Aircraft Airframe Part 2

indicating systems, the shaft to which the contact arm is fastened protrudes through the end of transmitter housing and is attached to the unit whose position is to be transmitted (e.g., flaps, landing gear). The transmitter is often connected to the moving unit through a mechanical linkage. As the unit moves, it causes the transmitter shaft to turn. The arm is turned so that voltage is applied through the brushes to any two points around the circumference of the resistance winding. The rotor shaft of DC selsyn systems, measuring other kinds of data, operates the same way, but may not protrude outside of the housing. The sensing device, which imparts rotary motion to the shaft, could be located inside the transmitter housing. Referring to Figure 10-47, note that the resistance winding of the transmitter is tapped off in three fixed places, usually 120° apart. These taps distribute current through the toroidial windings of the indicator motor. When current flows through these windings, a magnetic field is created. Like all magnetic fields, a definite north and south direction to the field exists. As the transmitter rotor shaft is turned, the voltage-supplying contact arm moves. Because it contacts the transmitter resistance winding in different positions, the resistance between the supply arm and the various tapoffs changes. This causes the voltage flowing through the tapoffs to change as the resistance of sections of the winding become longer or shorter. The result is that varied current is sent via the tapoffs to the three windings in the indicator motor. The resultant magnetic field created by current flowing through the indicator coils changes as each receives varied current from the tapoffs. The direction of the magnetic field also changes. Thus, the direction of the magnetic field across the indicating element corresponds in position to the moving arm in the transmitter. A permanent magnet is attached to the centered rotor shaft in the indicator, as is the indicator pointer. The magnet aligns itself with the direction of the magnetic field and the pointer does as well. Whenever the magnetic field changes direction, the permanent magnet and pointer realign with the new position of the field. Thus, the position of the aircraft device is indicated. Landing gear contain mechanical devices that lock the gear up, called an up-lock, or down, called a down-lock. When the DC selsyn system is used to indicate the position of the landing gear, the indicator can also show that the up-lock or down-lock is engaged. This is done by again varying the current flowing through the indicator’s coils. Switches located on the actual locking devices close when the locks engage. Current from the selsyn system described above flows through the switch and a small additional circuit. The circuit adds an additional resistor to one of the transmitter winding sections created by the rotor arm and a tapoff. This changes the total resistance of that section. The result is a change in the current flowing through one of the indicator’s motor coils. This, in turn, changes the magnetic field around that coil. Therefore, the combined magnetic field created by all three motor coils is also affected, causing a shift in the direction of the indicator’s magnetic field. The permanent magnet and pointer align with the new direction and shift to the locked position on the indicator dial. Figure 10-48 shows a simplified diagram of a lock switch in a three-wire selsyn system and an indicator dial. 10-28 26 volts 400 Hz power supply A B D C N S A B D C N S Up Down Transmitting magnesyn Indicating magnesyn Soft iron core Toroidal winding Permanent magnet 1/3 1/3 1/3 Figure 10-49. A magnasysn synchro remote-indicating system uses AC. It has permanent magnet rotors in the transmitter and indictor. Transmitter Indicator 26V. 400 Hz AC Electromagnetic rotor Stator windings Indicator pointer Figure 10-50. An autosyn remote-indicating system utilizes the interaction between magnetic fields set up by electric current flow to position the indicator pointer. AC Synchro Systems Aircraft with alternating current (AC) electrical power systems make use of autosyn or magnasysn synchro remote indicating systems. Both operate in a similar way to the DC selsyn system, except that AC power is used. Thus, they make use of electric induction, rather than resistance current flows defined by the rotor brushes. Magnasyn systems use permanent magnet rotors such a those found in the DC selsyn system. Usually, the transmitter magnet is larger than the indicator magnet, but the electromagnetic response of the indicator rotor magnet and pointer remains the same. It aligns with the magnetic field set up by the coils, adopting the same angle of deflection as the transmitter rotor. [Figure 10-49] Autosyn systems are further distinguished by the fact that the transmitter and indicator rotors used are electro-magnets rather than permanent magnets. Nonetheless, like a permanent magnet, an electro-magnet aligns with the direction of the magnetic field created by current flowing through the stator coils in the indicator. Thus, the indicator pointer position mirrors the transmitter rotor position. [Figure 10-50] AC synchro systems are wired differently than DC systems. The varying current flows through the transmitter and indicator stator coils are induced as the AC cycles through zero and the rotor magnetic field flux is allowed to flow. The important characteristic of all synchro systems is maintained by both the autosyn and magnasyn systems. That is, the position of the transmitter rotor is mirrored by the rotor in the indicator. These systems are used in many of the same applications as the DC systems and more. Since they are usually part of instrumentation for high performance aircraft, adaptations of autosyn and magnasyn synchro systems are frequently used in directional indicators and in autopilot systems. Remote Indicating Fuel and Oil Pressure Gauges Fuel and oil pressure indications can be conveniently obtained through the use of synchro systems. As stated previously, running fuel and oil lines into the cabin to direct reading gauges is not desirable. Increased risk of fire in the cabin and the additional weight of the lines are two primary deterrents. By locating the transmitter of a synchro system remotely, fluid pressure can be directed into it without a long tubing run. Inside the transmitter, the motion of a pressure bellows can be geared to the transmitter rotor in such a way as to make the rotor turn. [Figure 10-51] As in all synchros, the transmitter rotor turns proportional to the pressure sensed, which varies the voltages set up in the resistor windings of the synchro stator. These voltages are transmitted to the indicator coils that develop the magnetic field that positions the pointer. Often on twin-engine aircraft, synchro mechanisms for each engine can be used to drive separate pointers on the same indicator. By placing the coils one behind the other, the pointer shaft from the rear indicator motor can be sent through the hollow shaft of the forward indicator motor. Thus, each pointer responds with the magnet’s alignment in its own motor’s magnetic field while sharing the same gauge housing. Labeling the pointer’s engine 1 or 2 removes any doubt about which indicator pointer is being observed. A similar principle is employed in an indicator that has side-by-side indications for different parameters, such as oil pressure and fuel pressure in the same indicator housing. Each parameter has its own synchro motor for positioning its pointer. 