-term pitch correction of the rotor head. Digital Automatic Flight Control System (DAFCS) The central component of the DAFCS is the digital computer. The computer commands the Pitch Bias Actuator (PBA), the inner-loop SAS actuators, and the outer-loop trim actuators in all four control channels. The computer also provides self-monitoring, fault isolation, and failure advisory. The DAFCS uses two types of control— identified as inner loop and outer loop. The inner loop (SAS) uses rate damping to improve helicopter stability. This system is fast in response, is limited in authority, and operates without causing movement of the flight controls. The outer loop (AUTO PILOT) provides long-term inputs by trimming the flight controls to the position required to maintain the selected flight regime. It can drive the flight controls throughout their full range of travel (100-percent authority). The outer-loop drive rate is limited to 10 percent per second. Both inner and outer loops allow for complete pilot override through the normal use of the flight controls. The DAFCS computer processes incoming information from various sensors (Figure 8- 32) aboard the aircraft and stores this information in its memory. The Central Processing Unit (CPU) uses the sensor information to compute required correction signals. Inner-loop correction signals go to the SAS actuators, and outer-loop signals operate trim servos and actuators.

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Figure 8-32 — DAFCS input/output block diagram. TRIM SYSTEM—The parallel trim actuator assemblies provide the flight control force gradients, detent positions and the outer-loop autopilot control functions. The trim actuators command full control authority in all four control channels, but are rate-limited to 10 percent per second. Pressing the trim release switch (cyclic trim release, collective trim release, or pedal release) disengages the respective trim function and allows free control motion. Releasing the trim release switch reengages trim. For yaw trim release above 50 knots, the pilot must press the pedal microswitches and the cyclic trim switch. Below 50 knots, only the pedal microswitches have to be pressed. The pilot can override the trim control forces in all channels. AUTOPILOT— The autopilot maintains helicopter pitch and roll attitude, airspeed, and heading during cruise flight and provides a coordinated turn feature at airspeeds above 50 knots. To engage the autopilot function, the pilot presses the control panel SAS 1 or 8-42

SAS 2 switches, the TRIM switch, and then the AUTO PLT pushbutton. The autopilot may be disengaged by pressing the AUTO PLT pushbutton or pressing the AFCS release button. The computer also provides command signals to the trim actuators to reposition the flight controls using the trim system. ATTITUDE AND AIRSPEED HOLD— Attitude and airspeed holds are engaged with AUTO PLT. In the pitch channel, at airspeeds of less than 50 knots, attitude changes are commanded by changing the cyclic stick position. The pilot can use the TRIM REL switch or the four-direction (beeper) TRIM switch to change cyclic stick position. This causes the cyclic stick to move and the helicopter attitude to change about 5 degrees per second. When cyclic movement stops, the autopilot stabilizes the helicopter around the new stick position and attitude. At speeds above 50 knots and in bank angles less than 30 degrees, the system becomes airspeed sensitive in pitch. Operating the four- direction TRIM switch causes the cyclic stick to move and the helicopter to change airspeed reference at 6 knots per second. Because of variations in the pitot-static system during gusty conditions, integrated longitudinal acceleration is used for short- term correction. The airspeed sensor is used for long-term updates through a 3-second filter. The roll channel autopilot holds roll attitude of the helicopter. Attitude information is supplied to the computer from the pilot’s and copilot’s A/A24G vertical gyros. The command signal is applied to roll SAS 1 and SAS 2 and the roll trim system. When the pilot actuates the four-direction TRIM switch, the helicopter roll attitude will change at about 6 degrees per second. In addition to the attitude hold feature, the system includes an automatic wing-leveling capability. During transitions from hover to airspeeds above 50 knots, this feature automatically retrims the aircraft from a left roll attitude in a hover to a wings-level attitude at 50 knots. After establishing a level attitude, the attitude hold feature maintains that attitude until a new roll attitude is commanded by the pilot. HEADING HOLD— The yaw channel of the autopilot provides the heading hold feature for hover and forward flight. It is engaged whenever the AUTO PLT PBS is illuminated. Heading hold is an outer-loop function, operating through the yaw trim actuator; therefore, it will work only when the yaw trim is engaged. Releasing all pedal switches at a given heading synchronizes the trim system to the established heading. A potentiometer in the yaw trim actuator applies a trim position feedback signal to the computer. This signal cancels the drive signal at the desired position, stopping the motor. The yaw autopilot also uses a collective stick position sensor to hold reference heading for yaw excursions caused by main rotor torque changes. The collective stick position sensor is controlled by an airspeed signal that reduces its gain as airspeed increases. When the heading hold is engaged, the HDG TRIM (slew) switch on the collective lets the pilot make heading changes without retrimming. Below 50 KIAS, the aircraft slews at 3 degrees per second. Above 50 KIAS, actuation of the switch for less than 1 second provides a 1-degree heading change. Actuation for greater than 1 second provides a 1 degree per second coordinated turn. The heading hold is reengaged following a turn when the following conditions are maintained for 2 seconds:  Aircraft roll attitude is within 2 degrees of wings level.  Yaw rate is less than 2 degrees per second.  The heading hold is disengaged by the weight- on-wheels switch when the aircraft is on the ground. 8-43

ALTITUDE HOLD—Either barometric or radar altitude hold is selectable from the AFCS CONTROL panel. With the altitude hold mode on, the DAFCS computer uses a reference altitude. The reference altitude comes from either the air data transducer or the radar altimeter, depending on whether barometric altitude hold or radar altitude hold is selected. The DAFCS computer uses altitude and rate from the barometric or radar altitude systems (depending on which hold mode is selected) and vertical acceleration to command the collective SAS and trim actuators. The computer also monitors engine torque to prevent dual-engine torque from exceeding 116 percent whenever the collective trim is positioning the collective. Barometric altitude hold is engaged at any altitude and airspeed by depressing the BAR ALT PBS with SAS 2 and autopilot engaged. Depressing the collective trim release button temporarily disengages the mode. Upon release of the trim switch, barometer altitude hold automatically reengages and maintains the altitude at the time of reengagement. Radar altitude hold is engaged at any altitude from 0 to 5,000 feet AGL and at any airspeed by depressing the RDR ALT PBS with SAS 2 and autopilot engaged. When in the hover coupler mode, altitude hold is referenced to the altitude selected on the AFCS CONTROL panel HVR ALT potentiometer. Depressing the collective trim release temporarily disengages the mode. Upon release of the trim switch, radar altitude hold automatically reengages to the altitude selected on the AFCS CONTROL panel HVR ALT potentiometer. When in the hover coupler mode, transition from one altitude to another is made with the HVR ALT knob on the AFCS CONTROL panel. If the radar altitude mode should fail while engaged, the barometric altitude hold automatically engages. Integrated vertical acceleration provides short-term radar altitude corrections, and rate information from the radar altimeter altitude signal provides long-term updates. HOVER AUGMENTATION/GUST ALLEVIATION— An additional feature of the SAS, provided only through SAS 2, is hover augmentation/gust alleviation. It further improves aircraft stability at low airspeed using attitude retention and longitudinal and lateral acceleration to eliminate drift. TURN COORDINATION—Automatic turn coordination is provided at airspeeds greater than 50 knots. Turn coordination lets the pilot fly a coordinated turn with directional control provided by the AFCS. The AFCS uses lateral acceleration and roll rate to determine if the aircraft is out of balanced flight. It also provides the yaw SAS and yaw trim with the inputs necessary to maintain an automatic coordinated turn. Automatic turn coordination is engaged and heading hold disengaged when roll attitude is greater than 1 degree, and any of the following conditions exist:  Lateral cyclic force is greater than 3 percent stick displacement.  Cyclic trim release is pressed.  Roll attitude is beeped beyond 2.5 degrees of bank angle.  Actuation of the collective-mounted heading slew for greater than 1 second provides a 1 percent per second coordinated trim. MANEUVERING STABILITY—Pitch control forces are increased to increase pilot effort required for a given pitch rate at bank angles greater than 30 degrees. The higher pitch control forces help alert the pilot to g-loading during maneuvering flight and are provided through the longitudinal trim actuator. A linear longitudinal stick force gradient is provided by trimming 1 percent forward stick for each 1.5 degrees angle of bank between 30 degrees and 75 degrees. At 75 degrees angle of bank, the longitudinal stick 8-44

force is equivalent to 30 percent of stick displacement. The maneuvering stability feature is engaged whenever the AUTO PLT PBS is illuminated. AUTOMATIC APPROACH TO HOVER—The DAFCS provides the capability to perform an automatic approach to a zero longitudinal and any lateral groundspeed selected on the LAT VEL control knob. Also, the DAFCS can perform an automatic approach to any radar altitude selected on the HVR ALT control knob between 40 feet and 200 feet. If the HVR ALT is set below 40 feet, the approach will be made to 40 feet and then continued to the HVR ALT setting when the mode is switched from APPR to HVR. The automatic approach can start at any airspeed and altitude. The automatic approach is an outer-loop-only function, and it commands the aircraft to decelerate or descend until meeting the approach profile conditions. If the approach mode is selected when aircraft conditions are below the approach profile, the DAFCS commands the aircraft to decelerate. The aircraft will decelerate at 1 knot/second while in the radar altitude hold mode until the approach conditions are met. If the approach mode is selected when the aircraft is above the approach profile, the DAFCS commands the aircraft to descend. The descent occurs at 360 feet/minute when the aircraft is more than 50 feet above the approach profile. It occurs at 120 feet/minute when the aircraft is less than 50 feet above the profile. During these conditions, the DAFCS uses the radar altimeter until the approach profile conditions are met. When the approach profile conditions are met, the aircraft simultaneously decelerates at 1 knot/second and descends at 120 feet/minute. This profile is maintained until the aircraft attains 1 knot of Doppler groundspeed and comes to within 1 foot of the selected radar altitude. If the selected altitude is below 40 feet, the aircraft flies to 40 feet and zero longitudinal groundspeed and then descends to the selected attitude. When groundspeed equals 1 knot or less and the aircraft altitude is within 2 feet of the selected altitude, the hover coupler mode automatically engages. The aircraft then accelerates to the selected longitudinal groundspeed. HOVER COUPLER—The hover coupler provides longitudinal and lateral groundspeed control and stabilization about the selected groundspeed and automatic altitude retention. The longitudinal and lateral groundspeed and the altitude are selectable on the AFCS CONTROL panel. Longitudinal and lateral groundspeed can also be beeped ±10 knots with the cyclic trim switch about the groundspeed selected on the AFCS CONTROL panel. The hover coupler mode can automatically engage at the termination of the automatic approach. Also, the pilot can manually engage it when the aircraft is hovering with less than 5 knots longitudinal groundspeed. To do this, the pilot presses the APPR/HVR button on the AFCS CONTROL panel with SAS 2, TRIM, and AUTO PLT engaged. The hover coupler engages when the longitudinal groundspeed is less than 5 knots if engaged manually. After engagement, the aircraft accelerates to the longitudinal and lateral groundspeeds selected on the AFCS CONTROL panel. Pressing and releasing the cyclic TRIM REL removes cyclic trim switch inputs. This action returns the aircraft to the LONG VEL and LAT VEL settings on the AFCS CONTROL panel. Because of Doppler noise, short-term longitudinal and lateral groundspeed is obtained from integrated longitudinal and lateral inertial acceleration. Long-term correction is obtained from the Doppler sensor using a 7-second filter. AUTOMATIC DEPART—The automatic depart mode provides the capability to perform an automatic departure from a coupled hover or from an automatic approach. This mode can take the aircraft to a cruise airspeed of 100 KIAS and altitude of 500 feet. If the coupled hover or the automatic approach feature is engaged, the pilot engages the automatic depart mode by depressing the hover depart button on the cyclic grip. Depressing the DEPART PBS a second time disengages the automatic depart mode, but it doesn’t automatically reengage the RAD ALT hold. Depressing the hover depart 8-45

button a second time disengages the automatic depart mode and returns aircraft control to the pilot. Upon engagement, the aircraft accelerates at 2 knots/second and climbs at 480 feet/minute. During the departure, the DAFCS computer monitors engine torque to ensure that it does not exceed 116 percent. At 100 KIAS, the airspeed hold automatically engages; at a radar altitude of 500 feet, the radar altitude hold automatically engages. Any alternate cruise airspeed or altitude condition of less than 100 KIAS and 500 feet is available to the pilot. The pilot attains an alternate condition by depressing the cyclic trim release and collective trim release at the desired airspeed and altitude, respectively. If either trim release button is depressed and released, the hold mode (airspeed or altitude) associated with that control axis is engaged. The aircraft continues to follow the depart profile for the other axis until the final cruise condition for that axis is met. The automatic depart mode is an outer-loop function operating through the pitch, roll, and collective trim actuators. As in the automatic approach mode, above 60 KIAS, roll attitude is maintained, and below 60 KIAS, the DAFCS commands roll to eliminate lateral drift. CREW HOVER—The crew hover feature lets the crewmember position the helicopter during hoist and rescue operations. The crewmember controls the aircraft from the crew hover-trim panel. The crew hover controller has a control authority of ±5 knots. This authority is laterally and longitudinally about the reference values selected on the AFCS CONTROL panel plus the speeds beeped from the cyclic trim beep switch. The crew hover feature is activated from the AFCS CONTROL panel by depressing the CREW HVR button. It can be activated only if the hover coupler mode is already engaged. PITCH BIAS ACTUATOR (PBA)— The PBA provides longitudinal cyclic displacement proportional to airspeed. The DAFCS commands the PBA as a function of pitch attitude, pitch rate, and airspeed. The PBA is an electromechanical series actuator with ±15 percent control authority and ±3 percent per second rate limit. The PBA functions automatically upon application of power to the DAFCS computer, and it isn’t selectable on the AFCS control panel. The DAFCS computer monitors the PBA position to confirm correct response to the input commands. If the PBA fails, the DAFCS lights the BIAS advisory light and flashes the AFCS DEGRADED light. Also, the DAFCS commands the PBA to a predetermined position, depending on the type of failure. The PBA is driven by the DAFCS computer as a function of airspeed, pitch rate, and pitch attitude. PBA failure modes are as follows:  Attitude failure —bias actuator centered.  Pitch rate failure —faded out pitch rate component.  Airspeed failure —actuator goes to 120-knot position and attitude, and rate continues to function.  Actuator failure —power removed from actuator. If the system malfunctions, the BIAS Fail Advisory light on the AFCS CONTROL panel will go on, and, in some cases, remove power from the actuator. If the malfunction that caused the shutdown was of an intermittent nature, the actuator operation can be reset by pressing the appropriate MODE RESET button.

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End of Chapter 8 AUTOMATIC FLIGHT CONTROL AND STABILIZATION SYSTEMS Review Questions 8-1. What are the two main purposes of the AFCS?

A. To reduce flight time and maintenance B. To provide command and control C. To ease pilot workload and to provide aircraft stability at all speeds D. To lessen the time required to start control and movement

8-2. What is an airfoil?

A. Any part of an aircraft designed to produce lift B. Only the aircraft’s leading edge of rounded surfaces C. Any device mounted on the leading edge and trailing edge D. Only the Bottom surface of an aircraft

8-3. When the pressure on both sides of the airfoil is nearly equal, what happens to the airfoil?

A. It produces drag B. It remains constant C. It increases aircraft speed. D. It stalls

8-4. On a fixed-wing aircraft, what control surface(s) correct(s) for yaw?

A. Ailerons B. Rudders C. Elevators D. Wings

8-5. As fuel is used, what must the pilot do to lessen lift and prevent a gain in altitude?

A. Reduce angle of attack on the wing B. Control pitch C. Adjust ailerons D. Use the rudder to correct any tendency of the aircraft

8-6. In a helicopter with a constant speed rotor, how does the pilot increase lift?

A. By increasing blade angle of tail rotor B. By changing tail rotor pitch to compensate for increases or decreases C. By changing tail rotor speed D. By increasing the airfoil’s angle of attack (blade pitch)

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8-7. When does the action called coning take place?

A. When the pilot increases tail rotor speed B. When the pilot applies collective to the main rotor C. When there is a decrease in main rotor torque D. When there are changes to tail rotor pitch

8-8. What two types of altitude signals can the AFCS use to maintain a constant altitude?

A. Sensor and amplifier B. Data link and aileron C. Radar and barometric altimeter D. Fire control system and weapons control

8-9. Why is the AFCS synchronized with the flight controls before engaging the AFCS?

