AVIATION ELECTRICIAN'S MATE (AE)

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.

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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.

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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.

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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.

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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.

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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.

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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).

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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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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.

7-15

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.

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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

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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.

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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.

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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.

7-71

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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7-80

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.

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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.

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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.

8-46

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)

8-47

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

8-48

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

8-50

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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

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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