10-29 D C B A D C A B Indicator pointer Diaphragm Vent Rotor Rotor Stator Stator Pressure connector AC power Open to atmosphere To engine oil pump Oil pressure indicator (instrument panel) Oil pressure transmitter (engine) Figure 10-51. Remote pressure sensing indicators change linear motion to rotary motion in the sensing mechanism part of the synchro transmitter. Aircraft with digital instrumentation make use of pressuresensitive solid-state sensors that output digital signals for collection and processing by dedicated engine and airframe computers. Others may retain their analog sensors, but may forward this information through an analog to digital converter unit from which the appropriate computer can obtain digital information to process and illuminate the digital display. Many more instruments utilize the synchro remote-indicating systems described in this section or similar synchros. Sometimes simple, more suitable, or less expensive technologies are also employed. Mechanical Movement Indicators There are many instruments on an aircraft that indicate the mechanical motion of a component, or even the aircraft itself. Some utilize the synchro remote-sensing and indicating systems described above. Other means for capturing and displaying mechanical movement information are also used. This section discusses some unique mechanical motion indicators and groups instruments by function. All give valuable feedback to the pilot on the condition of the aircraft in flight. Tachometers The tachometer, or tach, is an instrument that indicates the speed of the crankshaft of a reciprocating engine. It can be a direct- or remote-indicating instrument, the dial of which is calibrated to indicate revolutions per minutes (rpm). On reciprocating engines, the tach is used to monitor engine power and to ensure the engine is operated within certified limits. Gas turbine engines also have tachometers. They are used to monitor the speed(s) of the compressor section(s) of the engine. Turbine engine tachometers are calibrated in percentage of rpm with 100 percent corresponding to optimum turbine speed. This allows similar operating procedures despite the varied actual engine rpm of different engines. [Figure 10-52] In addition to the engine tachometer, helicopters use a tachometer to indicator main rotor shaft rpm. It should also be noted that many reciprocating-engine tachometers also have built-in numeric drums that are geared to the rotational mechanism inside. These are hour meters that keep track of the time the engine is operated. There are two types of tachometer system in wide use today: mechanical and electrical. Mechanical Tachometers Mechanical tachometer indicating systems are found on small, single-engine light aircraft in which a short distance exists between the engine and the instrument panel. They consist of an indicator connected to the engine by a flexible drive shaft. The drive shaft is geared into the engine so that when the engine turns, so does the shaft. The indicator contains a flyweight assembly coupled to a gear mechanism that drives a pointer. As the drive shaft rotates, centrifugal force acts on the flyweights and moves them to an angular position. This angular position varies with the rpm of the engine. The amount of movement of the flyweights is transmitted through the gear mechanism to the pointer. The pointer rotates to indicate this movement on the tachometer indicator, which is directly related to the rpm of the engine. [Figure 10-53] 10-30 Flyweight Coil spring Rocking shaft Drive shaft Figure 10-53. The simplified mechanism of a flyweight type mechanical tachometer. 5 3 10 25 20 35 30 15 X100 HOURS 10 RPM 0 0 0 0 0 0 10 20 30 40 50 60 70 80 90 100 PERCENT RPM 0 5 9 1 8 7 6 4 5 5 Figure 10-52. A tachometer for a reciprocating engine is calibrated in rpm. A tachometer for a turbine engine is calculated in percent of rpm. A more common variation of this type of mechanical tachometer uses a magnetic drag cup to move the pointer in the indicator. As the drive shaft turns, it rotates a permanent magnet in a close-tolerance aluminum cup. A shaft attached to the indicating point is attached to the exterior center of the cup. As the magnet is rotated by the engine flex drive cable, its magnetic field cuts through the conductor surrounding it, creating eddy currents in the aluminum cup. This current flow creates its own magnetic field, which interacts with the rotating magnet’s flux field. The result is that the cup tends to rotate, and with it, the indicating pointer. A calibrated restraining spring limits the cup’s rotation to the arc of motion of the pointer across the scale on the instrument face. [Figure 10-54] Electric Tachometers It is not practical to use a mechanical linkage between the engine and the rpm indicator on aircraft with engines not mounted in the fuselage just forward of the instrument panel. Greater accuracy with lower maintenance is achieved through the use of electric tachometers. A wide variety of electric tachometer systems can be employed, so manufacturer’s 10-31 N S Dial Pointer Hairspring Drive cable from engine Drag cup Rotating magnet Figure 10-54. A simplified magnetic drag cup tachometer indicating device. A. Pad type B. Screw type Figure 10-56. Different types of tach generators. Drive coupling Indicating needle Drag cup Spring Permanent magnet Stator coil Permanent magnet rotor Tachometer generator Indicator Permanent magnet rotor Alternator Synchronous motor Figure 10-55. An electric tachometer system with synchronous motors and a drag cup indicator. instructions should be consulted for details of each specific tachometer system. A popular electric tachometer system makes use of a small AC generator mounted to a reciprocating engine’s gear case or the accessory drive section of a turbine engine. As the engine turns, so does the generator. The frequency output of the generator is directly proportional to the speed of the engine. It is connected via wires to a synchronous motor in the indicator that mirrors this output. A drag cup, or drag disk link, is used to drive the indicator as in a mechanical tachometer. [Figure 10-55] Two different types of generator units, distinguished by their type of mounting system, are shown in Figure 10-56. The dual tachometer consists of two tachometer indicator units housed in a single case. The indicator pointers show simultaneously, on one or two scales, the rpm of two engines. A dual tachometer on a helicopter often shows the rpm of the engine and the rpm of the main rotor. A comparison of the voltages produced by the two tach generators of this type of helicopter indicator gives