A. To prevent sudden and violent maneuvers upon engagement B. To maintain the aircraft at a constant altitude C. To use data-link systems to fly the aircraft during approaches and landings D. To program a pilot-desired maneuver that is within the capability of the system

8-10. What unit is considered the heart of the AFCS?

A. The control panel B. The air navigation computer C. The altitude reference D. The command coupler

8-11. What AFCS component provides signal outputs representing yaw, pitch, and roll rates?

A. Roll computer amplifier B. Pitch computer amplifier C. Heading computer D. Three-axis rate gyro

8-12. What type of heading signal does the AFCS receive from the compass/INS system?

A. Gear train B. Tension C. Reference D. Clutched

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8-13. What are the two modes of operation for electrohydraulic servo actuators?

A. Manual and hydraulics B. Manual and electrical signals C. Electrical and pneumatic D. Synchro and hydro

8-14. What unit(s) of the AFCS send(s) a signal that acts as a followup to flight control movement?

A. Synchro device on the boost package B. Load sensor C. Surface position transmitter D. Boost package

8-15. Name the three major systems that make up the AFCS for helicopters.

A. Longitudinal stick gradient augmentation, blade-fold assist, and stabilator control B. Stability augmentation system, stabilator system, and digital automatic flight control C. Stabilator control, automatic approach to hover, and heading hold D. Maneuvering stability, radar altitude hold, and heading hold

8-16. When only SAS 1 is engaged, what is the percentage of flight control authority?

A. 5 B. 10 C. 12 D. 15

8-17. What is the primary purpose of the stabilator system?

A. To remove long-term yaw rate signals B. To limit roll stability C. To stop undesirable noseup attitudes D. To control the internal yaw rate gyro

8-18. When is the stabilator system TEST pushbutton operational?

A. When airspeed is above 50 knots B. When airspeed is below 50 knots C. When airspeed is 60 knots D. When airspeed is between 60 and 70 knots

8-19. What is the purpose of the DAFCS trim actuators?

A. To provide actuator position feedback B. To improve helicopter stability C. To correct short-term pitch disturbances D. To provide self-monitoring, fault isolation 8-49

8-20. What is the rate limit, in percent per second, of the DAFCS trim actuators?

A. 5 B. 10 C. 15 D. 20

8-21. What is the effective altitude range in feet for the DAFCS radar altitude hold mode?

A. 0 to 5,000 B. 5,000 to 6,000 C. 10,000 to 15,000 D. 20,000 to 25,000

8-22. When the helicopter is less than 50 feet above the approach profile, what is the rate of degrees of descent, in feet per minute?

A. 100 B. 120 C. 360 D. 400

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8-51

APPENDIX I GLOSSARY

ACCELEROMETER —A device that measures the acceleration to which it is subjected and develops a signal proportional to it. AGONIC LINE— An imaginary line on the earth’s surface passing through points where the magnetic declination is 0 degrees; that is, a line of longitude where the compass points to true north. AMBIENT CONDITIONS —Physical conditions of the immediate environment, which may pertain to temperature, humidity, pressure, etc. AMPERE — The basic unit of electrical current. AMPERE-TURN —The magnetizing force produced by a current of 1 ampere flowing through a coil of 1 turn. AMPLIDYNE — A rotary magnetic or dynamoelectric amplifier used in servomechanism and control applications. AMPLIFICATION — (1) The process of increasing the strength (current, power, or voltage) of a signal. (2) The ratio of output magnitude to input magnitude in a device that is intended to produce an output that is an enlarged reproduction of its input. AMPLIFIER —A device used to increase the signal voltage, current, or power, generally composed of solid-state circuitry called a stage. It may contain several stages in order to obtain a desired gain. AMPLITUDE —The maximum instantaneous value of an alternating voltage or current, measured in either the positive or negative direction. ANALOG COMPUTER —A type of computer that provides a continuous solution to a mathematical problem with continuously changing inputs. Inputs and outputs are represented by physical quantities that may be easily generated or controlled. APPARENT DRIFT —The effect of the earth’s rotation on a gyro that causes the spinning axis to appear to make one complete rotation in 1 day. Also called APPARENT PRECESSION or APPARENT ROTATION. APPARENT PRECESSION —See APPARENT DRIFT. APPARENT ROTATION —See APPARENT DRIFT. ARC —A flash caused by an electric current ionizing a gas or vapor. ARMATURE —(1) In a relay, the movable portion of the relay. (2) The windings in which the output voltage is generated in a generator or in which input current creates a magnetic field that interacts with the main field in a motor. ATTENUATOR —A network of resistors used to reduce voltage, current, or power delivered to a load. ATTRACTION —The force that tends to make two objects approach each other. Attraction exists between two unlike magnetic poles (north and south) or between two unlike static charges. AI-1

AUTOTRANSFORMER —A transformer in which the primary and secondary are connected together in one winding. AVB —Avionic bulletin. AVC —Avionic change. AXIS —A straight line, either real or imaginary, passing through a body around which the body revolves. AZIMUTH —Angular measurement in the horizontal plane in a clockwise direction. BATTERY — A device for converting chemical energy into electrical energy, with two or more primary or secondary cells connected together electrically. The term does not apply to a single cell. BATTERY CAPACITY —The amount of energy available from a battery. Battery capacity is expressed in ampere-hours. BIAS —Difference of potential applied to a vacuum tube or transistor to establish a reference operating level. BLOCK DIAGRAM —A diagram in which the major components of an equipment or a system are represented by squares, rectangles, or other geometric figures, and the normal order of progression of a signal or current flow is represented by lines. BOLOMETER —A loading device that undergoes changes in resistance as changes in dissipated power occur. BRIDGE CIRCUIT —The electrical bridge circuit is a term referring to any one of a variety of electric circuit networks, one branch of which (the bridge proper) connects two points of equal potential; therefore, it carries no current when the circuit is properly adjusted or balanced. BRUSH —The conducting material, usually a block of carbon, bearing against the commutator or slip rings through which the current flows in or out. BUS BAR —A primary power distribution point connected to the main power source. CABLE HARNESS —A group of wires or ribbons of wiring used to interconnect electronic systems and subsystems. CAGING (GYRO) — The act of holding a gyro so that it cannot precess or change its attitude with respect to the body containing it. CAPACITOR — Two electrodes or sets of electrodes in the form of plates, separated from each other by an insulating material called the dielectric. CAPACITOR-START MOTOR —A type of single-phase, ac induction motor in which a starting winding and a capacitor are placed in series to start the motor. The values of and R are such that the main-winding and starting winding currents are nearly 90 degrees apart, and the starting torque is produced as in a two-phase motor. CARDIOPULMONARY RESUSCITATION – Procedure designed to restore breathing after cardiac arrest. Includes clearing air passages to lungs and heart massage. CELL — A single unit that transforms chemical energy into electrical energy. Batteries are made up of cells. AI-2

CHOKE COIL —A coil of low ohmic resistance and high impedance to alternating current. CIRCUIT —The complete path of an electric current. CIRCUIT BREAKER —An electromagnetic or thermal device that opens a circuit when the current in the circuit exceeds a predetermined amount. Circuit breakers can be reset. CIRCULAR MlL —An area equal to that of a circle with a diameter of 0.001 inch. It is used for measuring the cross section of wires. COAXIAL CABLE — A transmission line consisting of two conductors concentric with and insulated from each other. COMMUTATION — The act of a commutator in converting generator output from an ac voltage to a dc voltage. COMMUTATOR — The copper segments on the armature of a motor or generator. It is cylindrical in shape and is used to pass power into or from the brushes. This mechanical device reverses armature connections in motors and generators at the proper instant so that current continues to flow in only one direction. In effect, the commutator changes ac to dc. COMPARATOR — A circuit that compares two signals or values, and indicates agreement or variance between them. COMPENSATING WINDINGS —Windings embedded in slots in pole pieces, connected in series with the armature, whose magnetic field opposes the armature field and cancels armature reaction. COMPOUND-WOUND MOTORS AND GENERATORS — Machines that have a series field in addition to a shunt field. Such machines have characteristics of both series- and shunt-wound machines. CONDUCTANCE —The ability of a material to conduct or carry an electric current. It is the reciprocal of the resistance of the material, and is expressed in mhos. CONDUCTIVITY —The ease with which a substance transmits electricity. CONDUCTOR —Any material suitable for carrying electric current. CORE —A magnetic material that affords an easy path for magnetic flux lines in a coil. COUNTER EMF —Counter electromotive force; an EMF induced in a coil or armature that opposes the applied voltage. COUNTING CIRCUIT —A circuit that receives uniform pulses representing units to be counted and produces a voltage in proportion to their frequency. COVALENT BOND —A type of linkage between atoms in which the atoms share valence electrons. CPR — See CARDIOPULMONARY RESUSCITATION. CURRENT —The movement of electrons past a reference point. The passage of electrons through a conductor. Measured in amperes. CURRENT LIMITER —A protective device similar to a fuse, usually used in high amperage circuits. AI-3

CYCLE —One complete positive and one complete negative alternation of a current or voltage. D’ARSONVAL METER MOVEMENT —The permanent-magnet moving-coil movement used in most meters. DEGREES OF FREEDOM (GYRO) — A term applied to gyros to describe the number of variable angles required to specify the position of the rotor spin axis relative to the case. DELTA — (1) A three-phase connection in which windings are connected end to end, forming a closed loop that resembles the Greek letter delta. A separate phase wire is then connected to each of the three junctions. (2) A term that is also used to describe a difference between two quantities or measurements. DEMODULATOR — A circuit used in servo systems to convert an ac signal to a dc signal. The magnitude of the dc output is determined by the magnitude of the ac input signal, and its polarity is determined by whether the ac input signal is in or out of phase with the ac reference voltage. DIELECTRIC — An insulator; a term that refers to the insulating material between the plates of a capacitor. DIFFERENTIAL —A mechanical computing device used to add or subtract two quantities. DIGITAL COMPUTER —A type of computer in which quantities are represented in numerical form. It is generally made to solve complex mathematical problems by use of the fundamental processes of addition, subtraction, multiplication, and division. Its accuracy is limited only by the number of significant figures provided. DIODE — A material of either germanium or silicon that is manufactured to allow current to flow in only one direction. Diodes are used as rectifiers and detectors. DIRECT CURRENT —An electric current that flows in one direction only. DISCRIMINATOR — A dual-input circuit in which the output is dependent on the variation of one input from the other input or from an applied standard. DOPPLER EFFECT —An apparent change in the frequency of a sound wave or electromagnetic wave reaching a receiver when there is relative motion between the source and the receiver. EDDY CURRENT —Induced circulating currents in a conducting material that are caused by a varying magnetic field. EFFICIENCY — The ratio of output power to input power, generally expressed as a percentage. ELECTROLYSIS — A type of corrosion (chemical decomposition) caused by current flow resulting from contact of dissimilar metals. ELECTROLYTE — A solution of a substance that is capable of conducting electricity. An electrolyte may be in the form of either a liquid or a paste. ELECTROMAGNET — A magnet made by passing current through a coil of wire wound on a soft iron core. ELECTROMOTIVE FORCE (EMF) —The force that produces an electric current in a circuit. ELECTRON —A negatively charged particle of matter. AI-4

ELECTRON SHELL —A group of electrons that have a common energy level that forms part of the outer structure (shell) of an atom. ENERGY —The ability or capacity to do work. EQUIVALENT CIRCUIT —A diagrammatic arrangement of component parts representing, in simplified form, the effects of a more complicated circuit to permit easier analysis. ERECTING (A GYRO) —The placing of a gyro into a desired position and the maintaining of that position. ERROR SIGNAL — (1) In servo systems, the signal whose amplitude and polarity or phase are used to correct the alignment between the controlling and the controlled elements. (2) The name given to the electrical output of a control transformer. E-TRANSFORMER —A magnetic device with an E configuration, used as an error detector. FARAD —The unit of capacitance. FEEDBACK —A transfer of energy from the output circuit of a device back to its input. FIELD —The space containing electric or magnetic lines of force. FIELD WINDING —The coil used to provide the magnetizing force in motors and generators. FLUX —(1) In electrical or electromagnetic devices, a general term used to designate collectively all the electric or magnetic lines of force in a region. (2) A solution that removes surface oxides from metals being soldered. FLUX DENSITY —The number of magnetic lines of force passing through a given area. FLUX FIELD —All electric or magnetic lines of force in a given region. FREE ELECTRONS —Electrons that are loosely held; consequently, they tend to move at random among the atoms of the material. FREE GYRO — A gyro so gimbaled that it assumes and maintains any attitude in space. The free gyro has two degrees of freedom; torque cannot be applied to the rotor of a truly free gyro. FREQUENCY —The number of complete cycles per second existing in any form of wave motion, such as the number of cycles per second of an alternating current. FULL-WAVE RECTIFIER CIRCUIT — A circuit that uses both the positive and the negative alternations of an alternating current to produce a direct current. FUSE — A protective device inserted in series with a circuit. It contains a metal that will melt or break when current is increased beyond a specific value for a definite period of time. GAIN —The ratio of the output power, voltage, or current to the input power, voltage, or current, respectively. GALVANOMETER —An instrument used to measure small dc currents. GENERATOR — A machine that converts mechanical energy into electrical energy. GIMBAL —A frame in which the gyro wheel spins, and that allows the gyro wheel to have certain freedom of movement. It permits the gyro motor to incline freely and retain that position when the support is tipped or repositioned. AI-5

GROUND —(1) A metallic connection with the earth to establish ground potential. (2) a common return to a point of zero potential. GROUND POTENTIAL —Zero potential with respect to the ground or earth. GYROSCOPE —A wheel or disk so mounted as to spin rapidly about one axis and be free to move about one or both of the two axes mutually perpendicular to the axis of spin. HALF-WAVE RECTIFIER —A rectifier using only one-half of each cycle to change ac to pulsating dc. HEAT SHUNT — A device (preferably a clip-on type) used to absorb heat and protect heat-sensitive components during soldering. HERO —Hazardous electromagnetic radiation to ordnance. HERTZ —A unit of frequency equal to one cycle per second. HMI —Handbook maintenance instructions. HORSEPOWER —The English unit of power, equal to work done at the rate of 550 foot-pounds per second. Equal to 746 watts of electrical power. HYSTERESIS — A lagging of the magnetic flux in a magnetic material behind the magnetizing force that is producing it. HZ —See HERTZ. IMPEDANCE —The total opposition offered to the flow of an alternating current. It may consist of any combination of resistance, inductive reactance, and capacitive reactance. INDUCED CURRENT —Current caused by the relative motion between a conductor and a magnetic field. INDUCTANCE — The property of a circuit that tends to oppose a change in the existing current. INDUCTION — The act or process of producing voltage by the relative motion of a magnetic field across a conductor. INDUCTION MOTOR —A simple, rugged, ac motor with desirable characteristics. The rotor is energized by transformer action (induction) from the stator. Induction motors are used more than any other type. INDUCTIVE REACTANCE — The opposition to the flow of alternating or pulsating current caused by the inductance of a circuit. It is measured in ohms. INERTIA — The physical tendency of a body in motion to remain in motion and a body at rest to remain at rest unless acted upon by an outside force (Newton’s first law of motion). lNFINITE —(1) Extending indefinitely, endless. (2) Boundless, having no limits. (3) An incalculable number. IN PHASE —This term is applied to the condition that exists when two waves of the same frequency pass through their maximum and minimum values of like polarity at the same instant. lNSULATION —A material used to prevent the leakage of electricity from a conductor and to provide mechanical spacing or support as protection against accidental contact with the conductor. AI-6

INTEGRATING CIRCUIT —A circuit whose output voltage is proportional to the product of the instantaneous applied input voltages and their durations. INTEGRATOR —A computing device used for summing up an infinite number of minute quantities. INVERSELY — Inverted or reversed in position or relationship. ISOGONIC LINE —An imaginary line drawn through points on the earth’s surface where the magnetic variation is equal. JOULE — Unit of energy or work. A joule of energy is liberated by 1 ampere flowing for 1 second through a resistance of 1 ohm. JUNCTION — (1) The connection between two or more conductors. (2) The contact between two dissimilar metals or materials, as in a thermocouple. JUNCTION BOX — A box with a cover that serves the purpose of joining different runs of wire or cable and provides space for the connection and branching of the enclosed conductors. KINETIC ENERGY —Energy that a mass possesses by virtue of its motion. KNEE (OF A CURVE) —An abrupt change in direction between two fairly straight segments of a curve. LAG — The amount one wave is behind another in time; expressed in electrical degrees. LAMINATED CORE — A core built up from thin sheets of metal and used in transformers and relays. LEAD — (1) The opposite of LAG. (2) a wire or connection. LEAD-ACID CELL —A cell in an ordinary storage battery in which electrodes are grids of lead containing an active material consisting of certain lead oxides that change in composition during charging and discharging. The electrodes or plates are immersed in an electrolyte of diluted sulfuric acid. LINE OF FORCE — A line in an electric or magnetic field that shows the direction of the force. LOAD — (1) The power that is being delivered by any power-producing device. (2) The equipment that uses the power from the power-producing device. LOGIC CIRCUITS — Digital computer circuits used to store information signals and/or to perform logical operations on those signals. MAGNETIC AMPLIFIER — A saturable reactor-type device that is used in a circuit to amplify or control. MAGNETIC CIRCUIT —The complete path of magnetic lines of force. MAGNETIC FIELD —The space in which a magnetic force exists. MAGNETIC FLUX —The total number of lines of force issuing from a pole of a magnet. MAGNETIZE —To convert a material into a magnet by causing the molecules to rearrange. MAGNETO —A generator that produces alternating current and has a permanent magnet as its field. AI-7

MATTER —Any physical entity that possesses mass. METEOROLOGY AUTOMATED SYSTEM FOR UNIFORM RECALL AND REPORTING (MEASURE) —The Navy data processing system designed to provide a standardized system for the recall, scheduling, and documenting of test equipment into calibration facilities. MEGGER. — A test instrument used to measure insulation resistance and other high resistances. It is a portable hand-operated dc generator used as an ohmmeter. MEGOHM — A million ohms. METER —A device used to measure a specific quantity, such as current, voltage, or frequency. METER MOVEMENT —The part of the meter that moves to indicate some value. METER SHUNT — A resistor placed in parallel with the meter terminals; used to provide increased range capability. MICRO —A prefix meaning one-millionth. MICROMETER —A unit of length equal to 10-6 meter. Formerly a micron. MICRON —See MICROMETER. MIL —The diameter of a conductor equal to 1/1000 (.001) inch. MIL-FOOT —A unit of measurement for conductors (diameter of 1 mil, 1 foot in length). MILITARY SPECIFICATIONS (MILSPEC) —Technical requirements and standards adopted by the Department of Defense (DoD) that must be met by vendors selling materials to the DoD. MILITARY STANDARDS (MILSTD) —Standards of performance for components or equipment that must be met to be acceptable for military systems. MILLI —A prefix meaning one-thousandth. MILLIAMMETER —An ammeter that measures current in thousandths of an ampere. MOTOR —A machine that converts electrical energy to mechanical energy. It is activated by ac or dc voltage, depending on the design. MOTOR-GENERATOR —A motor and a generator with a common shaft used to convert line voltages to other voltages or frequencies. MULTICONDUCTOR —More than one conductor, as in a cable. MULTIMETER — A single meter combining the functions of an ammeter, a voltmeter, and an ohmmeter. MULTIPHASE —See POLYPHASE. MUTUAL INDUCTANCE —A circuit property existing when the relative position of two inductors causes the magnetic lines of force from one to link with the turns of the other. NAMP —The Naval Aviation Maintenance Program. NANOMETER —A unit of length equal to 10-9 meter. Formerly millimicron. NEGATIVE CHARGE —The electrical charge carried by a body that has an excess of electrons. AI-8

NEGATIVE FEEDBACK —Feedback in which the feedback signal is out of phase with the input signal. NEGATIVE TEMPERATURE COEFFICIENT—A characteristic of a semiconductor material, such as silver sulfide, in which resistance to electrical current flow decreases as temperature increases. NEUTRON —A particle having the weight of a proton but carrying no electric charge. It is located in the nucleus of an atom. NOISE — (1) Any undesired disturbance within the useful frequency band. (2) That part of the modulation of a received signal (or an electrical or electronic signal within a circuit) representing an undesirable effect of transient conditions. NUCLEUS —The central part of an atom that is mainly made up of protons and neutrons. It is the part of the atom that has the most mass. NULL — A point or position where a variable strength signal is at its minimum value (or zero). OHM —The unit of electrical resistance. That value of electrical resistance through which a constant potential difference of 1 volt across the resistance will maintain a current flow of 1 ampere through the resistance. OHM’S LAW —The current in an electrical circuit is directly proportional to the electromotive force in the circuit. The most common form of the law is E = IR, where E is the electromotive force or voltage across the circuit, I is the current flowing in the circuit, and R is the resistance of the circuit. OVERLOAD — A load greater than the rated load of an electrical device. PACKAGING —An all-inclusive term covering cleaning, preserving, packaging, packing, and marking required to protect items during every phase of shipment, handling, and storage. PARAMETERS —In electronics, the design or operating characteristics of a circuit or device. PERMALLOY —An alloy of nickel and iron having an abnormally high magnetic permeability. PERMEABILITY — A measure of the ease with which magnetic lines of force can flow through a material as compared to air. PHASE — The angular relationship between two alternating currents or voltages when the voltage or current is plotted as a function of time. When the two are in phase, the angle is zero; both reach their peak simultaneously. When out of phase, one will lead or lag the other; that is, at the instant when one is at its peak. The other phase will not be at peak value and (depending on the phase angle) may differ in polarity as well as magnitude. PHASE DIFFERENCE —The time in electrical degree by which one wave leads or lags another. PHOTON (hv) — An elementary quantity of radiant energy (quantum) whose value is equal to the product of Plank’s constant and the frequency of the electromagnetic radiation. AI-9

PICKOFF —In gyros, a sensing device that measures the angle of the spin axis with respect to its reference and provides an error signal that indicates the direction and (in most cases) the magnitude of the displacement. PLANCK’S CONSTANT —A constant that gives the unvarying ratio of the energy of a quantum of radiation to its frequency and that has an approximate value of 6.626 × 10-34 J·s —Symbol h. POLARITY —The character of having magnetic poles, or electric charges. POLYPHASE —A circuit that uses more than one phase of alternating current. POSITIVE CHARGE —The electrical charge carried by a body that has become deficient in electrons. POSITIVE FEEDBACK —Feedback in which the feedback signal is in phase with the input signal. POSITIVE TEMPERATURE COEFFICIENT— The characteristic of a conductor in which the resistance increases as temperature increases. POTENTIAL — The amount of charge held by a body as compared to another point or body. Usually measured in volts. POTENTIOMETER —A variable voltage divider; a resistor that has a variable contact arm so that any portion of the potential applied between its ends may be selected. POWER —The rate of doing work or the rate of expending energy. The unit of electrical power is the watt. POWER FACTOR — The ratio of the actual power of an alternating or pulsating current, as measured by a wattmeter, to the apparent power, as indicated by ammeter and voltmeter readings. The power factor of an inductor, capacitor, or insulator is an expression of their losses. POWER SUPPLY —A unit that supplies electrical power to another unit. It changes ac to dc and maintains a constant voltage output within limits. PRECESSION —The reaction of a gyro to an applied torque, which causes the gyro to tilt itself at right angles to the direction of the applied torque in such a manner that the direction of spin of the gyro rotor will be in the same direction as the applied torque. PRIMARY WINDING —The winding of a transformer connected to the electrical source. PRIME MOVER —The source of the turning force applied to the rotor of a generator. This may be an electric motor, a gasoline engine, a steam turbine, and so forth. PROTON — A positively charged particle in the nucleus of an atom. RADAR —An acronym for radio detecting and ranging. RADAR ALTIMETER —Airborne radar that measures the distance of the aircraft above the ground. RADIAN — In a circle, the angle included within an arc equal to the radius of the circle. A complete circle contains 2 radians. One radian equals 57.3 degrees, and 1 degree equals 0.01745 radian. RATE GYRO —A gyro with one degree of freedom that has an elastic restraint, with or without a damper, and whose output will be proportional to the rate of the applied torque. AI-10

RATIO —The value obtained by dividing one number by another, indicating their relative proportions. REACTANCE —The opposition offered to the flow of an alternating current by the inductance, capacitance, or both, in any circuit. RECTIFIERS — Devices used to change alternating current to unidirectional current. These may be vacuum tubes; semiconductors, such as germanium and silicon; and dry- disk rectifiers, such as selenium and copper oxide. RELUCTANCE —A measure of the opposition that a material offers to magnetic lines of force. RESISTANCE — The opposition to the flow of current caused by the nature and physical dimensions of a conductor. RETENTIVITY — The measure of the ability of a material to hold its magnetism. RHEOSTAT —A variable resistor. RIGIDITY —In gyros, the characteristics of a spinning body that causes it to oppose all attempts to tilt it away from the axis in which it is spinning. ROTATING FIELD —The magnetic field in a multiphase ac motor that is the result of field windings being energized by out-of-phase currents. In effect, the magnetic field is made to rotate electrically rather than mechanically. ROTOR — (1) The revolving part of a rotating electrical machine. The rotor may be either the field or the armature, depending on the design of the machine. (2) The rotating member of a synchro that consists of one or more coils of wire wound on a laminated core. Depending on the type of synchro, the rotor functions similarly to the primary or secondary winding of a transformer. SATURABLE REACTOR — A control device that uses a small dc current to control a large ac current by controlling core flux density. SATURATION — The condition existing in any circuit when an increase in the driving signal produces no further change in the resultant effect. SCHEMATIC — A diagram that shows, by means of graphic symbols, the electrical connections and functions of a specific circuit arrangement. SECONDARY — The output coil of a transformer. SELF-INDUCTION —The process by which a circuit induces an EMF into itself by its own magnetic field. SERIES CIRCUIT — An arrangement where electrical devices are connected so that the total current must flow through all the devices; electrons have one path to travel from the negative terminal to the positive terminal. SERIES-WOUND —A motor or generator in which the armature is wired in series with the field winding. SERVO — A device used to convert a small movement into one of greater movement or force. SERVOMECHANISM —A closed-loop system that produces a force to position an object according to the information that originates at the input. AI-11

SERVOMOTOR —An ac or dc motor used in servo systems to move a load to a desired position or at a desired speed. The ac motor is usually used to drive light loads at a constant speed, while the dc motor is used to drive heavy loads at varying speeds. SERVO SYSTEM —An automatic feedback control system that compares a required condition (desired value, position, etc.) with an actual condition and uses the difference to drive a control device to achieve the required condition. SHIELDING —(1) A metallic covering used to prevent magnetic or electromagnetic fields from affecting an object. (2) A technique designed to minimize internal and external interference. SHUNT —A resistive device placed in parallel with another component. Appreciable current may flow through it, and an appreciable voltage may exist across it. SLIP RINGS — (1) Contacts that are mounted on the shaft of a motor or generator to which the rotor windings are connected and against which the brushes ride. (2) Devices for making electric connections between stationary and rotating contacts. SOLENOID —An electromagnetic coil that contains a movable plunger. SOLID-STATE DEVICE —An electronic device that operates by the movement of electrons within a solid piece of semiconductor material. SOURCE —(1) The object that produces the waves or disturbance. (2) The name given to the end of a two-wire transmission line that is connected to a source. (3) The device that furnishes the electrical energy used by a load. SPECIFIC GRAVITY —The ratio between the density of a substance and that of pure water at a given temperature. STATOR — (1) The stationary part of a rotating electrical machine. The stator may be either the field or the armature, depending on the design of the machine. (2) The stationary member of a synchro that consists of a cylindrical structure of slotted laminations on which three Y-connected coils are wound with their axes 120 degrees apart. Depending on the type of synchro, the stator’s functions are similar to the primary or secondary windings of a transformer. STRANDED CONDUCTOR —A conductor composed of a group of wires. The wires in a stranded conductor are usually twisted together and not insulated from each other. STRANDS —Fine metallic filaments twisted together to form a single wire. SUBASSEMBLY —Two or more parts that form a portion of an assembly or a unit. SUPPORT EQUIPMENT (SE) —All the equipment on the ground or ship needed to support aircraft in a state of readiness for flight. SYNCHRO — A small motor-like analog device that operates like a variable transformer and is used primarily for the rapid and accurate transmission of data among equipments and stations. SYNCHRO SYSTEM —An electrical system that gives remote indications or control by means of self-synchronizing motors. TACHOMETER — An instrument for indicating revolutions per minute. TEMPERATURE COEFFICIENT —The amount of change of resistance in a material per unit change in temperature. AI-12

TERTIARY WINDING —A third winding on a transformer or magnetic amplifier that is used as a second control winding. THERMISTOR —A resistor that is used to compensate for temperature variations in a circuit. See also BOLOMETER. THERMOCOUPLE —A junction of two dissimilar metals that produces a voltage when heated. TINNING — The process of applying a thin coat of solder to materials prior to their being soldered; for example, application of a light coat of solder to the filaments of a conductor to hold the filaments in place prior to soldering of the conductor. TOROID(AL) — (1) A surface generated by a closed plane curve rotated about a line that lies in the same plane as the curve but does not intersect it. (2) A body whose surface has the form of a toroid. TOROIDAL —Of, relating to, or shaped like a torus or toroid: doughnut-shaped <a toroidal resistance coil>. TORQUE —The turning effort or twist that a shaft sustains when transmitting power. A force tending to cause rotational motion; the product of the force applied times the distance from the force to the axis of rotation. TOTAL RESISTANCE ( ) —The equivalent resistance of an entire circuit. For a series circuit: = + + . . . . For parallel circuits: 1 = 1 + 1 + 1 + … 1 [ 1 ]. RT R1 R2 R3 Rn TRANSFORMER —A device composed of two or more coils, linked by magnetic lines of force, used to transfer energy from one circuit to another. TRANSFORMER EFFICIENCY —The ratio of output power to input power, generally expressed as a percentage: Efficiency = P out x 100. P in TRANSFORMER, STEP-DOWN —A transformer constructed so that the number of turns in the secondary winding is less than the number of turns in the primary winding. This construction will provide less voltage in the secondary circuit than in the primary circuit. TRANSFORMER, STEP-UP — A transformer constructed so that the number of turns in the secondary winding is more than the number of turns in the primary winding. This construction will provide more voltage in the secondary circuit than in the primary circuit. TRUE BEARING —Angle between a target and true north measured clockwise in the horizontal plane. TRUE NORTH —Geographic north. TUMBLE (GYRO) —To subject a gyro to torque so that it presents a precession violent enough to cause the gyro rotor to spin end over end. TURNS RATIO — The ratio of the number of turns in the primary winding to the number of turns in the secondary winding of a transformer. AI-13

VALENCE —The measure of the extent to which an atom is able to combine directly with other atoms. It generally depends on the number and arrangement of the electrons in the outermost shell of the atom. VALENCE SHELL —The electrons that form the outermost shell of an atom. VECTOR —A line used to represent both direction and magnitude. VOLT —The unit of electromotive force or electrical pressure. One volt is the pressure required to send 1 ampere of current through a resistance of 1 ohm. VOLTAGE —(1) The term used to signify electrical pressure. Voltage is a force that causes current to flow through an electrical conductor. (2) The voltage of a circuit is the greatest effective difference of potential between any two conductors of the circuit. VOLTAGE DIVIDER —A series network in which desired portions of the source voltage may be tapped off for use in the circuit. WATT — The unit of electrical power. WHEATSTONE BRIDGE —An ac bridge circuit used to measure unknown values of resistance, inductance, or capacitance. WIRING DIAGRAM —A diagram that shows the connections of an equipment or its component devices or parts. It may cover internal or external connections, or both, and contains such detail as is needed to make or trace connections that are involved. WORK — The product of force and motion. WYE (Y) —A three-phase connection in which one end of each phase winding is connected to a common ground. X-AXIS — In a gyro, the spin axis of the gyro. Y-AXIS —In a gyro, an axis through the center of gravity and perpendicular to the spin axis. Z-AXIS — In a gyro, an axis through the center of gravity and mutually perpendicular to both the X (spin) and Y axes. ZENER DIODE — A p-n junction diode designed to operate in the reverse-bias breakdown region. AI-14

APPENDIX II SYMBOLS

AII-1

AII-2

AII-3

AII-4

AII-5

AII-6

AII-7

AII-8

NOTE Although the following references were current when this NRTC was published, their continued currency cannot be assured. When consulting these references, keep in mind that they may have been revised to reflect new technology or revised methods, practices, or procedures; therefore, you need to ensure that you are studying the latest references. If you find an incorrect or obsolete reference, please use the Rate Training Manual User Update Form provided at the end of each chapter to contact the CNATT Rate Training Manager. APPENDIX III REFERENCES