information concerning clutch slippage. A third indication showing this slippage is sometimes included in the helicopter tachometer. [Figure 10-57] Some turbine engines use tachometer probes for rpm indication, rather than a tach generator system. They provide a great advantage in that there are no moving parts. They are sealed units that are mounted on a flange and protrude into the compressor section of the engine. A magnetic field is set up inside the probe that extends through pole pieces and out the end of the probe. A rotating gear wheel, which moves at the same speed as the engine compressor shaft, alters the magnetic field flux density as it moves past the pole pieces at close proximity. This generates voltage signals in coils inside the probe. The amplitude of the EMF signals vary directly with the speed of the engine. The tachometer probe’s output signals need to be processed in a remotely located module. They must also be amplified to drive a servo motor type indicator in the cockpit. They may 10-32 Pole piece Electrical connector Gear wheel Tacho probe Axis of polarization Permanent magnet Sensing coils Pole pieces Figure 10-58. A tacho probe has no moving parts. The rate of magnetic flux field density change is directly related to engine speed. Synchroscope SLOW FAST Figure 10-59. This synchroscope indicates the relative speed of the slave engine to the master. ENGINE RPM ROTOR RPM 10 15 20 25 30 5 0 25 30 5 10 40 20 15 35 2 E T Figure 10-57. A helicopter tachometer with engine rpm, rotor rpm, and slippage indications. also be used as input for an automatic power control system or a flight data acquisition system. [Figure 10-58] Synchroscope The synchroscope is an instrument that indicates whether two or more rotating devices, such as engines, are synchronized. Since synchroscopes compare rpm, they utilize the output from tachometer generators. The instrument consists of a small electric motor that receives electrical current from the generators of both engines. Current from the faster running engine controls the direction in which the synchroscope motor rotates. If both engines are operating at exactly the same speed, the synchroscope motor does not operate. If one engine operates faster that the other, its tach generator signal causes the synchroscope motor to turn in a given direction. Should the speed of the other engine then become greater than that of the first engine, the signal from its tach generator causes the synchroscope motor to reverse itself and turn in the opposite direction. The pilot makes adjustments to steady the pointer so it does not move. One use of synchroscope involve designating one of the engines as a master engine. The rpm of the other engine(s) is always compared to the rpm of this master engine. The dial face of the synchroscope indicator looks like Figure 10-59. “Slow” and “fast” represent the other engine’s rpm relative to the master engine, and the pilot makes adjustments accordingly. Accelerometers An accelerometer is an instrument that measures acceleration. It is used to monitor the forces acting upon an airframe. Accelerometers are also used in inertial reference navigation systems. The installation of accelerometers is usually limited to high-performance and aerobatic aircraft. Simple accelerometers are mechanical, direct-reading instruments calibrated to indicate force in Gs. One G is equal to one times the force of gravity. The dial face of an 10-33 4 2 -2 PUSH TO SET Sheave pulley (top) Mass shafts Control cord Mass Main pulley Main pointer centering spring Driver arm Sheave pulley (bottom) H T S Pawl Pointer reset shaft Auxiliary pointer return spring Ratchet Main pointer Auxiliary pointer (plus G indication) Auxiliary pointer (minus G indication) Figure 10-60. The inner workings of a mass-type accelerometer. accelerometer is scaled to show positive and negative forces. When an aircraft initiates a rapid climb, positive G force tends to push one back into one’s seat. Initiating a rapid decent causes a force in the opposite direction, resulting in a negative G force. Most accelerometers have three pointers. One is continuously indicating the acceleration force experienced. The other two contain ratcheting devices. The positive G pointer follows the continuous pointer and stay at the location on the dial where the maximum positive force is indicated. The negative G pointer does the same for negative forces experienced. Both max force pointers can be reset with a knob on the instrument face. The accelerometer operates on the principle of inertia. A mass, or weight, inside is free to slide along a shaft in response to the slightest acceleration force. When a maneuver creates an accelerating force, the aircraft and instrument move, but inertia causes the weight to stay at rest in space. As the shaft slides through the weight, the relative position of the weight on the shaft changes. This position corresponds to the force experienced. Through a series of pulleys, springs, and shafts, the pointers are moved on the dial to indicate the relative strength of the acceleration force. [Figure 10-60] Forces can act upon an airframe along the three axes of flight. Single and multi-axis accelerometers are available, although most cockpit gauges are of the single-axis type. Inertial reference navigation systems make use of multi-axis accelerometers to continuously, mathematically calculate the location of the aircraft in a three dimensional plane. Electric and digital accelerometers also exist. Solid-state sensors are employed, such as piezoelectric crystalline devices. In these instruments, when an accelerating force is applied, the amount of resistance, current flow, or capacitance changes in direct relationship to the size of the force. Microelectric signals integrate well with digital computers designed to process and display information in the cockpit. 10-34 Figure 10-61. A reed-type stall warning device is located behind this opening in the leading edge of the wing. When the angle of attack increases to near the point of a stall, low-pressure air flowing over the opening causes a suction, which audibly vibrates the reed. Switch tab held down High angle of attack near stall point Low/normal angle of attack Direction of travel Relative wind Relative wind Point of stagnation Switch tab pushed up Point of stagnation Figure 10-62. A popular stall warning switch located in the wing leading edge. Stall Warning and Angle of Attack (AOA) Indicators An aircraft’s angle of attack (AOA) is the angle formed between the wing cord centerline and the relative wind. At a certain angle, airflow over the wing surfaces is insufficient to create enough lift to keep the aircraft flying, and a stall occurs. An instrument that monitors the AOA allows the pilot to avoid such a condition. The simplest form of AOA indicator is a stall warning device that does not have a gauge located in the cockpit. It uses an aural tone to warn of an impending stall due to an increase in AOA. This is done by placing a reed in a cavity just aft of the leading edge of the wing. The cavity has an open passage to a precise point on the