Chapter 1 Navy Electricity and Electronics Training Series (NEETS), Module 1, Introduction to Matter, Energy, and Direct Current, NAVEDTRA 14173, Naval Education and Training Professional Development and Technology Center (NETPDTC), Pensacola, FL, January 2003. Chapter 2 Engineering Change Proposal ECP-MDC-F/A-18-06313 for AN/APN/-194 RADAR Altimeter Mounting Tray Modification, Navy Model FA-18E/F and EA-18G, Commander, Naval Air Systems Command, Washington, DC, 22 April 2009. Fluke Model 8808A Digital Multimeter User’s Manual, Fluke Corporation, P.O. Box 9090, Everett, WA 98206, USA, December 2009. Fluke Model Series IVDigital Multimeter User’s Manual, Fluke Corporation, P.O. Box 9090, Everett, WA 98206, USA, September 2006. Handbook, Operation, Service and Overhaul Instructions with Illustrated Parts Breakdown, Tachometer Indicator-Generator Test Set Part Number TTU-27/E, NAVWEPS 17-15CM-3, Direction of the Chief of the Bureau of Naval Weapons, 15 April, 1965. NEETS, Module 16, Introduction to Test Equipment, NAVEDTRA 14188, NETPDTC, Pensacola, FL, September 1998. Navy Safety and Occupational Health Program Manual, OPNAVINST 5100.23G, CH-1, Chief of Naval Operations, Washington, DC, 21 July 2011. Navy Training System Plan, A/F 37T-21 Aircraft Engine Test Stand, N88-NTSP-A-50- 0005/A, Commander, Naval Air Systems Command, Washington, DC, 1 October 2002. Organizational Maintenance with Illustrated Parts Breakdown, Operating Instructions Test Set, Air Data TS-4508/U, NSN 4920-01-449-8072, NAVAIR 17-15CA-62, Naval Air Systems Command, Washington, DC, 15 January 1998. AIII-1

Technical Manual, Cleaning and Corrosion Control, NAVAIR 01-1A-509 (series), Naval Air Systems Command, Washington, DC, 1 July 2009. Technical Manual, Installation and Repair Practices, NAVAIR 01-1A-505 (series), Naval Air Systems Command, Washington, DC, 15 September 2009. Technical Manual, Intermediate Maintenance Instructions with Illustrated Parts Breakdown, Test Set, Air Data TS-4508/U, NAVAIR 17-15CA-62.1, Naval Air Systems Command, Washington, DC, 1 December 2010. Technical Manual, Intermediate Maintenance Instructions with Illustrated Parts Breakdown, Propeller Synchronizer Tester TTK-512/E24T-204(v), Part Number 1383AS200-1, AG-240AB-OMP-000, Commander, Naval Air Systems Command, Washington, DC, 1 September 1992. Technical Manual, Maintenance Instructions, Organizational, Flight Instruments, Navy Model P-3C Aircraft, NAVAIR 01-75PAC-2-13.1.3, Commander, Naval Air Systems Command, Washington, DC, 1 June 2011. Technical Manual, Maintenance Instructions, Organizational, Power Plant Related Electrical Systems, Navy Model P-3C Aircraft, NAVAIR 01-75PAC-2-13.1.2, Commander, Naval Air Systems Command, Washington, DC, 1 January 2004. Technical Manual, Maintenance Instructions, Organizational , Utility Systems, Navy Models P-3A, P-3B, and P-3C Aircraft, NAVAIR-01-75PAA-2-2.4, Commander, Naval Air Systems Command, Washington, DC, 15 July 2011. Technical Manual, Operation and Intermediate Maintenance with Illustrated Parts Breakdown, Temperature Control System Test Set, Part Number BR-61-103, AG- 513SA-OMP-000, Naval Air Systems Command, Washington, DC, 1 January 2005. Technical Manual, Operation and Maintenance Instructions with Illustrated Parts Breakdown, Test Set Indicator TTU-378A/E 361-046-001, NAVAIR 17-15BD-46, Naval Air Systems Command, Washington, DC, CH-1, 1 July 1993. Technical Manual, Operation and Service Instructions with Illustrated Parts Breakdown, JetCal® Analyzer, BH112J-46 4920-00-090-3409, NAVAIR 17-15A-503, Commander, Naval Air Systems Command, Washington, DC, 1 December 1991. Technical Manual, Organizational and Intermediate Maintenance with Illustrated Parts Breakdown, Fuel Control Test Set Assembly (TTU-597/E), Part Number BC849A6010- 000, NAVAIR 17-15MT-4, Naval Air Systems Command, Washington, DC, 1 October 2010. Technical Manual, Organizational and Intermediate Maintenance, Wiring Repair with Parts Data General Wiring Repair Procedures, Navy Model F/A-18A and F/A-18B, 161353 and Up, A1-F18AC-WRM-001 (series), Commander, Naval Air Systems Command, Washington, DC, 1 August 2009. Technical Manual, Organizational Maintenance, General Aircraft Information (GAI), Navy Model F/A-18A/B/C/D, 161353 and Up, A1-F18AC-GAI-000, Commander, Naval Air Systems Command, Washington, DC, 1 November 2009. Technical Manual, Organizational Maintenance, Principles of Operation, Electrical System, Navy Model F/A-18A/B/C/D, 161353 and Up, A1-F18AC-420-100, Commander, Naval Air Systems Command, Washington, DC, CH-18, 1 August 2008. Technical Manual, Organizational Maintenance, System Maintenance with Illustrated Parts Breakdown, Electrical System, Navy Model F/A-18A/B/C/D, 161353 and Up, A1- AIII-2

F18AC-420-300, Commander, Naval Air Systems Command, Washington, DC, CH-22, 1 April 2010. Technical Manual, Organizational Maintenance, Wiring Diagrams, Navy Model F/A-18A, 161353 thru 163175, A1-F18AC-WDM-000 (series), Commander, Naval Air Systems Command, Washington, DC, 15 May 2010. Technical Manual, Standard Maintenance Practices, Miniature/Microminiature (2M) Electronic Assembly Repair, Organizational/Intermediate/Depot Level, NAVAIR-01-1A- 23, Revision 3, Commander, Naval Air Systems Command, Washington, DC, 1 October 2006. The Naval Aviation Maintenance Program, COMNAVAIRFORINST 4790.2 (series), Commander Naval Air Forces, Washington, DC, 10 November 2009. Tools and Their Uses, NAVEDTRA 14256, NETPDTC, Pensacola, FL, June 1992. Chapter 3 Handbook, Operation and Service Instructions, Inverter, Type DMZ3508M-PM3508G-X, AN 03-5HE-1, Secretary of the Air Force and Chief of the Bureau of Aeronautics, 1 December 1952. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Electrical System, AC Power. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Electrical System, DC Power. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Electrical System, Ground Power System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Electrical System, Power Distribution System. NEETS, Module 5, Introduction to Generators and Motors, NAVEDTRA 14177, NETPDTC, Pensacola, FL, March 2003. Technical Manual, Maintenance Instructions, Organizational, Electrical System, Navy Model P-3A/B Aircraft, NAVAIR 01-75PAA-2-29, Commander, Naval Air Systems Command, Washington, DC, 15 March 1985. Technical Manual, Operation and Maintenance Instructions with Illustrated Parts Breakdown, Navy and Air Force Aircraft and Aircraft Support Equipment Storage Batteries, NAVAIR 17-15BAD-1, Naval Surface Warfare Center Crane, Indiana, Commander, Naval Air Systems Command, Washington, DC, 15 September 2011. Technical Manual, Organizational Maintenance, Auxiliary Power Unit, Navy Models P- 3A, P-3B, and P-3C Aircraft, NAVAIR 01-75PAA-2-4.4, Commander, Naval Air Systems Command, Washington, DC, 1 May 2009. Technical Manual, Organizational Maintenance, Electrical Power and Lighting Systems, Navy Model EA-6B Aircraft, NAVAIR 01-85ADC-2-12, Commander, Naval Air Systems Command, Washington, DC, 1 March 2010. Technical Manual, Organizational Maintenance System Schematics, Electrical System, Navy Model F/A-18E/F and EA-18G, 165533 and Up, A1-F18EA-420-500, Naval Air Systems Command, Washington, DC, 15 March 2011. AIII-3

Chapter 4 Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, Air Data System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, Environmental Control System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, Interior Lighting System, Exterior Lighting System, Inflight Refueling System. NATOPS Flight Manual, Navy Model F/A-18E/F, 165533 and Up Aircraft, A1-F18EA- NFM-000, Naval Air Systems Command, Washington, DC, 1 February 2011. NATOPS Flight Manual, Navy Model P-3C Aircraft, NAVAIR 01-75PAC-1, Commander, Naval Air Systems Command, Washington, DC, 15 March 2010. Organizational Maintenance, Principles of Operation, Exterior Lighting System, Effectivity: HH-60H Helicopter Serial No. 163783 and Subsequent HH-60J Helicopter Serial No. 163801 and Subsequent, A1-H60HA-420-100, Commander, Naval Air Systems Command, Washington, DC, 30 July 2004. Organizational Maintenance, Principles of Operation, Stability Augmentation System (SAS 1), Effectivity: HH-60H Helicopter Serial No. 163783 and Subsequent HH-60J Helicopter Serial No. 163801 and Subsequent, A1-H60HA-560-100, Commander, Naval Air Systems Command, Washington, DC, 27 February 2009. Technical Manual, Maintenance Instructions, Organizational, Airframe Related Electrical Systems, Navy Models P-3C Aircraft, NAVAIR 01-75PAA-2-13-1-1, Commander, Naval Air Systems Command, Washington, DC, 15 June 2011. Technical Manual, Maintenance Instructions, Organizational, Integrated Flight Station Systems Navy Model P-3C Aircraft, NAVAIR 01-75PAC-2-9, Commander, Naval Air Systems Command, Washington, DC, 15 April 2011. Technical Manual, Organizational Maintenance, Instrument Systems, Navy Model EA- 6B Aircraft, NAVAIR 01-85ADC-2-10, Commander, Naval Air Systems Command, Washington, DC, 1 February 2011. Technical Manual, Organizational Maintenance, System Schematics, Environmental Control Systems, Navy Model F/A-18E/F and EA-18G, 165533 and Up, A1-F18EA-410- 500, Commander, Naval Air Systems Command, Washington, DC, 1 October 2010. Chapter 5 Technical Manual, Organizational Maintenance, Principles of Operation, Fuel Systems, Navy Model F/A-18A and F/A-18B, 161353 and Up, A1-F18AC -460-100, Commander, Naval Air Systems Command, Washington DC, CH-4, 1 July 2009. Technical Manual, Organizational Maintenance, Principles of Operation, Power Plants and Related Systems, A1-F18AC-270-100, Commander, Naval Air Systems Command, Washington, DC, CH-4, 1 February 2004. Technical Manual, Organizational Maintenance, Principles of Operation, Rotor Systems, Blade Fold System, A1-H60CA-150-100, Commander, Naval Air Systems Command, Washington, DC, 30 July 2004. AIII-4

Technical Manual, Maintenance Instructions, Organizational, Utility Systems, Navy Models P-3A, P-3B, and P-3C Aircraft, NAVAIR 01-75PAA-2-2.4, Commander, Naval Air Systems Command, Washington, DC, 15 July 2011. Technical Manual, Organizational Maintenance, Principles of Operation, Secondary Power System, Navy Model F/A-18A/B/C/D, A1-F18AC-240-100, Commander, Naval Air Systems Command, Washington, DC, 1 August 2002. Chapter 6 Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Air Data System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Backup Attitude and Navigation System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Cabin Pressurization System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Hydraulic System. Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Instrument System. NEETS, Module 4, Introduction to Electrical Conductors, Wiring Techniques and Schematic Reading, NAVEDTRA 14176, NETPDTC, Pensacola, FL, September 1998. NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187, NETPDTC, Pensacola, FL, July 2003. Technical Manual, Maintenance Instructions, Organizational, Electrical Wiring Data, Navy Models P-3A and P-3B Aircraft, NAVAIR 01-75PAA-2-13.2, Commander, Naval Air Systems Command, Washington, DC, CH-10, 1 August 2005. Technical Manual, Organizational Maintenance, Aircraft Electromechanical System Theory, Navy Model E2-C Aircraft, NAVAIR 01-E2AAA-2-1.1, Commander, Naval Air Systems Command, Washington, DC, CH-9, 15 July 2008. Technical Manual, Organizational Maintenance, Principles of Operation, Air Data Computer System, Navy Model F/A-18A/B/C/D 161353 and Up, A1-F18AC-560-100, Commander, Naval Air Systems Command, Washington, DC, CH-3, 1 July 2001. Technical Manual, Organizational Maintenance, Principles of Operation, Inertial Navigation and Backup Attitude and Navigation System, Navy Model F/A-18A/B/C/D 161353 and Up, A1-F18AC-730-100, Commander, Naval Air Systems Command, Washington, DC, 15 June 2007. Technical Manual, Organizational Maintenance, Principles of Operation, Instrument Systems, Navy Model F/A-18A/B/C/D 161353 and Up, A1-F18AC -510-100, Commander, Naval Air Systems Command, Washington, DC, CH-4, 1 December 1999. Chapter 7 Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Inertial Navigation System and Accurate Navigation System (ANAV), Global Positioning System (GPS). AIII-5

Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Mission Computer (MC)/Fiber Channel Network (FCN). Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Multipurpose Display Group (MDG), Flight Incident Recorder and Monitoring System (FIRAMS). Interactive Electronic Technical Manual (IETM), A1-F18E/F/G, Buno 166817, Lot 30, OFP H5E, Tactical Moving Map Capability (TAMMAC) System, Display Symbology. NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187, NETPDTC, Pensacola, FL, July 2003. Technical Manual, Organizational Maintenance, Principles of Operation, Navigation Systems, Navy Model CH-53E, MH-53E, A1-H53CE-700-100, Commander, Naval Air Systems Command, Washington, DC, 1 November 2010. Technical Manual, Organizational Maintenance, Testing and Troubleshooting Procedures Flight Reference and Automatic Flight Control System, Effectivity: SH-60B Helicopter Serial No. 161553 and Subsequent, A1-H60BB-560-200, Naval Air Systems Command, Washington, DC, 1 November 2007. Chapter 8 NATOPS Flight Manual, Navy Model SH-60B, NAVAIR A1-H60BB-NFM-000, Naval Air Systems Command, Washington, DC, 15 February 2010. NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA 14187, NETPDTC, Pensacola, FL, July 2003. Technical Manual, Organizational Maintenance, Principles of Operation, Stability Augmentation System (SAS 1), Effectivity: HH -60H Helicopter Serial No. 163783 and Subsequent HH-60J Helicopter Serial No. 163801 and Subsequent, A1-H60HA-560- 100, Naval Air Systems Command, Washington, DC, 27 February 2009. Technical Manual, Organizational Maintenance, Principles of Operation, Stability Augmentation System (SAS 1), Effectivity: MH-60S Helicopter Serial No. 165742 and Subsequent, A1-H60SA-560-100, Naval Air Systems Command, Washington, DC, 1 October 2009. Technical Manual, Organizational Maintenance, Principles of Operation, Integrated Flight Controls, Navy Model F/A-18A/B/C/D 161353 and Up, A1-F18AC -570-100, Naval Air Systems Command, Washington, DC, 1 October 2003. AIII-6

APPENDIX IV Answers to End of Chapter Questions Chapter 1 – Basic Physics

1-1. B 1-2. C 1-3. B 1-4. D 1-5. C 1-6. A 1-7. B 1-8. A 1-9. A 1-10. A 1-11. C 1-12. D 1-13. C 1-14. B 1-15. D 1-16. C 1-17. A 1-18. B 1-19. B 1-20. D 1-21. B 1-22. D 1-23. A 1-24. B 1-25. C 1-26. D 1-27. C 1-28. B 1-29. D 1-30. C 1-31. A 1-32. D 1-33. B 1-34. C 1-35. A 1-36. D 1-37. C 1-38. B 1-39. A 1-40. C 1-41. A 1-42. D 1-43. D 1-44. C

AIV-1

Chapter 2 – Electrical Maintenance and Troubleshooting

2-1. B 2-2. C 2-3. A 2-4. D 2-5. B 2-6. C 2-7. D 2-8. C 2-9. A 2-10. D 2-11. C 2-12. B 2-13. A 2-14. D 2-15. B 2-16. C 2-17. A 2-18. D 2-19. B 2-20. C 2-21. B 2-22. A 2-23. D 2-24. B 2-25. A 2-26. D 2-27. B 2-28. B 2-29. C 2-30. A 2-31. B 2-32. D 2-33. B 2-34. B 2-35. A 2-36. B 2-37. A 2-38. D 2-39. B 2-40. D 2-41. A 2-42. A 2-43. C 2-44. A 2-45. C 2-46. B 2-47. C 2-48. D 2-49. D 2-50. A 2-51. C 2-52. D 2-53. B 2-54. D 2-55. A 2-56. C 2-57. B 2-58. C

AIV-2

Chapter 3 – Power Generation and Control Systems

3-1. C 3-2. A 3-3. C 3-4. D 3-5. B 3-6. C 3-7. C 3-8. D 3-9. A 3-10. B 3-11. C 3-12. D 3-13. B 3-14. D 3-15. C 3-16. A 3-17. D 3-18. B 3-19. B 3-20. A 3-21. D 3-22. B 3-23. A 3-24. B 3-25. D 3-26. C 3-27. C 3-28. A 3-29. C 3-30. B 3-31. D 3-32. C 3-33. D 3-34. B 3-35. D 3-36. C

AIV-3

Chapter 4 – Aircraft Electrical Systems

4-1. B 4-2. A 4-3. D 4-4. B 4-5. C 4-6. D 4-7. B 4-8. C 4-9. A 4-10. B 4-11. C 4-12. A 4-13. D 4-14. D 4-15. B 4-16. D 4-17. C 4-18. C 4-19. D 4-20. C 4-21. D 4-22. A 4-23. A 4-24. B 4-25. C 4-26. A 4-27. A 4-28. D 4-29. B 4-30. C 4-31. D 4-32. D 4-33. B 4-34. B 4-35. B 4-36. A 4-37. C 4-38. A 4-39. A 4-40. C 4-41. B 4-42. D 4-43. C

AIV-4

Chapter 5 – Aircraft Power Plant Electrical Systems

5-1. C 5-2. A 5-3. B 5-4. D 5-5. B 5-6. C 5-7. B 5-8. A 5-9. B 5-10. C 5-11. C 5-12. D 5-13. B 5-14. C 5-15. B 5-16. A 5-17. A 5-18. B 5-19. C 5-20. C 5-21. D 5-22. B 5-23. D 5-24. D 5-25. C 5-26. B 5-27. B 5-28. D 5-29. C 5-30. A 5-31. D 5-32. A 5-33. C 5-34. D 5-35. B 5-36. A 5-37. C 5-38. D 5-39. B 5-40. A 5-41. C 5-42. D 5-43. A 5-44. C 5-45. B 5-46. D 5-47. B 5-48. B 5-49. C 5-50. D 5-51. A

AIV-5

Chapter 6 – Aircraft Instruments

6-1. C 6-2. B 6-3. A 6-4. D 6-5. A 6-6. A 6-7. A 6-8. C 6-9. B 6-10. D 6-11. B 6-12. B 6-13. C 6-14. C 6-15. B 6-16. A 6-17. A 6-18. B 6-19. A 6-20. D 6-21. D 6-22. B 6-23. A 6-24. A 6-25. D 6-26. B 6-27. D 6-28. A 6-29. B 6-30. B 6-31. A 6-32. D 6-33. A 6-34. B 6-35. C

Chapter 7 – Compass and Inertial Navigation Systems

7-1. C 7-2. A 7-3. C 7-4. B 7-5. D 7-6. C 7-7. B 7-8. A 7-9. C 7-10. A 7-11. D 7-12. B 7-13. B 7-14. C 7-15. A 7-16. C 7-17. A 7-18. D 7-19. B 7-20. D 7-21. B 7-22. C 7-23. C 7-24. C 7-25. B 7-26. D 7-27. A 7-28. C 7-29. B 7-30. B 7-31. C 7-32. C 7-33. D 7-34. C

AIV-6

Chapter 8 – Automatic Flight Control and Stabilization Systems

8-1. C 8-2. A 8-3. D 8-4. B 8-5. A 8-6. D 8-7. B 8-8. C 8-9. A 8-10. B 8-11. D 8-12. D 8-13. B 8-14. C 8-15. B 8-16. A 8-17. C 8-18. B 8-19. B 8-20. B 8-21. A 8-22. B

AIV-7

ASSIGNMENT Chapter 1 BASIC PHYSICS 1-1. What element is being measured when using the term meter?

A. Distance B. Mass C. Time D. Heat

1-2. What is the English equivalent to 1 meter?

A. 2.5 feet B. Approximately 1 yard C. Approximately 3 yards D. 4.5 feet

1-3. What is the difference between speed and velocity?

A. Speed is the time of travel and velocity is the distance. B. Speed is not measured when using velocity. C. Velocity is a vector quantity; it is speed in a given direction. D. Velocity is the fastest time an object travels; speed is the average time.

1-4. What are the four types of temperature scales?

A. Celsius, Fahrenheit, Kelvin, and Rankine B. Hot, Cold, Fahrenheit, and Celsius C. Atmospheric, Liquid, Solid, and Gas D. Freezing, Boiling, Steam, and Vapor

1-5. What two terms describe the range of sound the human ear can distinguish?

A. Frequency and pressure B. Normal and upper C. Threshold of audibility and threshold of feeling D. Intensity and decibel

1-6. What are the three parts of the atom?

A. Positive, Negative, and Neutral B. Nucleus, Electrons, and Number C. Outer, Inner, and Central D. Proton, Neutron, and Electron

1-7. How is the atomic weight of an element determined?

A. By the subatomic particles B. By the number of protons and neutrons in its nucleus C. By the deuterium D. By the isotopes

1-8. In forming a compound, what part of the atom changes?

A. Proton B. Electron C. Neutron D. Electron shell

1-9. Which of the following is an advantage of liquids when it is applied to aviation?

A. Liquid can withstand extreme heat. B. Hydraulic energy is transmitted around corners without gears or levers. C. Liquids can be used in hydraulic hoses. D. Liquids weigh less than mechanical components.

1-10. What person formulated the following conclusion, "For a constant temperature, the product of the volume and pressure of an enclosed gas remains constant”?

A. Kelvin B. Gas C. Boyle D. Charles

1-11. What is the concept defined by the statement, "Two objects can't occupy the same space at the same time."?

A. Energy is constantly exchanged from one object to another. B. Matter possesses energy. C. Anything that occupies space has mass. D. Impenetrability of matter.

1-12. What is meant by the term acceleration?

A. A decrease in mass B. A change in the state of rotational motion C. An increase or decrease in speed and/or a change in direction of motion D. An increase of mass when an object is in motion

1-13. How is the specific gravity of a substance described?

A. The ratio of the density of the substance to the density of water B. By its porosity C. The pressure of the substance D. By its kinetic energy

1-14. What are the two classes of circular motion?

A. Radial and circular B. Rotation and revolution C. Linear and center D. Gyro and ball

1-15. When does an object have potential energy?

A. When it can radiate energy B. When the object can expand C. When it can do work, such as a wound clock spring D. Objects only transfer energy

1-16. What type of bearing is used in many types of machinery to minimize friction and maximize efficiency?

A. Flat B. Sliding C. Self-lubricating D. Machine

1-17. When an object is revolving, what force tries to oppose centripetal force?

A. Inertial friction B. Centrifugal C. Gravity D. Motion

1-18. Why are wood handles used on soldering irons?

A. Wood handles transfer heat evenly. B. They are poor conductors of heat. C. Wood is an excellent conductor of heat. D. They are used for ease of design.

1-19. For an object to become a good absorber of heat, what color is it normally painted?

A. Blue B. Silver C. Yellow D. Dull black

1-20. Which is an effect on light waves when they meet a substance?

A. Absorbed by the substance B. Substance changes color of the light waves C. Increases the light waves’ brightness D. Light wave divides into the basic colors

1-21. What is measured by the foot-candle?

A. Lumen B. Watt C. Intensity of incident light D. Porosity

1-22. Which objects act as refractors?

A. Prisms, positive lenses, and negative lenses B. Solids and Liquids C. Bright solid objects D. Dull colored objects

1-23. What action must be applied to an object to overcome inertia?

A. A decrease in uniform motion B. A push or pull that exerts a force on the body C. A greater tendency of a body to maintain uniform motion D. A decrease of gravitational force

1-24. Which is used in airborne installations to aid convection?

A. Aluminum and copper B. Insulators and conductors C. Liquids and steam D. Fans and blowers

1-25. What happens when light passes through a transparent substance?

A. It is stopped. B. It is refracted. C. It is absorbed. D. It does not change.

1-26. What is lost when energy is expended?

A. Gravitational force B. Thermal expansion C. Efficiency D. Energy

1-27. What is meant by the term intensity of illumination?

A. The amount of reflection of the light waves B. The amount of light received per unit area at a distance from the source C. The refraction level of the light waves when it travels through a transparent substance D. The amount of the ray of light as it passes through a flat sheet of glass

1-28. What is the minimum change of sound level that the human ear perceives?

A. A decibel B. 2 decibels C. 5 decibels D. 10 decibels

1-29. The number of protons in the nucleus of the atom is referred to as the ________.

A. Basic substance B. Characteristics of an element C. Atomic number of the element D. Nuclei

1-30. What is defined as a gas element that does not combine chemically with any other element?

A. Atom B. Electron C. Inert element D. Proton

1-31. The attracting force that holds ions together in the molecular form is known as?

A. Ion B. Valence bond C. Compound D. Negative charge

1-32. The natural tendency of a moving body is to move in a manner so that the center of gravity travels in a straight line is known as?

A. Gravitational effect B. Center of rotation C. Circular motion D. Linear motion

ASSIGNMENT Chapter 2 ELECTRICAL MAINTENANCE AND TROUBLESHOOTING 2-1. Under what category (ies) can all maintenance performed on naval aircraft be grouped?

A. Preventive only B. Unscheduled only C. Scheduled and unscheduled D. Scheduled and preventive

2-2. An accident-free naval career can best be achieved by following which of the following courses of action?

A. Constantly reading technical manuals B. Reading all naval rules and regulations C. Making a list of all potential work hazards D. Taking a common-sense approach towards safety

2-3. If electrical equipment is to be repaired, what action should you take before beginning the actual work?

A. Remove the fuses for the associated circuits B. Short out the main supply switches C. Secure the main power switches in the open position and properly tag them D. Station an individual with a fire extinguisher near the work area

2-4. If you are working on high-voltage circuits or around wires having exposed surfaces, you should keep tools and equipment that have metal parts what minimum number of feet from the work area?

A. 9 feet B. 2 feet C. 5 feet D. 4 feet

2-5. The intensity of electrical shock is determined by which of the following properties?

A. Current B. Voltage C. Impedance D. Electromagnetic force

2-6. Why should a person not move after receiving an electrical shock?

A. The heart is temporarily weakened B. Muscles have been damaged C. Nerves have been damaged D. The brain is impaired

2-7. When fighting an electrical fire, you should use which of the following fire- extinguishing agents?

A. Foam B. Water (H2O) C. Soda and water D. Carbon dioxide (CO2)

2-8. If you swallow gasoline, which of the following actions should you take?

A. Swallow three glasses of salt water to induce vomiting B. Drink large amounts of milk or water and take 4 tablespoons of vegetable oil, if available C. Take two aspirins and two glasses of water D. Swallow a solution of bicarbonate of soda and water

2-9. Which of the following statements describes the hazards of compressed air?

A. It can inject minute foreign bodies into the skin B. It can cool the cell tissue and clean infected wounds C. It can pass through thin sheet metal and cause fatal injury D. It causes a disturbance of dust particles and cause faulty equipment

2-10. When using compressed air to clean out fixtures and jigs, you should observe the proper safety precautions. Also, you should maintain the air pressure below what maximum value PSI?

A. 10 B. 20 C. 30 D. 40

2-11. When using tools, you should observe which of the following rules?

A. Use tools for their intended purpose B. Maintain tools in faulty condition and repair as much as possible C. Use the tool even if not working properly until it’s replaced D. Maximize tool usage by using it for other purposes not designed

2-12. Which of the following is the cause of most accidents in electrical and electronic work centers?

A. Moving machinery B. Carelessness C. Improper grounding D. Exposed electrical fixtures

2-13. If one of your tools becomes worn, damaged, or broken, you should report this fact to what person?

A. Crew leader B. Division officer C. Work center supervisor D. Material control officer

2-14. What alloy is used to make most nonmagnetic tools?

A. Cadmium B. Nickel-iron C. Beryllium-copper D. Copper-Constantan

2-15. When you find a damaged power tool electrical cord, what action should you take?

A. Cover it with rudder tape B. Shorten the cord to remove the damaged part C. Repair the damage with insulating tape D. Replace the cord

2-16. Which of the following is NOT a safety practice to follow when using a soldering iron?

A. Provide ventilation for the iron while it is on its rest rack B. Hold small soldering jobs with pliers or clamps C. Disconnect the iron during temporary absences from the work area D. Shake the iron to get rid of excess solder

2-17. While using an electric drill, you experience an electrical shock. Which of the following conditions is the most likely cause?

A. The voltage source is too high B. The voltage source is too low C. An incorrectly grounded drill D. An overloaded drill

2-18. What color is the safety ground wire for electrical tools?

A. Black B. White C. Green D. Red

2-19. You need to apply voltage to a power tool having a three-wire system. The receptacle is a two-wire type. To connect the tool to this voltage source, you should use an adapter with an external ground wire and connect it in which of the following configurations?

A. Tape the exposed ground wire terminal B. Connect the safety ground to a good ground before plugging in the tool C. Connect the safety ground wire of the adapter to the tool case D. Connect the safety ground wire to the center screw of the receptacle before plugging in

2-20. Discrepancies found before, during, or after a flight require what type of maintenance?

A. Preventive B. Unscheduled C. Scheduled D. Field

2-21. You are troubleshooting an electrical device that is not receiving any power. What check should you make first?

A. Check fuse or circuit breaker B. Check the power source C. Check for loose connector pins D. Check for visible indications of trouble

2-22. Which of the following meters should you use when troubleshooting an open circuit?

A. Voltmeter B. Wattmeter C. Digital multimeter D. Ammeter

2-23. Before replacing a major component in an aircraft, the AE should make which of the following determinations?

A. Whether the component is defective B. Whether the intended replacement is a suitable substitute C. Whether the repair will require a test flight D. Whether the appropriate work center has been assigned the replacement task

2-24. Before making an adjustment to any system, you should consult which of the following publications?

A. MIM B. IPB C. NAVSUP 2002 D. NATOPS

2-25. An ammeter is connected into a circuit in which of the following ways?

A. In parallel with the circuit B. In series-parallel with the circuit C. In parallel with the power supply D. In series with the circuit

2-26. Which of the following conditions is the most probable cause for a grounded circuit?

A. A blown fuse B. A tripped circuit breaker C. Frayed insulation on wiring D. Loose terminal lugs

2-27. Which of the following types of meters are contained in the digital fluke multimeter?

A. Voltmeter, frequency meter, and ohmmeter B. Frequency meter, ammeter, and voltmeter C. Frequency meter, ohmmeter, and ammeter D. Ammeter, voltmeter, and ohmmeter

2-28. A permanent-magnet, moving-coil meter mechanism can be adapted to measure alternating current and voltage if it is used with which of the following devices?

A. A transformer B. A transponder C. A rectifier D. A reactor

2-29. To test insulation for high resistance, grounds, and leakage, what meter should you use?

A. VTVM B. Megger C. Ohmmeter D. Digital multimeter

2-30. A megger is prevented from exceeding its rated output voltage by the action of a ________.

A. friction clutch B. voltage regulator C. diode limiter D. variable resistor

2-31. Which of the following values can be measured by using an oscilloscope?

A. Frequency and voltage amplitude B. Phase differences and temperature C. Electronic circuit’s stages D. Synchronizing time-delay circuits

2-32. What term is used to define abnormal resistance or impedance that interferes with the normal signal flow?

A. Discontinuity B. Distortion C. Reflectometry D. Reduction

2-33. What instrument should you use to troubleshoot fuel quantity coaxial cables?

A. Time-domain reflectometer B. Whetstone bridge C. Ammeter D. Phase-angle voltmeter

2-34. The output of the phase detector in a phase-angle voltmeter is proportional to the signal amplitude multiplied by which of the following angles of phase difference?