leading edge. In flight, air flows over and under a wing. The point on the wing leading edge where the oncoming air diverges is known as the point of stagnation. As the AOA of the wing increases, the point of stagnation moves down below the open passage that leads inside the wing to the reed. Air flowing over the curved leading edge speeds up and causes a low pressure. This causes air to be sucked out of the inside of the wing through the passage. The reed vibrates as the air rushes by making a sound audible in the cockpit. [Figure 10-61] Another common device makes use of an audible tone as the AOA increases to near the point where the aircraft will stall. This stall warning device includes an electric switch that opens and closes a circuit to a warning horn audible in the cockpit. It may also be wired into a warning light circuit. The switch is located near the point of stagnation on the wing leading edge. A small lightly sprung tab activates the switch. At normal AOA, the tab is held down by air that diverges at the point of stagnation and flows under the wing. This holds the switch open so the horn does not sound nor the warning light illuminate. As the AOA increases, the point of stagnation moves down. The divergent air that flows up and over the wing now pushes the tab upward to close the switch and complete the circuit to the horn or light. [Figure 10-62] A true AOA indicating system detects the local AOA of the aircraft and displays the information on a cockpit indicator. It also may be designed to furnish reference information to other systems on high-performance aircraft. The sensing 10-35 0 10 30 ANGLE OF ATTACK O F F Figure 10-63. Angle of attack indicator. Air flow Air flow Probe type Probe Flush mounted in side of forward fuselage Vane Air holes Figure 10-64. A slotted AOA probe and an alpha vane. Paddle chamber Probe Paddle Separator Potentiometer Air passages e Slots Orifice Figure 10-65. The internal structure of a slotted probe airstream direction detector. mechanism and transmitter are usually located on the forward side of the fuselage. It typically contains a heating element to ensure ice-free operation. Signals are sent from the sensor to the cockpit or computer(s) as required. An AOA indicator may be calibrated in actual angle degrees, arbitrary units, percentage of lift used, symbols, or even fast/ slow. [Figure 10-63] There are two main types of AOA sensors in common use. Both detect the angular difference between the relative wind and the fuselage, which is used as a reference plane. One uses a vane, known as an alpha vane, externally mounted to the outside of the fuselage. It is free to rotate in the wind. As the AOA changes, air flowing over the vane changes its angle. The other uses two slots in a probe that extends out of the side of the fuselage into the airflow. The slots lead to different sides of movable paddles in a chamber of the unit just inside the fuselage skin. As the AOA varies, the air pressure ported by each of the slots changes and the paddles rotate to neutralize the pressures. The shaft upon which the paddles rotate connects to a potentiometer wiper contact that is part of the unit. The same is true of the shaft of the alpha vane. The changing resistance of the potentiometer is used in a balanced bridge circuit to signal a motor in the indicator to move the pointer proportional to the AOA. [Figures 10-64 and 10-65] Modern aircraft AOA sensor units send output signals to the ADC. There, the AOA data is used to create an AOA indication, usually on the primary flight display. AOA information can also be integrated with flap and slat position information to better determine the point of stall. Additionally, AOA sensors of the type described are subject to position error since airflow around the alpha vane and slotted probe changes somewhat with airspeed and aircraft attitude. The errors are small, but can be corrected in the ADC. To incorporate a warning of an impending stall, many AOA systems signal a stick shaker motor that literally shakes the control column to warn the pilot as the aircraft approaches a stall condition. Electrical switches are actuated in the AOA indicator at various preset AOA to activate the motor that drives an unbalanced weighted ring, causing the column to shake. Some systems include a stick pusher actuator that pushes the control yoke forward, lowering the nose of the aircraft when the critical AOA is approached. Regardless of the many existing variations for warning of an impending stall, the AOA system triggers all stall warnings in high performance aircraft. Temperature Measuring Instruments The temperature of numerous items must be known for an aircraft to be operated properly. Engine oil, carburetor mixture, inlet air, free air, engine cylinder heads, heater ducts, and exhaust gas temperature of turbine engines are all items requiring temperature monitoring. Many other temperatures must also be known. Different types of thermometers are used to collect and present temperature information. 10-36 20 0 -20 -40 -60 40 60 100 140 120 80 F 20 0 -20 -40 40 60 C −40 −30 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110 120 Bimetallic temperature gauge Bimetallic coil of bonded metals with dissimilar coefficients of expansion Figure 10-66. A bimetallic temperature gauge works because of the dissimilar coefficients of expansion of two metals bonded together. When bent into a coil, cooling or heating causes the dissimilar metal coil to tighten, or unwind, moving the pointer across the temperature scale on the instrument dial face. Non-Electric Temperature Indicators The physical characteristics of most materials change when exposed to changes in temperature. The changes are consistent, such as the expansion or contraction of solids, liquids, and gases. The coefficient of expansion of different materials varies and it is unique to each material. Most everyone is familiar with the liquid mercury thermometer. As the temperature of the mercury increases, it expands up a narrow passage that has a graduated scale upon it to read the temperature associated with that expansion. The mercury thermometer has no application in aviation. A bimetallic thermometer is very useful in aviation. The temperature sensing element of a bimetallic thermometer is made of two dissimilar metals strips bonded together. Each metal expands and contracts at a different rate when temperature changes. One end of the bimetallic strip is fixed, the other end is coiled. A pointer is attached to the coiled end which is set in the instrument housing. When the bimetallic strip is heated, the two metals expand. Since their expansion rates differ and they are attached to each other, the effect is that the coiled end tries to uncoil as the one metal expands faster than the other. This moves the pointer across the dial face of the instrument. When the temperature drops, the metals contract at different rates, which tends to tighten the coil and move the pointer in the opposite direction. Direct reading bimetallic temperature gauges are often used in light aircraft to measure free air temperature or outside air