A. Sine B. Cosine C. Tangent D. Cotangent

2-35. Maximum deflection of the phase angle voltmeter occurs when what phase relationships exists between the two signals?

A. 0° or 90° B. 0° or 180° C. 90° or 120° D. 90° or 180°

2-36. To be classified as a cable, a single conductor must have which of the following characteristics?

A. Be insulated and designed to carry RF energy B. Be insulated and have a metallic braided shield C. Be covered by a metal shield and designed to carry RF energy D. Be covered by a metal shield and have at least a 00 wire size

2-37. To replace an aircraft electrical wire, you must determine the correct size and type of wire to use. To make this determination, what publication should you consult first?

A. Maintenance Instructions Manual B. Aircraft IPB C. Military Specifications, MIL-W-5088 (latest edition) D. NAVAIR 01-1A-505 (series)

2-38. Aluminum has the tendency to flow away from a point where pressure is applied. This tendency is known as________.

A. flowing B. crystallization C. creep D. feed through

2-39. When stamping wires or cables, the distance between markings should not exceed what maximum distance?

A. 24 inches B. 15 inches C. 3 inches D. 6 inches

2-40. You are reading a wire identification code. Which of the following types of information can you gain?

A. Circuit function B. Circuit manufactures C. Wire age D. Wire length

2-41. Which of the following letters is NOT used to identify a wire segment?

A. E B. O C. X D. Z

2-42. What letter suffix in the wire identification identifies the wire as being a ground?

A. A B. B C. C D. N

2-43. Heat-shrinkable tubing has which of the following advantages?

A. It insulates wire terminals and waterproofs wire splices B. It provides wire strength and support C. It reduces D. It replaces wire terminal covers

2-44. When used on aircraft electrical wiring, the recommended power rating range for general-use soldering irons is within which of the following ranges?

A. 13 to 60 watts B. 20 to 500 watts C. 55 to 600 watts D. 60 to 200 watts

2-45. Before reusing items of mounting hardware, what determination should you first make?

A. Whether they exceed the specifications for their intended use B. That they are the same size and shape of the specified items C. If their reuse is prohibited by existing directives D. That they are not damaged and exceed the specification for the items required by the IPB

2-46. You have temporarily installed suitable substitute mounting parts. When should these parts be replaced?

A. During the next periodic inspection B. During the time the aircraft is at NADEP C. When the substitute parts become defective D. When the required parts become available

2-47. Which of the following considerations should you observe when substituting mounting parts?

A. Color B. Availability C. Magnetic properties D. Accessibility

2-48. What is the reason for moisture proofing solder-type electrical connectors?

A. To reinforce the connector B. To reduce the possibility of the connector cracking C. To improve the connector’s dielectric characteristics D. To protect the connector for future use

2-49. The preferred method for attaching cable terminals to terminal blocks requires the use of what items of hardware?

A. An anchor nut and lock washer B. An anchor nut and flat washer C. A standard nut and lock washer D. A standard nut and washer

2-50. Shielded conduit should be supported by the use of what type of clamp?

A. AN 742 B. Strap C. Bonded D. Nonbonded

2-51. When long runs of cable between panels need to be supported, which of the following types of clamps is preferred?

A. Strap B. Plastic C. Bonded D. AN 742

2-52. When installing a cable through a lightening hole, you should use a grommet (rubber cushion) if the cables’ distance from the edge of the hole is less than what minimum distance?

A. 1/4 inch B. 3/8 inch C. 1/2 inch D. 5/8 inch

2-53. In addition to supporting and protecting electrical wires, what other advantage does conduit offer?

A. It protects against heat radiation B. It provides radio shielding C. It provides bullet deflection D. It supports adjacent cables

2-54. What type of safety wire should you use to secure an oxygen regulator?

A. Lockwire B. Seal wire C. Shear wire D. Soft steel wire

2-55. If an aircraft were improperly bonded, which of the following conditions would exist?

A. An increased likelihood of fire and a noisy radio receiver B. An increased likelihood of lightning strikes C. A decreased chance of lightning strikes D. An increased likelihood of electrical failures

2-56. A primary objective of bonding is to provide an electrical path of:

A. High dc resistance and high RF impedance B. Low dc resistance and high RF impedance C. High dc resistance and low RF impedance D. Low dc resistance and low RF impedance

2-57. When using methyl chloroform to clean electrical equipment, you should remove the equipment from the solution within what maximum length of time?

A. 5 minutes B. 5 to 15 minutes C. 15 to 30 minutes D. 30 to 60 minutes

2-58. When a printed circuit is manufactured by the photoetching process, what portions of the plastic or phenolic sheet are actually photographically exposed?

A. All areas covered by light sensitive enamel not covered by the circuit template B. Only areas covered by the circuit template C. Only areas where circuit components, such as resistors, are attached D. Only areas that act as wires

2-59. Exposed copper is removed during the etching process, and the unexposed copper surfaces are protected by?

A. Enamel B. Solder C. The printed circuitry overlay D. The exposed copper smear

2-60. What color paint is normally used for line test equipment?

A. Yellow B. Orange C. White D. Red

2-61. The Huntron Tracker 2000 quadrant 3 shows which of the following displays?

A. Positive voltage and negative current B. Negative current and negative voltage C. Positive current and negative voltage D. Positive current and positive voltage

2-62. When using the Huntron Tracker 1000 or 2000, you should begin testing in which of the following ranges?

A. Low B. High C. Medium D. Medium-low

2-63. When testing analog circuits or devices with a Huntron Tracker, you should use the low range of the tester for which of the following reasons?

A. Defects will show easier and the internal impedance makes it more likely the device under test will load the tester B. Defects are easier to find in the medium range C. The internal impedance will cause the device under test to load the tester D. Defects will show easier and the internal 54 ohm impedance makes it less likely that parts in parallel with the device under test will load the tester

2-64. When using JETCAL to perform a functional ground test of the EGT system, heat for the thermocouples is provided by ________.

A. heater probes B. exhaust gas C. the JETCAL potentiometer D. the aircraft heating and air conditioning system

2-65. An external ac power supply is required to supply electrical power to a JETCAL analyzer that is being used to make which of the following checks?

A. Engine speed B. The EGT circuit for shorts and grounds C. EGT indicators D. The resistance of the EGT circuit

2-66. Which of the following pulses is/are generated by the synchrophaser test set?

A. Slave pulse only B. Master pulse only C. Slave and master pulses D. Tachometer pulse

2-67. The gain test readout from the synchrophaser test set is provided by what front panel component?

A. The null meter B. The galvanometers C. The calibrated potentiometer D. The feedback potentiometer

2-68. Which of the following test sets provides regulated pitot and static pressure for evaluating the performance characteristics of air data systems, aircraft pneumatic instruments, and other auxiliary equipment?

A. TTU-378 A/E B. TS-4508/U C. TTU-27/E D. AN/PSM-17A

2-69. The AN/PSM-21A air-conditioning test set is used to troubleshoot and check electrical components of which following systems?

A. Cabin pressure B. Pitot temperature C. Aircraft indexer D. Stall warning

2-70. When operating the TTU-597/E test set, the temperature range should be between what degrees?

A. -40° to +55° C B. 0 to 30,000° C C. -55° to +40° C D. 55° C and up

2-71. The variable speed drive pad accommodates the tachometer generator during testing. In what rpm range does the control knob vary the speed of the tach-gen?

A. 60 - 400 rpm B. 115 rpm C. 0 - 5,000 rpm D. 0 - 50,000 rpm

2-72. Which function of the TTU-597/E uses a bank of very accurate and stable capacitors arranged parallel and selected under software control to provide 11 to 10,000pF for TANK and 11 to 1000pF for COMP for simulating fuel quantities?

A. Voltage Measurement B. Resistance Measurement C. AC Capacitance Measurement D. Capacitance Simulation

2-73. Test Program Sets are developed for a unique unit under test, contain four basic elements and are used to verify the performance of a unit under test and ________.

A. isolate failures to a required level B. determine the exact time of failure C. provide corrective action D. provide repair procedures

2-74. Under which of the following conditions is generated static electricity decreased?

A. Dry air B. Humid air C. Cold air D. Hot air

2-75. Which of the following is NOT an ESD prime generator?

A. Synthetic mats B. Vinyl C. Common solder suckers D. Carbon impregnated polyethylene

2-76. What is the minimum resistance for personnel ground straps?

A. 25,000 ohms B. 150,000 ohms C. 250,000 ohms D. 500,000 ohms

2-77. Which of the following procedures should NOT be performed when working on ESD-sensitive devices?

A. Ground the work area, wrist straps, and equipment B. After testing, replace shorting devices and protective packaging C. Perform dielectric strength tests D. Use of a Simpson 260 meter or equivalent to test components

ASSIGNMENT Chapter 3 POWER GENERATION AND CONTROL SYSTEMS 3-1. What are the advantages of using transformers, transformer rectifiers vice inverters, and dynamotors for supplying aircraft power?

A. Accurate and requires less current B. Lightweight, more reliable, and simple to maintain C. Poles are fixed quantity and uses smaller wiring D. Provides control and protection

3-2. What power system requirements do modern naval aircraft have?

A. Three phase, 15-/120-volt, 500-hertz ac B. Three phase, 120-/208-volt, 400-hertz ac C. Three phase, 115-/380-volt, 400-hertz ac D. Three phase, 500-volt, 400-hertz ac

3-3. What determines the voltage frequency of an ac generator?

A. The size of the rotors B. The number of transformers and inverters C. The number of magnetic poles and the rotor rpm D. The number of transformer rectifiers

3-4. What are the two types of ac generators?

A. Stationary and rotating B. Brush and brushless C. Rotating magnetic and fixed polarity D. 120 volts and 208 volts

3-5. What supplies the output ac power of a generator?

A. Poles B. Phase windings C. Rotor D. Stator

3-6. How many phases is an ac generator?

A. 1 B. 2 C. 3 D. 4

3-7. What are the advantages of a brushless generator over a brush-type generator?

A. Weighs less and is interchangeable B. Increased reliability and greater operating time between overhaul C. Easily attaches to the engine pad D. Has a generator shear section to prevent possible damage

3-8. What is a prime mover?

A. Propeller mechanical governor B. Three-phase differential transformer C. Hydromechanical unit D. Device that provides the driving force for a generator

3-9. What are inverters?

A. Speed-governed dc motor B. An emergency source of ac power when normal ac power fails C. Indicator that receives power D. Permanent magnet type generator

3-10. What converts electrical energy to mechanical energy in an inverter?

A. The control box B. The dc motor C. The relays D. The shunt winding

3-11. On what does the rating of an aircraft inverter depend?

A. The dc load B. Speed of the dc motor C. The shunt winding D. The equipment that it will supply

3-12. What is the purpose of a transformer?

A. Convert an input ac voltage into a usable ac output voltage by either stepping up or stepping down that voltage B. Maintain voltage limits to reasonable level than can be used C. Increase the voltage output for ac power D. Decrease ac and dc current to true power

3-13. What is the function of a transformer rectifier?

A. Provides cooling for equipment B. Maintains straight-line dc voltage C. Converts ac input voltage to a dc output voltage D. Connects input ac voltage to the load

3-14. What is the only moving part in a transformer rectifier?

A. Thermostat B. Relay C. Diodes D. Cooling fan

3-15. What is the difference between an autotransformer and an ordinary transformer?

A. The power load in an autotransformer is much less than the ordinary transformer B. The autotransformer does not have a control shaft C. An autotransformer has one winding that is common to both primary and secondary windings D. The ordinary transformer has a continuous variable tap

3-16. Autotransformers offer savings in both size and costs over conventional units. When are these savings greatest?

A. When there is no isolation between primary and secondary positions B. When the turns ratio is less than 2:1 C. When power comes from the primary by magnetic field D. The saving between the cost and weight of an entire winding

3-17. What are the two general types of instrument transformers?

A. Current and potential B. High voltage and high current C. High current and low voltage D. Primary and secondary

3-18. What are the ways in which emergency power is supplied to naval aircraft?

A. AC generator and transformer-rectifier B. Hydraulically driven ac/dc generator and a motor-generator control unit C. DC output from the stationary rectifier D. Storage batteries, auxiliary power units, and hydraulic motor-driven generator

3-19. What components maintain the aircraft (a/c) storage battery in a charged state?

A. The hydraulic motor B. The transformer C. AC generator and transformer rectifier D. DC motor

3-20. What manual is used for information on securing, storing, servicing, and handling a/c storage batteries?

A. NAVAIR 7-5 AD-1 B. NAVAIR 7-15BA C. NAVAIR 17-15BAD-1 D. NAVAIR 17-75BAD-1

3-21. Why can the hydraulic-driven emergency generator only power a small number of circuits?

A. The kVA rating is higher than the primary generators B. The kVA rating is much lower than the primary generators C. The output ac power is limited to three circuits at any one time D. The output ac/dc power is limited to four emergency circuits at one time

3-22. How is the generator cooled?

A. Ambient air B. Coolant fluids C. By hydraulic fluid D. Forced air

3-23. What is the purpose of the motor-generator control?

A. Provides power flow to the rectifier B. Provides signals to the dc panel C. Protects the circuits in the motor generator D. Provides voltage regulation for, and detection of, motor-generator output

3-24. What is the function of the solenoid control valve?

A. De-energize and open pressure switches B. When primary electrical power fails, the valve de-energizes and opens, routing hydraulic pressure to the hydraulic motor, driving the generator C. Routes hydraulic pressure to the hydraulic motor D. Prevents out-of-tolerance power from being connected

3-25. What is the function of an APU?

A. Provide compressed air at the output end of the engine B. To furnish electrical power when engine-driven generators are not operating, external power is not available, or the engine- driven generator fails C. Develops power by compressing ambient air with a two-stage centrifugal compressor D. Funnels compressed air from a gas-turbine compressor onto the turbine blades

3-26. What is needed for APU starting?

A. Aircraft battery and fuel B. Compressed air and fuel C. Hydraulic power, bleed air, and aircraft battery D. Electrical power or aircraft battery

3-27. What drives the generator on a GTCP-95?

A. Accessory assembly B. Compressor and turbine assembly C. Rotating shaft of the turbine wheel D. Shaft power at the main output drive pad

3-28. What is the purpose of the centrifugal speed switch assembly?

A. Controls the sequence of operation of various electrical components B. Adjusts current to the generator to vary the load C. Increases voltage and the load on the generator D. Prevents the possibility of one circuit feeding another

3-29. How many sub-switches make up the centrifugal speed switch assembly?

A. 2 B. 3 C. 4 D. 5

3-30. How are centrifugal speed switches adjusted?

A. By the changing the holding relays, oil pressure, and centrifugal switch B. By turning the knife-edged fulcrum C. By applying spring tension to the lever arm with the three-adjustment screws D. By actuating the shaft to move the lever arm

3-31. Changing the setting of the 35 percent switch also affects what?

A. The lowest percent speed adjustment B. The overspeed switch C. The setting of the 95- and 106-percent switches D. The fuel flow to the combustion chamber

3-32. At what rpm is the starter disengaged?

A. 35 percent B. 95 percent C. 106 percent D. 135 percent

3-33. What does the battery voltmeter indicate when the ac buses are powered?

A. Maintenance battery voltage B. Battery charger voltage C. 115 vac D. Undervoltage

3-34. What component provides 28 vdc from the essential bus backup to the essential bus?

A. Left power supply B. Right power supply C. Isolated permanent magnet generator D. Battery

ASSIGNMENT Chapter 4 AIRCRAFT ELECTRICAL SYSTEMS 4-1. What are the two most common types of bulb bases?

A. Single wire and single contact B. Doubled filament and index C. Single and double contact bayonet D. Clear and frosted

4-2. On which wing tip will the green navigation light be found?

A. Left B. Right C. Both left and right D. Wing tips do not have navigation lights

4-3. The science of liquid pressure and flow is known as ________.

A. Rotary-mechanical B. Hydraulics C. Fluid power D. Actuating pressure

4-4. What system provides the pilot with adequate directional control during ground operations?

A. Push button switch B. Rudder pedal C. Hydraulic steer-damp D. Nose wheel steering

4-5. The pneumatic system compressor comes on when system pressure drops below what PSI value?

A. 100 B. 1,500 C. 2,750 D. 3,150

4-6. Where is the resistance heating element installed in electrically heated windshields?

A. Outer surface of windshield B. Current buss bars C. Inner surface of the outer pane of glass D. Lower surface of the inner pane of glass

4-7. What type of light provides the landing safety officer with a visual indication of a carrier aircraft’s safe or unsafe landing configuration?