temperature (OAT). In this application, a collecting probe protrudes through the windshield of the aircraft to be exposed to the atmospheric air. The coiled end of the bimetallic strip in the instrument head is just inside the windshield where it can be read by the pilot. [Figures 10-66 and 10-67] A bourdon tube is also used as a direct reading non-electric temperature gauge in simple, light aircraft. By calibrating the dial face of a bourdon tube gauge with a temperature scale, it can indicate temperature. The basis for operation is the consistent expansion of the vapor produced by a volatile liquid in an enclosed area. This vapor pressure changes directly with temperature. By filling a sensing bulb with such a volatile liquid and connecting it to a bourdon tube, the tube causes an indication of the rising and falling vapor pressure due to temperature change. Calibration of the dial face in degrees Fahrenheit or Celsius, rather than psi, provides a temperature reading. In this type of gauge, the sensing bulb is placed in the area needing to have temperature measured. A long capillary tube connects the bulb to the bourdon tube in the instrument housing. The narrow diameter of the capillary tube ensures that the volatile liquid is lightweight and stays primarily in the sensor bulb. Oil temperature is sometimes measured this way. Electrical Temperature Measuring Indication The use of electricity in measuring temperature is very common in aviation. The following measuring and indication systems can be found on many types of aircraft. Certain temperature ranges are more suitably measured by one or another type of system. Electrical Resistance Thermometer The principle parts of the electrical resistance thermometer are the indicating instrument, the temperature-sensitive element (or bulb), and the connecting wires and plug connectors. Electrical resistance thermometers are used widely in many types of aircraft to measure carburetor air, oil, free air temperatures, and more. They are used to measure low and medium temperatures in the –70 °C to 150 °C range. 10-37 Figure 10-67. A bimetallic outside air temperature gauge and its installation on a light aircraft. Figure 10-68. An electric resistance thermometer sensing bulb. + − Heat-sensitive element or bulb 14Volts Indicator A B L M X C Cal. res. Y D Figure 10-69. The internal structure of an electric resistance thermometer indicator features a bridge circuit, galvanometer, and variable resistor, which is outside the indicator in the form of the temperature sensor. For most metals, electrical resistance changes as the temperature of the metal changes. This is the principle upon which a resistance thermometer operates. Typically, the electrical resistance of a metal increases as the temperature rises. Various alloys have a high temperature-resistance coefficient, meaning their resistance varies significantly with temperature. This can make them suitable for use in temperature sensing devices. The metal resistor is subjected to the fluid or area in which temperature needs to be measured. It is connected by wires to a resistance measuring device inside the cockpit indicator. The instrument dial is calibrated in degrees Fahrenheit or Celsius as desired rather than in ohms. As the temperature to be measured changes, the resistance of the metal changes and the resistance measuring indicator shows to what extent. A typical electrical resistance thermometer looks like any other temperature gauge. Indicators are available in dual form for use in multiengine aircraft. Most indicators are self-compensating for changes in cockpit temperature. The heat-sensitive resistor is manufactured so that it has a definite resistance for each temperature value within its working range. The temperature-sensitive resistor element is a length or winding made of a nickel/manganese wire or other suitable alloy in an insulating material. The resistor is protected by a closed-end metal tube attached to a threaded plug with a hexagonal head. [Figure10-68] The two ends of the winding are brazed, or welded, to an electrical receptacle designed to receive the prongs of the connector plug. The indicator contains a resistance-measuring instrument. Sometimes it uses a modified form of the Wheatstonebridge circuit. The Wheatstone-bridge meter operates on the principle of balancing one unknown resistor against other known resistances. A simplified form of a Wheatstonebridge circuit is shown in Figure 10-69. Three equal values of resistance [Figure 10-69A, B, and C] are connected into a diamond shaped bridge circuit. A resistor with an unknown value [Figure 10-69D] is also part of the circuit. The unknown resistance represents the resistance of the temperature bulb of the electrical resistance thermometer system. A galvanometer is attached across the circuit at points X and Y. 10-38 A B N S N S N S R Sensitive element (bulb) Figure 10-70. A ratiometer temperature measuring indicator has two coils. As the sensor bulb resistance varies with temperature, different amounts of current flow through the coils. This produces varying magnetic fields. These fields interact with the magnetic field of a large permanent magnet, resulting in an indication of temperature. When the temperature causes the resistance of the bulb to equal that of the other resistances, no potential difference exists between points X and Y in the circuit. Therefore, no current flows in the galvanometer leg of the circuit. If the temperature of the bulb changes, its resistance also changes, and the bridge becomes unbalanced, causing current to flow through the galvanometer in one direction or the other. The galvanometer pointer is actually the temperature gauge pointer. As it moves against the dial face calibrated in degrees, it indicates temperature. Many indicators are provided with a zero adjustment screw on the face of the instrument. This adjusts the zeroing spring tension of the pointer when the bridge is at the balance point (the position at which the bridge circuit is balanced and no current flows through the meter). Ratiometer Electrical Resistance Thermometers Another way of indicating temperature when employing an electric resistance thermometer is by using a ratiometer. The Wheatstone-bridge indicator is subject to errors from line voltage fluctuation. The ratiometer is more stable and can deliver higher accuracy. As its name suggests, the ratiometer electrical resistance thermometer measures a ratio of current flows. The resistance bulb sensing portion of the ratiometer electric resistance thermometer is essentially the same as described above. The circuit contains a variable resistance and a fixed resistance to provide the indication. It contains two branches for current flow. Each has a coil mounted on either side of the pointer assembly that is mounted within the magnetic field of a large permanent magnet. Varying current flow through the coils causes different magnetic fields to form, which react with the larger magnetic field of the permanent magnet. This interaction rotates the pointer against the dial face that is calibrated in degrees Fahrenheit or Celsius, giving a temperature indication. [Figure 10-70] The magnetic pole ends of the permanent magnet are closer at the top than they are at the bottom. This causes the magnetic field lines of flux between the poles to be more concentrated at the top. As the two coils produce their magnetic fields, the stronger field interacts and pivots downward into the weaker, less concentrated part of the permanent magnet field, while the weaker coil magnetic field shifts upward toward the more concentrated flux field of the large magnet. This provides a balancing effect that changes but stays in balance as the coil field strengths vary with temperature and the resultant current flowing through the coils. For example, if the resistance of the temperature bulb is equal to the value of the fixed resistance (R), equal values of current flow through the coils. The torques, caused by the magnetic field each coil creates, are the same and cancel any movement in the larger magnetic field. The indicator pointer will be in the vertical position. If the bulb temperature increases, its resistance also increases. This causes the current flow through coil A circuit branch to increase. This creates a stronger magnetic field at coil A than at coil B. Consequently, the torque on coil A increases, and it is pulled downward into the weaker part of the large magnetic field. At the same time, less current flows through the sensor bulb resistor and coil B, causing coil B to form a weaker magnetic field that is pulled upward into the stronger flux area of the permanent magnet’s magnetic field. The pointer stops rotating when the fields reach a new balance point that is directly related to the resistance in the sensing bulb. The opposite of this action would take place if the temperature of the heat-sensitive bulb should decrease. Ratiometer temperature measuring systems are used to measure engine oil, outside air, carburetor air, and other temperatures in many types of aircraft. They are especially in demand to measure temperature conditions where accuracy is important, or large variations of supply voltages are encountered. Thermocouple Temperature Indicators A thermocouple is a circuit or connection of two unlike metals. The metals are touching at two separate junctions. If one of the junctions is heated to a higher temperature than the other, an electromotive force is produced in the circuit. This voltage is directly proportional to the temperature. So, by measuring the amount of electromotive force, temperature can be determined. A voltmeter is placed across the colder of the two junctions of the thermocouple. It is calibrated in 10-39 Back of indicating instrument Connectors Thermocouple leads Black Voltmeter inside forms cold junction Hot junction Constantan (chrome on turbine engine) Copper or iron (alumel on turbine engines) Typical Thermocouple Figure 10-71. Thermocouples combine two unlike metals that cause current flow when heated. degrees Fahrenheit or Celsius, as needed. The hotter the hightemperature junction (hot junction) becomes, the greater the electromotive force produced, and the higher the temperature indication on the meter. [Figure 10-71] Thermocouples are used to measure high temperatures. Two common applications are the measurement of cylinder head temperature (CHT) in reciprocating engines and exhaust gas temperature (EGT) in turbine engines. Thermocouple leads are made from a variety of metals, depending on the maximum temperature to which they are exposed. Iron and constantan, or copper and constantan, are common for CHT measurement. Chromel and alumel are used for turbine EGT thermocouples. The amount of voltage produced by the dissimilar metals when heated is measured in millivolts. Therefore, thermocouple leads are designed to provide a specific amount of resistance in the thermocouple circuit (usually very little). Their material, length, or cross-sectional size cannot be altered without compensation for the change in total resistance that would result. Each lead that makes a connection back to the voltmeter must be made of the same metal as the part of the thermocouple to which it is connected. For example, a copper wire is connected to the copper portion of the hot junction and a constantan wire is connected to the constantan part. The hot junction of a thermocouple varies in shape depending on its application. Two common types are the gasket and the bayonet. In the gasket type, two rings of the dissimilar metals are pressed together to form a gasket that can be installed under a spark plug or cylinder hold down nut. In the bayonet type, the metals come together inside a perforated protective sheath. Bayonet thermocouples fit into a hole or well in a cylinder head. On turbine engines, they are found mounted on the turbine inlet or outlet case and extend through the case into the gas stream. Note that for CHT indication, the cylinder chosen for the thermocouple installation is the one that runs the hottest under most operating conditions. The location of this cylinder varies with different engines. [Figure 10-72] The cold junction of the thermocouple circuit is inside the instrument case. Since the electromotive force set up in the circuit varies with the difference in temperature between the hot and cold junctions, it is necessary to compensate the indicator mechanism for changes in cockpit temperature which affect the cold junction. This is accomplished by using a bimetallic spring connected to the indicator mechanism. This actually works the same as the bimetallic thermometer described previously. When the leads are disconnected from the indicator, the temperature of the cockpit area around the instrument panel can be read on the indicator dial. [Figure 10-73] Numeric LED indictors for CHT are also common in modern aircraft. Turbine Gas Temperature Indicating Systems EGT is a critical variable of turbine engine operation. The EGT indicating system provides a visual temperature indication in the cockpit of the turbine exhaust gases as 10-40 x 100 °C 3 2 1 0 °C 3 1 2 0 50 150 250 350 Figure 10-73. Typical thermocouple temperature indicators. Gasket type thermocouple Engine cylinder spark plug bore A B Figure 10-72. A cylinder head temperature thermocouple with a gasket type hot junction is made to be installed under the spark plug or a cylinder hold down nut of the hottest cylinder (A). A bayonet type thermocouple is installed in a bore in the cylinder wall (B). they leave the turbine unit. In certain turbine engines, the temperature of the exhaust gases is measured at the entrance to the turbine unit. This is referred to as a turbine inlet temperature (TIT) indicating system. Several thermocouples are used to measure EGT or TIT. They are spaced at intervals around the perimeter of the engine turbine casing or exhaust duct. The tiny thermocouple voltages are typically amplified and used to energize a servomotor that drives the indicator pointer. Gearing a digital drum indication off of the pointer motion is common. [Figure 10-74] The EGT indicator shown is a hermetically sealed unit. The instrument’s scale ranges from 0 °C to 1,200 °C, with a vernier dial in the upper right-hand corner and a power off warning flag located in the lower portion of the dial. A TIT indicating system provides a visual indication at the instrument panel of the temperature of gases entering the turbine. Numerous thermocouples can be used with the average voltage representing the TIT. Dual thermocouples exist containing two electrically independent junctions within a single probe. One set of these thermocouples is paralleled to transmit signals to the cockpit indicator. The other set of parallel thermocouples provides temperature signals to engine monitoring and control systems. Each circuit is electrically independent, providing dual system reliability. A schematic for the turbine inlet temperature system for one engine of a four-engine turbine aircraft is shown in Figure 10-75. Circuits for the other three engines are identical to this system. The indicator contains a bridge circuit, a chopper circuit, a two-phase motor to drive the pointer, and a feedback potentiometer. Also included are a voltage reference circuit, an amplifier, a power-off flag, a power supply, and an over temperature warning light. Output of the amplifier energizes the variable field of the two-phase motor that positions the indicator main pointer and a digital indicator. The motor also drives the feedback potentiometer to provide a humming signal to stop the drive motor when the correct pointer position, relative to the temperature signal, has been reached. The voltage reference circuit provides a closely regulated reference voltage in the bridge circuit to preclude error from input voltage variation to the indicator power supply. 10-41 CX 100 1 3 4 2 5 0 6 4 2 0 6 7 8 9 10 11 OFF DA C AMP 115 V.A.C. bus Indicator CR AL Chromel Alumel Turbine outlet circuit breaker Figure 10-74. A typical exhaust gas temperature thermocouple system. 1 1 1 M R Bridge Chopper Power supply Zener voltage reference Amplifier B A FEGD Digital indicator 0 C 1200 Power oft warning flag Overtemp warning light Thermocouples Engine No. 1 Overtemp warning light test switch Eng. 1 Eng. 2 Eng. 3 Eng. 4 Figure 10-75. A typical analog turbine inlet temperature indicating system. 10-42 20 0 -20 -40 -60 40 60 100 140 120 80 F 20 0 -20 -40 40 60 C Total air temp + TAT °C OFF TAT + 85.0 TO 96.1 10 2 6 0 4 8 85.0 96.1 10 2 6 0 4 8 450 EGT 10 2 6 0 4 8 450 10 2 6 0 4 8 N1 TAT - 15 c 50 OIL ORESS 50 105 OIL TEMP 100 97.0 N2 97.0 20 OIL QTY 20 8.4 FF 8.4 1.9 N2 VIB 1.9 FAN °C Mode annunciation Function selector push button Failure flag Balanced bridge indicator Servo driven indicator LCD indicator Full digital display Figure 10-76. Different cockpit TAT displays. The overtemperature warning light in the indicator illuminates when the TIT reaches a predetermined limit. An external test switch is usually installed so that over temperature warning lights for all the engines can be tested at the same time. When the test switch is operated, an overtemperature signal is simulated in each indicator temperature control bridge circuit. Digital cockpit instrumentation systems need not employ resistance-type indicators and adjusted servo-driven thermocouple gauges to provide the pilot with temperature information. Sensor resistance and voltage values are input to the appropriate computer, where they are adjusted, processed, monitored, and output for display on cockpit display panels. They are also sent for use by other computers requiring temperature information for the control and monitoring of various integrated systems. Total Air Temperature Measurement Air temperature is a valuable parameter that many performance monitoring and control variables depend on. During flight, static air temperature changes continuously and accurate measurement presents challenges. Below 0.2 Mach, a simple resistance-type or bimetallic temperature gauge can provide relatively accurate air temperature information. At faster speeds, friction, the air’s compressibility, and boundary layer behavior make accurate temperature capture more complex. Total air temperature (TAT) is the static air temperature plus any rise in temperature caused by the highspeed movement of the aircraft through the air. The increase in temperature is known as ram rise. TAT-sensing probes are constructed specifically to accurately capture this value and transmit signals for cockpit indication, as well as for use in various engine and aircraft systems. Simple TAT systems include a sensor and an indicator with a built-in resistance balance circuit. Air flow through the sensor is designed so that air with the precise temperature impacts a platinum alloy resistance element. The sensor is engineered to capture temperature variations in terms of varying the resistance of the element. When placed in the bridge circuit, the indicator pointer moves in response to the imbalance caused by the variable resistor. More complex systems use signal correction technology and amplified signals sent to a servo motor to adjust the indicator in the cockpit. These systems include closely regulated power supply and failure monitoring. They often use numeric drum type readouts, but can also be sent to an LCD driver to illuminate LCD displays. Many LCD displays are multifunctional, capable of displaying static air temperature and true airspeed. In fully digital systems, the correction signals are input into the ADC. There, they can be manipulated appropriately for cockpit display or for whichever system requires temperature information. [Figure 10-76] TAT sensor/probe design is complicated by the potential of ice forming during icing conditions. Left unheated, a probe may cease to function properly. The inclusion of a heating element threatens accurate data collection. Heating the probe must not affect the resistance of the sensor element. [Figure 10-77] Close attention is paid to airflow and materials conductivity during the design phase. Some TAT sensors channel bleed air through the units to affect the flow of outside air, so that it flows directly onto the platinum sensor without gaining added energy from the probe heater. 