A. Index B. Position C. Formation D. Approach

4-8. What is the landing configuration of an aircraft if the Landing Signal Officer (LSO) observes flashing approach lights?

A. The landing gear only is up B. The landing and arresting gears are up C. The arresting gear is not fully extended D. Normal operation

4-9. The fuselage formation lights are connected in parallel with and controlled by the same switches as what other lights?

A. The fuselage signal lights only B. The wingtip formation lights only C. The fuselage signal and wingtip formation lights only D. The fuselage signal, wingtip formation, and navigation lights

4-10. What lighting feature is provided to aid a crew member who is reading a chart?

A. Red floodlights B. White floodlights C. Extension lights D. Momentary contact switches to bypass the rheostats on floodlights

4-11. In aircraft hydraulic systems, the AE maintains circuits that control the fluid ________.

A. viscosity B. flow C. shape D. pressure

4-12. Which of the following components directs the fluid flow in a hydraulic system?

A. Reservoir B. Pump C. Selector valve D. Actuating unit

4-13. The catapulting system’s launch bar warning light illuminates when which of the following conditions exist?

A. The launch bar is up and locked with weight off the landing gear B. The launch bar is up and locked with weight on the landing gear C. Solenoid B is energized D. Solenoid A is energized

4-14. Which of the following is a correct temperature equivalent?

A. 100°C = 212°F B. 32°C = 212°F C. 32°F = 100°C D. 0°F = 32°C

4-15. In an aircraft, cabin air pressure is controlled by the operation of which of the following components?

A. A safety valve and a manual dump control B. A pressure regulator and a safety valve C. A manual dump control and a pressure regulator D. Bleed air shut off valve

4-16. Which knob controls warm, dry air used as muscle air for the Avionics flow control valve and the cabin flow valve?

A. BLEED AIR B. TEMP CABIN C. CABIN PRESS D. ANTI ICE

4-17. Windshield overheating is prevented by the combined actions of a shutoff valve and what other component?

A. A thermostat B. A dump valve C. A thermistor D. An electronic temperature controller

4-18. What method is used to deice the empennage of P-3 aircraft?

A. A powered, controlled system provides constant heat to the empennage B. Hot engine bleed air is circulated under the surfaces C. Leading edges of the vertical and horizontal stabilizers are electrically heated D. Heat from the cabin routed through control valves heats the empennage

4-19. How many components are in a basic hydraulic system?

A. One B. Two C. Three D. Four

4-20. Which component of the Wing Flap Asymmetry system will allow a confidence check of the Wing Flap Asymmetry Shutoff valve and Shutoff Valve Test light without tripping the system?

A. Wing Flap Control Valve B. Wing Flap Asymmetry Relay C. Wing Flap Brake Relay D. Test/Reset Switch

4-21. The Wing Flap system is made up of the Wing Flap Asymmetry system and the Wing Flap Position indicator system. What do both systems combined detect for?

A. Automatic arrest of asymmetry and unequal extension B. Unequal extension of the wing flaps and the amount of wing flap extension from 0 to 50 percent C. Automatic arrest of asymmetry of the wing flap and amount of wing flap extension D. Amount of wing flap extension from 0 to 50 percent

4-22. The Landing Gear Warning System uses the landing gear control levers as a visual indication for the landing gear position? What is the other type of indication to visually check the landing gear is up or down?

A. Digital display B. Flashing wheels indication C. Three down and locked lights D. Warning flags

4-23. What mechanically prevents movement of the landing gear control lever from the wheels down position when the aircraft weight is on the gear?

A. A solenoid B. Torque-link switch C. Armature pin D. Shock strut oleo

4-24. To maintain cabin temperature, the Cabin Temperature Channel of the temperature controller uses how many bridge circuits?

A. one B. two C. three D. four

ASSIGNMENT Chapter 5 AIRCRAFT POWER PLANT ELECTRICAL SYSTEMS 5-1. For a fire to occur, what elements must be present?

A. Oxygen and heat only B. Heat and a combustible material only C. Oxygen and a combustible material only D. Heat, oxygen, and a combustible material

5-2. In an electronic ignition system, what component develops the voltage that produces a spark?

A. exciter B. dynamotor C. transformer D. booster coil

5-3. In an electronic ignition system, ignition is discontinued when what percentage of the rated engine speed is reached?

A. Between 15 and 20 percent B. Between 25 and 45 percent C. Between 45 and 65 percent D. Between 65 and 75 percent

5-4. In turbine-powered aircraft, what relationship, if any, exists between engine power and turbine temperature?

A. They are directly proportional B. They are inversely proportional C. They tend to cancel each other D. There is no relationship between engine power and turbine temperature.

5-5. By what means does an engine temperature control system on turboprop engines control engine torque?

A. A converging and diverging panel circuit B. Power and condition levers C. A restrictive airflow fan circuit D. Variable airflow

5-6. The fuel shutoff valve is electrically closed on engine shutdown by placing the condition lever in what position?

A. Feather B. Run C. Ground stop D. Air start

5-7. The engine coordinators function to coordinate the power and condition levers along with what other component(s)?

A. The fuel control only B. The propeller and fuel control only C. The electronic fuel trimming circuit only D. The fuel control, the propeller, and the electronic fuel trimming circuit

5-8. The discriminating device will complete the feather cycle when the condition lever is placed in feather and the power lever is placed in what position?

A. In any position below the flight idle position only B. In any position above the flight idle position only C. In any taxi range position only D. In any position

5-9. The reference temperature and turbine inlet temperature signals sent to the temperature datum control indicate a difference greater than 1.9°C, and a control signal is sent to the temperature datum valve. With the power lever in the temperature controlling range and the TEMP DATUM switch in the AUTO position, the control signal causes the temperature datum valve to ________.

A. adjust the power lever assembly, regulating fuel flow to the engine being controlled B. control a fuel-flow stabilizing pump on the engine being controlled C. regulate the fuel flow to the engine being controlled D. readjust the engine coordinator and trimming circuit on the engine being controlled

5-10. If engine speed is less than 94% and the engine coordinator is set above 66°, to what temperature is the normal limiting temperature automatically set?

A. 730°C B. 830°C C. 978°C D. 1,077°C

5-11. Dual unit thermocouples are radially mounted in what part of the engine?

A. Turbine inlet case B. Compressor section C. Inlet guide vanes D. Turbine section

5-12. By which of the following methods are thermocouples electrically connected to provide an average temperature?

A. In series B. In parallel C. Either 1 or 2 above, as both will provide an average temperature D. In series-parallel

5-13. An air turbine starter can be operated by compressed air from a GTC, an APU, or what other device?

A. Bleed air from an operating engine B. The starter C. An outside air vent D. An emergency air tank

5-14. What is the function of the engine start system’s speed-sensitive control?

A. To synchronize all engine speeds B. To control engine RPM during start cycles C. To prevent the engine from exceeding maximum RPM D. To activate internal switches at predetermined intervals relative to the engine’s normal speed

5-15. The ignition exciter provides which of the following voltages?

A. A stepped-up voltage for firing the ignition plugs B. A 28-volt excitation voltage for closing the ignition relay C. A 28-volt excitation voltage for activating the speed-sensitive control D. A sine-wave voltage for application to the ignition relay solenoid

5-16. What behavior of the paralleling lamp indicates that the secondary element of a fuel pump has failed?

A. It illuminates continuously B. It never illuminates C. It illuminates above 65% RPM D. It illuminates between 16% and 65% RPM

5-17. Along with the temperature datum valve, the fuel control functions to provide a starting fuel flow schedule and to ________.

A. reduce nominal fuel requirements B. prevent engine over temperature and compressor surge C. operate hydraulically actuated fuel cutoff valves D. close the fuel shutoff valve during compressor surges

5-18. At what percentage, if any, of the rated engine RPM does the fuel control shutoff valve open to permit fuel flow to the engine?

A. 16% RPM B. 65% RPM C. 94% RPM D. None

5-19. The temperature datum valve is located in which of the following components?

A. Between the fuel tank and the fuel control B. Between the fuel control and the engine fuel nozzles C. Between the primary fuel pump and the secondary fuel pump D. Between the fuel tank and the primary fuel pump

5-20. What is the function of compressor bleed-air valves?

A. To sequentially close the eight parts on the compressor housing B. To bleed air from the compressor’s fifth stage into its tenth stage C. To reduce the compressor load during starts D. To minimize the starter load on the compressor

5-21. For any slow-cycle operation, the timing cycle for propeller deicers is such that current is supplied to the heating elements for approximately what time period?

A. 17 to 22 seconds B. 25 to 35 seconds C. 40 to 75 seconds D. 80 to 120 seconds

5-22. The propeller deice timer motor is changed from fast speed to slow speed by which of the following actions?

A. Switching a filter in the motor circuit B. Bypassing the variable resistor with two fixed resistors C. Adjusting the variable resistor to provide maximum resistance D. Switching an additional fixed resistor in series with the variable resistor

5-23. Under what condition will a fire warning light illuminate?

A. When the control unit does not monitor resistance changes B. When the resistance of the sensing element does not change with a change in engine compartment temperature C. When the resistance of the sensing element decreases to a predetermined level due to an increase in temperature D. When the resistance of the sensing element increases to a predetermined level due to an increase in temperature

5-24. What is the purpose of the short discriminator circuit in the fire warning system?

A. To illuminate the fire warning lights during a test B. To activate the fire warning system if a short occurs in the system C. To prevent the fire warning system from actuating when an open occurs in the circuit D. To prevent the fire warning system from actuating when a short occurs in the circuit

5-25. Which of the following statements describes the means by which Halon extinguishes an aircraft engine fire?

A. It forms a blanket around the engine’s air passages, smothering the fire B. It cools the burning area to an extremely low temperature, extinguishing the fire C. It uses the oxygen in the compartment at a rapid rate, making the air incapable of supporting a fire D. It displaces the air in the nacelle, making the air incapable of supporting a fire

5-26. What component controls the oil cooler door position when the oil cooler switch is in the automatic mode?

A. Thermostat B. Thermistor C. Magnetic brake D. Solenoid valve

5-27. On an aircraft, what control system serves to vary the exhaust escape area to obtain the desired thrust and to maintain safe operating conditions?

A. Afterburner control system B. Variable exhaust nozzle system C. Main fuel control system D. MIL control system

5-28. What VEN system components serve to schedule, compute, and control engine operation?

A. VEN power unit B. ECA C. VEN actuator D. MFC

5-29. When the throttle is moved to the idle position, the VEN area rapidly moves to almost full open. The full open VEN area has which of the following effects on engine performance?

A. Reduces running temperature and exhaust temperature B. Allows higher idle speed; reduces acceleration time C. Aids starting; lowers thrust D. Reduces idle speed

5-30. Decreasing the VEN area has what effect, if any, on EGT?

A. Increases B. Decreases C. Increases up to 9,000 feet; decreases above 9,000 feet of altitude D. None

5-31. What component provides feedback to the ECA to ensure the VEN is positioned correctly?

A. The VEN synchronizing shaft B. The VEN position transmitter LVDT C. The VEN torque motor feedback circuit D. The metering valve position transmitter LVDT

5-32. What is the function of the VEN power unit?

A. To provide electromechanical feedback to position the VEN B. To provide hydraulic pressure to the actuator to position the VEN C. To provide power for the electrical control assembly D. To provide hydraulic pressure to drive the servomotor when signaled by the ECA

5-33. What is the function of the propeller governor?

A. To control engine speed by varying the pitch of the propeller B. To control engine speed by varying the rate of fuel flow to the fuel nozzles C. To control propeller pitch by varying hydraulic pressure D. To control the propeller by varying engine speed

5-34. The pitch of the propeller blade is varied by porting hydraulic fluid directly to which of the following parts of the propeller piston?

A. Inboard side B. Outboard side only C. Inboard and outboard D. Geared cam

5-35. The synchrophaser will function in which of the following modes?

A. Normal only B. Synchrophasing C. Normal and synchrophasing D. Mechanical

5-36. What is the function of the pulse generator in the synchrophaser system?

A. Provide pulses for phase and speed control of the propellers B. Establish the phase relationship between propellers C. Translate synchrophaser electrical signals into mechanical bias D. Provide pulses directly to the corresponding servomotor

5-37. What is the purpose of the phase and trim control in the synchrophaser system?

A. Control the servomotor train B. Provide pulses for speed control C. Set the phase relationship between master and slave propellers D. Convert electrical signals into mechanical motion

5-38. What is the function of the speed bias servo assembly?

A. To set the phase relationship between the master and slave propellers B. To translate synchrophaser electrical signals into mechanical bias C. To establish the reference pulse for the propellers D. To provide pulses for speed and phase control

5-39. The synchrophaser provides which of the following servomotor control voltages?

A. An ac voltage 90° or 270° out of phase with the excitation voltage B. An ac voltage in phase or 180° out of phase with the voltage applied to the reference C. An ac voltage in phase or 180° out of phase with the excitation voltage D. A negative or positive dc signal voltage

5-40. In a synchrophaser, what governing mode is used to provide improved engine response to transient RPM changes?

A. Throttle lever anticipation mode B. Master governing mode C. Synchrophasing mode D. Normal mode

5-41. In the synchrophaser, the speed derivative circuit senses changes in engine RPM and generates signals having what function?

A. To change the propeller pitch to align the slave propellers to the master propeller B. To dampen engine RPM changes C. To change the propeller pitch to correspond to engine RPM D. To dampen propeller pitch transients

5-42. When the synchrophasing mode of operation is used, all engines except the master engine operate with synchrophasing. With what does the master engine operate?

A. Normal governing only B. Mechanical governing only C. Normal and mechanical D. Hydraulic governing

5-43. What synchrophaser circuit prevents the slave engine from following an overspeeding or underspeeding master engine?

A. The 2% limiting circuit B. The 20% limiting circuit C. The speed bias servo assembly D. The throttle lever anticipation potentiometer

5-44. What is the purpose of resynchrophasing?

A. To overcome small errors of lead and lag about a set point B. To provide for correcting an overshoot C. To move the feedback potentiometer and cancel a portion of the error signal D. To correct for accumulated one direction errors

5-45. What assembly maintains the minimum desired low-pitch angle?

A. Magnetic latching B. Hydraulic lever C. Electrical solenoid D. Low-pitch stop

5-46. What is the purpose of the Beta follow-up stop?

A. Provide a secondary stop setting at the 15° blade angle only B. Provide a secondary low-pitch stop C. Prevent negative-torque system failure D. Prevent minor reductions in blade angle

5-47. The function of the pitchlock mechanism in the propeller is to prevent which of the following conditions?

A. Propeller overspeeding B. Low-pitch oscillations C. Propeller gyrations D. High-pitch oscillations

5-48. Pitchlock is blocked out between blade angles of +57° and +86°. What action does this permit?

A. RPM surges during approaches B. RPM changes during landings C. The blade angle reduction for takeoffs and landings D. The blade angle reduction for air starting

5-49. If the fuel governor and the propeller pitchlock test switch were to be placed in the TEST position, what would be the result?

A. The blade angle would be reset to the proper angle for an air start B. The propeller governor RPM would be reset to permit ground check of pitchlock and fuel governor functions C. A momentary blockout of the pitchlock mechanism D. An increase in pitch

5-50. When needed, the negative torque system functions in what way?

A. To eliminate propeller cycling actions B. To generate a negative torque C. To limit positive horsepower D. To increase blade angle

ASSIGNMENT Chapter 6 AIRCRAFT INSTRUMENTS 6-1. Which of the following statements describes the earth’s atmosphere?

A. The air molecules are closer together at the bottom of the atmosphere than at the top B. The weight of the air pressing down from above determines the air pressure at any given altitude C. The air is denser on the earth’s surface than at an altitude of 1,000 feet D. The air is denser at higher altitudes

6-2. In the pitot-static system, the term pitot represents what type of pressure?

A. Impact B. Ambient C. Barometric D. Stationary

6-3. To compute airspeed, the airspeed indicator uses ________.

A. static pressure only B. impact pressure only C. the sum of impact and static pressures D. the difference between static and impact pressures

6-4. The accuracy of the airspeed indicator readings may be affected by which of the following conditions?

A. Temperature changes in the instrument B. Air turbulence around the pitot tube C. Imperfect scaling of the indicator dial D. Type of instrument indicator

6-5. An aircraft’s Mach indicator reads 0.5 when the airspeed indicator shows 300 knots. If the airspeed were to double to 600 knots, which of the following statements would reflect the aircraft’s speed and Mach indication?

A. The aircraft is at the speed of sound, and the Mach indication is 0.25 B. The aircraft is at the speed of sound, and the Mach indication is 1 C. The aircraft is at one-half the speed of sound, and the Mach indication is 1 D. The aircraft is at the speed of sound, and the Mach indication is 1.5

6-6. Which of the following is a meaning of the term altitude?

A. The distance above the terrains average level B. The distance above mean sea level C. The distance above terrains highest D. The difference in distance from terrains lowest and highest level

6-7. The altitude reading of a properly calibrated altimeter referenced to 29.92 inches of mercury (Hg) is known as the ________.

A. true altitude B. absolute altitude C. pressure altitude D. indicated altitude

6-8. If, at sea level, there is a barometric change of 0.03 inches, the altimeter reading would change by how many feet?

A. 9 B. 15 C. 27 D. 36

6-9. At what altitude, in feet, does the barometric pressure setting for aircraft altimeters change from the local barometric pressure to 29.92 inches?

A. 8,000 B. 13,000 C. 18,000 D. 23,000

6-10. For the pilot to obtain the best performance from the aircraft engine, which of the following altitudes should the pilot know?

A. True B. Density C. Indicated D. Calibrated

6-11. In a Vertical Speed Indicator (VSI), the pointer is driven by ________.

A. the difference between diaphragm and case pressures B. the difference between pitot and static pressures C. a pneumatically driven motor D. a calibrated leak

6-12. What is the purpose of the Air Data Computer (ADC)?

A. To provide accurate air data that is free of aircraft configuration errors B. To compute air data information that is gathered from sensors isolated from air data disturbances C. To sense the characteristics of the air surrounding the aircraft D. To sense the characteristics of the air surrounding the aircraft and correct the data to compensate for aircraft-induced errors

6-13. What are the four data sense inputs to the ADC?

A. Pitot pressure, static pressure, total temperature, and angle of attack (AOA) B. Pitot pressure, static pressure, AOA, and total pressure C. Pitot pressure, total temperature, cabin pressure, and AOA D. Static pressure, cabin pressure, total pressure, and pneumatic differential pressure

6-14. What is the purpose of the potentiometers in the AOA transmitter?

A. To provide an electrical indication of the aircraft fuselage in reference to the angle of attack B. To provide a means for converting voltage into mechanical motion C. To convert electrical signals into mechanical signals indicative of the angle of attack D. To convert mechanical motion into proportional electrical signals

6-15. What is the purpose of the AOA system?

A. To indicate aircraft total pressure with respect to ambient pressure B. To indicate the aircraft attitude with respect to the surrounding air mass C. To provide indications of the air data sensor outputs D. To compute the air data information for the Air Data Computer (ADC)

6-16. Mach number and what other variable cause the most significant errors in indicated static pressure, as detected by the aircraft static ports?

A. AOA B. altitude C. total temperature D. ambient temperature

6-17. What is the definition of impact pressure (Qc)?

A. The weight of the air on the aircraft B. Atmospheric pressure, including disturbances C. The force of the air against the aircraft D. Atmospheric pressure free of disturbances

6-18. In the automatic altitude reporting system, position and altitude reporting is accomplished by which of the following components?

A. Transponders B. Transformers C. Flashing beacons D. Tachometer generators

6-19. The automatic altitude reporting system provides the aircraft’s altitude in which of the following increments of feet?

A. 50 B. 100 C. 150 D. 250

6-20. What total number of dimensions is/are automatically displayed on a radar presentation by the semiautomatic air traffic control system?

A. One B. Two C. Three D. Four

6-21. The AAU-19/A altimeter operates as a standard altimeter when it is placed in which of the following modes?

A. Reset B. Servoed C. Baroset D. Standby

6-22. What components in the AAU-24/A altimeter overcome the effects of the stop- and-jump friction?

A. Servos B. Synchros C. Transponders D. Vibrators

6-23. When used in naval aircraft, the angle-of-sideslip system is used along with which of the following systems?

A. Crosswind landing system B. Gun firing system C. Bombing system D. Rocket firing system

6-24. The AOA indicating system operates by detecting ________.

A. earth field variation B. airflow differential pressure C. the altitude/speed pressure gradient D. the airflow/altitude change rate

6-25. If the AOA of an aircraft is changed, which of the following actions occur(s) in the self-balancing bridge circuit of the AOA system?

A. The transmitter and receiver prevents movement of the potentiometers in the circuit B. A servomotor drives the receiver potentiometer to return the bridge circuit to null C. The receiver prevents movement of the potentiometers in the circuit D. A servomotor drives the transmitter potentiometer to return the bridge circuit to null

6-26. On most naval aircraft, the stall warning system is activated by which of the following systems?

A. Pitot-static B. AOA C. ADC D. Angle-of-sideslip

6-27. Which of the following general characteristics are the most desirable for an instrument gyroscope?

A. Light weight, small size, and low speed of rotation B. Light weight, small size, and high speed of rotation C. Heavy weight, large size, and high speed of rotation D. Heavy weight, small size, and high speed of rotation

6-28. For a gyro to have two degrees of freedom, the platform must have what total number of gimbals?

A. One B. Two C. Three D. Four

6-29. What are the two fundamental properties for gyroscopic action?

A. Stability and rigidity in space B. Precession and centrifugal force C. Stability and centrifugal force D. Rigidity in space and precession

6-30. A spinning gyro precesses when subjected to a deflecting force. Which of the following actions will make the gyro precess at a faster rate?

A. A decrease in the speed of the rotor B. An increase in the force applied C. A decrease in force and speed on the rotor D. An increase in the speed of the rotor

6-31. What action does the gyro horizon perform, relative to its case, to indicate aircraft attitude?

A. The case revolves around the gyro B. The case and gyro spin axis are both free to move with respect to the aircraft C. The gyro spin axis revolves about the case D. The gyro and case are held stationary, and the needles are free to move

6-32. The sphere in the attitude indicator may be centered by a control on the face of the indicator to correct for which of the following flight attitudes?

A. Yaw only B. Roll only C. Pitch only D. Yaw, roll, and pitch

6-33. In a turn-and-bank indicator, what factor(s) determine(s) the position of the ball?

A. Natural forces B. Gyroscopic precession C. Earth’s magnetic lines of force D. Electrical tilting of the plate on which the indicator mounts

6-34. In a flight-coordinated turn, the ball of a turn-and-bank indicator will be in which of the following positions?

A. The center B. Always to the left, showing a slip C. Always to the right, showing a skid D. Either B or C above, depending on the direction of the turn

6-35. As an aircraft turns, the indicator needle moves in the direction of the turn for which of the following reasons?

A. Centrifugal force and gyro precession are in opposition to each other B. Gravity is greater than centrifugal force in any aircraft attitude other than straight-and-level flight C. The needle is connected to the frame in such a manner to cause the two to move in opposition D. The centrifugal force applied to the needle is in the direction opposite to the centrifugal force applied to the frame

6-36. What instrument does the pilot use to reduce the possibility of damage to the aircraft from excessive stress?

A. The turn-and-bank indicator B. The accelerometer C. The AOA indicator D. The vertical gyro indicator

6-37. The standard Navy aircraft clock has what type of movement?

A. 12-hour, 8-day B. 24-hour, 8-day C. 12-hour, 12-day D. 24-hour, 12-day

6-38. What is the purpose of filling the bowl of an aircraft direct-reading compass with a liquid?

A. To keep the surface of the compass card smooth B. To slow the movement of the compass card C. To compensate for pressure changes D. To magnify the compass card

ASSIGNMENT Chapter 7 COMPASS AND INERTIAL NAVIGATION SYSTEMS 7-1. Planes that pass through the earth perpendicular to the earth’s rotational axis and intersect with the earth’s surface to form circles are known as ________.

A. parallels B. perpendiculars C. meridians D. prime meridians

7-2. Relative to the earth’s surface, if an aircraft were at latitude 37° S and longitude 83° E, it would be at which of the following positions?

A. 37° south of Greenwich, England, and 83° east of the equator B. 83° east and 37° south of Greenwich, England C. 83° east of Greenwich, England, and 37° south of the equator D. 37° south and 83° east of the equator

7-3. Convert the following coordinates from decimal form to degree/minutes/seconds form: Latitude--47.7° N Longitude--131.45° E

A. 47°60'4" N and 131°45' E, respectively B. 42°36' N and 48°36' E, respectively C. 47°42' N and 131°27' E, respectively D. 48°15’ N and 131°75’ E, respectively

7-4. What is an irregular line connecting points on a map of the earth, indicating where a compass points to true north?

A. Variation B. Agonic line C. Isogonic line D. Deviation

7-5. What is the angular difference between the directions of true north and magnetic north at a particular location?

A. Variation B. Agonic line C. Isogonic line D. Deviation

7-6. What is the angular difference between the direction of the earth’s magnetic field and the compass reading due to nearby electromagnetic influences?

A. Variation B. Agonic line C. Isogonic line D. Deviation

7-7. If variation is 6° west and deviation is 1° west, the compass error is equal to________.

A. -6° + 1° = 5° east B. -6° - 1 ° = 7° east C. 6° + 1° = 7° west D. 6° - 1° = 5° west

7-8. The difference between the direction of the earth’s magnetic field and the horizontal at any location on the earth’s surface is known as the ________.

A. magnetic dip B. magnetic variation C. surface variation D. surface deviation

7-9. A line on a map that connects all places having equal dip angles is known as an ________.

A. agonic line B. aclinic line C. isobaric line D. isoclinic line

7-10. The aircraft navigator is plotting the present position by using aircraft course and speed, last known position, elapsed time, and any changes in speed and course since the last known position. What type of navigation is the navigator using?

A. Mapping B. Pilotage C. Inertial D. Dead reckoning

7-11. In the compass system, what unit senses the direction of the flux lines of the earth’s magnetic field?

A. The gyro amplifier B. The flux valve C. The magnetic amplifier D. The direct-reading compass

7-12. A displacement gyro provides which of the following electrical signals?

A. Azimuth and pitch only B. Pitch and roll only C. Roll and azimuth only D. Azimuth, pitch, and roll

7-13. The directional gyro pitch gimbal is maintained perpendicular to the surface of the earth by a motor-generator that is driven by the amplified output of ________.

A. microswitches B. a generator dampened by a pendulum-type weight C. the vertical gyro’s pitch servo control transmitter D. electrolytic switches mounted on the directional gyro’s pitch gimbal

7-14. The two basic categories of navigation are known as ________.

A. dead reckoning and compass B. position fixing and celestial C. position fixing and dead reckoning D. celestial and inertial

7-15. Dead reckoning is the process by which position is calculated from what known data?

A. Course, speed, elapsed time, and previous position B. Course, elapsed time, and previous position C. Course, speed, and elapsed time D. Course, speed, and previous position

7-16. The inertial navigation system is a unique navigation system for which of the following reasons?

A. It produces its own electrical power. B. It relies on information external to the vehicle. C. It is independent of its operating environment. D. It beams a laser signal to navigational satellites.

7-17. Every body continues its state of rest, or of uniform motion in a straight line, unless it is compelled to change that state by forces impressed upon it. This statement defines which of the following laws?

A. Boyle’s law B. Charles’ law C. Newton’s second law of motion D. Newton’s first law of motion

7-18. What quantity does an INS derive when acceleration is integrated twice over a specific period of time?

A. Displacement B. Acceleration C. Position D. Velocity

7-19. The mathematical process of summing all minute values of a variable function over a given time is known as ________.

A. Integration of acceleration B. displacement of acceleration C. differentiation of acceleration D. value of acceleration

7-20. All compass systems must be periodically ________.

A. charged B. replaced C. calibrated D. lubricated

7-21. A minimum of how many satellites are observable from anywhere on earth?

A. Two B. Three C. Four D. Five

7-22. Global Positioning System (GPS) is funded by and controlled by whom?

A. U. S. Air Force B. U. S. Air and Space Command C. U. S. Department of Defense D. U. S. Space and Warfare Command

7-23. The difference between the direction of the earth’s magnetic field and the horizontal at any location is known as what?

A. Magnetic poles B. Magnetic dip C. Isoclinic line D. Magnetic equator

7-24. What measures the linear acceleration of the inertial reference that only is measured relative to a moving system?

A. X-axis B. Integrator C. Gyro D. Accelerometers

7-25. What component below functions as a selectable display and is used to provide visual information of steering and navigation?

A. Horizontal Situation Indicator B. Tactical aircraft moving map C. Digital map computer D. PCMCIA card

7-26. GPS aircraft position data is not susceptible to local atmospheric pressure variations or other environmental effects but can be affected by what types of atmospheric delays?

A. Mesospheric and Tropospheric delay B. Tropospheric and Ionospheric delay C. Tropospheric and Stratospheric delay D. Ionspheric and Stratospheric delay

7-27. Global Positioning System is a ________ system?

A. one-way (listen only) B. two-way (worldwide) C. three-way (receiving, sending and correcting) D. four satellite

ASSIGNMENT Chapter 8 AUTOMATIC FLIGHT CONTROL AND STABILIZATION SYSTEMS 8-1. Which of the following aircraft components are examples of airfoils?

A. Engines B. Propellers C. Vertical stabilizers D. Wheels

8-2. By what means does an airfoil produce lift?

A. Decreasing drag B. Decreasing the angle of attack C. Creating low pressure on the rounded surface D. Creating high pressure on the rounded surface

8-3. A movable device attached to the trailing edge of an airfoil increases lift by ________.

A. increasing airspeed. B. increasing the angle of attack. C. decreasing drag. D. decreasing the angle of attack.

8-4. If the control stick of a fixed-wing aircraft were moved aft, what would be the result?

A. The nose of the aircraft would move down B. The elevators would move up C. The elevators would move down D. The ailerons would move down

8-5. When banking an aircraft, what flight controls must be coordinated?

A. Elevators and ailerons only B. Rudder and ailerons only C. Elevators and rudder only D. Elevators, ailerons, and rudder

8-6. To compensate for an aircraft’s continuous right yaw condition, the pilot should take which of the following actions?

A. Push the left rudder pedal and hold B. Move the rudder trim tab to the left C. Push the right rudder pedal and hold D. Move the rudder trim tab to the right

8-7. To return the aircraft to laterally level flight from a right bank, what aileron control is necessary?

A. The left aileron must go down, and the right aileron must go up. B. The left aileron must go up, and the right aileron must go down. C. Both ailerons must go up. D. Both ailerons must go down.

8-8. The movement of the aircraft nose in the opposite direction of an intended turn is caused by aileron drag and is known as ________.

A. adverse yaw B. ballooning C. slip D. skid

8-9. A major advantage of a helicopter over fixed-wing aircraft is that a helicopter can ________.

A. fly at low altitudes B. fly at zero or very low airspeed C. make sharp turns D. climb at high rates of speed

8-10. Operation of the collective control causes a change in the pitch angle of the ________.

A. rudder B. ailerons C. main rotor blades D. tail rotor blades

8-11. In hovering flight only, helicopter heading is controlled by which of the following controls?

A. Cyclic stick only B. Collective stick only C. Cyclic stick and collective stick D. Rudder pedals

8-12. Pitch and directional control of a rotary-wing aircraft are accomplished by the use of the ________.

A. ailerons B. rudder C. cyclic stick D. collective stick

8-13. What is the function of the amplifiers and computers in the AFCS?

A. To determine how much and in which direction correction is necessary B. To provide a reference for standard conditions C. To produce a reference for nonstandard conditions D. To determine the magnitude and direction of the correction

8-14. The pilot controls the operator modes of the AFCS through the use of ________.

A. a control panel B. a control valve C. a control surface D. an air navigation computer

8-15. In helicopter flight (except hovering flight), the main rotor provides altitude, bank, and directional control through use of the collective and ________ controls.

A. torque B. cyclic C. engine speed D. stator

8-16. During forward flight, blade pitch is greatest as it passes which position?

A. 0 degrees B. 90 degrees C. 180 degrees D. 270 degrees

8-17. The vane assembly of the dynamic vertical sensor consists of a viscous-damped pendulum mechanically connected to the rotor shaft. The damping effect of the fluid gives long term sensing characteristic that makes the unit insensitive to?

A. Transient oscillations B. Torque C. Displacement angle D. Lateral force

8-18. Where does the AFCS receive control signals from when altitude hold is engaged?

A. Air navigation computer B. Mission computer C. Flight control computer D. Air data computer