10-43 FWD FWD Figure 10-77. Total air temperature (TAT) probes. Magnetic lines of flux North magnetic pole South magnetic pole North geographic pole South geographic pole Geographic pole Magnetic pole Equator Prime Meridian Axis of rotation Equator Figure 10-78. The earth and its magnetic field. Direction Indicating Instruments A myriad of techniques and instruments exist to aid the pilot in navigation of the aircraft. An indication of direction is part of this navigation. While the next chapter deals with communication and navigation, this section discusses some of the magnetic direction indicating instruments. Additionally, a common, reliable gyroscopic direction indicator is discussed in the gyroscopic instrument section of this chapter. Magnetic Compass Having an instrument on board an aircraft that indicates direction can be invaluable to the pilot. In fact, it is a requirement that all certified aircraft have some sort of magnetic direction indicator. The magnetic compass is a direction finding instrument that has been used for navigation for hundreds of years. It is a simple instrument that takes advantage of the earth’s magnetic field. Figure 10-78 shows the earth and the magnetic field that surrounds it. The magnetic north pole is very close to the geographic North Pole of the globe, but they are not the same. An ordinary permanent magnet that is free to do so, aligns itself with the direction of the earth’s magnetic field. Upon this principle, an instrument is constructed that the pilot can reference for directional orientation. Permanent magnets are attached under a float that is mounted on a pivot so it is free to rotate in the horizontal plane. As such, the magnets align with the earth’s magnetic field. A numerical compass card, usually graduated in 5° increments, is constructed around the perimeter of the float. It serves as the instrument dial. The entire assembly is enclosed in a sealed case that is filled with a liquid similar to kerosene. This dampens vibration and oscillation of the moving float assembly and decreases friction. On the front of the case, a glass face allows the numerical compass card to be referenced against a vertical lubber line. The magnetic heading of the aircraft is read by noting the graduation on which the lubber line falls. Thus, direction in any of 360° can be read off the dial as the magnetic float compass card assembly holds its alignment with magnetic north, while the aircraft changes direction. The liquid that fills the compass case expands and contracts as altitude changes and temperature fluctuates. A bellows diaphragm expands and contracts to adjust the volume of the space inside the case so it remains full. [Figure 10-79] There are accuracy issues associated with using a magnetic compass. The main magnets of a compass align not only with the earth’s magnetic field, they actually align with the composite field made up of all magnetic influences around them, meaning local electromagnetic influence from metallic structures near the compass and operation aircraft’s electrical system. This is called magnetic deviation. It causes a magnet’s alignment with the earth’s magnetic field to be altered. Compensating screws are turned, which move small permanent magnets in the compass case to correct for this 10-44 N 33 30 27 N-S E-W Float Filler hole Outer case Compensating magnet Instrument lamp Bellows expansion unit Lubber line Compass card Lens Sensing magnet Compensating screws t s Jewel post Jewel spring Jewel Pivot Figure 10-79. The parts of a typical magnetic compass. 5°W 20°E 15°E 10°E 5°E 0° 10°W 15°W 20°W Add to magnetic heading Subtract from magnetic heading Figure 10-80. Aircraft located along the agonic line have 0° of variation between magnetic north and true north. Locations on and between the isogonic lines require addition or subtraction, as shown, to magnetic indications to arrive at a true geographic direction. magnetic deviation. The two set-screws are on the face of the instrument and are labeled N-S and E-W. They position the small magnets to counterbalance the local magnetic influences acting on the main compass magnets. The process for knowing how to adjust for deviation is known as swinging the compass. It is described in the instrument maintenance pages near the end of this chapter. Magnetic deviation cannot be overlooked. It should never be more than 10°. Using nonferrous mounting screws and shielding or twisting the wire running to the compass illuminating lamp are additional steps taken to keep deviation to a minimum. Another compass error is called magnetic variation. It is caused by the difference in location between the earth’s magnetic poles and the geographic poles. There are only a few places on the planet where a compass pointing to magnetic north is also pointing to geographic North. A line drawn through these locations is called the Agonic line. At all other points, there is some variation between that which a magnetic compass indicates is north and geographic (true) North. Isogonic lines drawn on aeronautical charts indicate points of equal variation. Depending on the location of the aircraft, airmen must add or subtract degrees from the magnetic indication to obtain true geographic location information. [Figure 10-80] The earth’s magnetic field exits the poles vertically and arches around to extend past the equator horizontally or parallel to the earth’s surface. [Figure 10-78] Operating an aircraft near the magnetic poles causes what is known as dip error. The compass magnets pull downward toward the pole, rather than horizontally, as is the case near the equator. This downward motion causes inaccuracy in the indication. Although the compass float mechanism is weighted to compensate, the closer the aircraft is to the north or south magnetic poles, the more pronounced the errors. Dip errors manifest themselves in two ways. The first is called acceleration error. If an aircraft is flying on an east-west path and simply accelerates, the inertia of the float mechanism causes the compass to swing to the north. Rapid deceleration causes it to swing southward. Second, if flying toward the North Pole and a banked turn is made, the downward pull of the magnetic field initially pulls the card away from the direction of the turn. The opposite is true if flying south from the North Pole and a banked turn is initiated. In this case, there is initially a pull of the compass indicator toward the direction of the turn. These kinds of movements are called turning errors. 10-45 N W S E 33 30 24 21 15 12 6 3 Figure 10-81. A vertical magnetic direction indicator provides a realistic reference of headings. Another peculiarity exists with the magnetic compass that is not dip error. Look again at the magnetic compass in Figure 10-79. If flying north or toward any indicated heading, turning the aircraft to the left causes a steady decrease in the heading numbers. But, before the turn is made, the numbers to the left on the compass card are actually increasing. The numbers to the right of the lubber line rotate behind it on a left turn. So, the compass card rotates opposite to the direction of the intended turn. This is because, from the pilot’s seat, you are actually looking at the back of the compass card. While not a major problem, it is more intuitive to see the 360° of direction oriented as they are on an aeronautical chart or a hand-held compass. Vertical Magnetic Compass Solutions to the shortcomings of the simple magnetic compass described above have been engineered. The vertical magnetic compass is a variation of the magnetic compass that eliminates the reverse rotation of the compass card just described. By mounting the main indicating magnets of the compass on a shaft rather than a float, through a series of gears, a compass card can be made to turn about a horizontal axis. This allows the numbers for a heading, towards which the pilot wants to turn, to be oriented correctly on the indicating card. In other words, when turning right, increasing numbers are to the right; when turning left, decreasing numbers rotate in from the left. [Figure 10-81] Many vertical magnetic compasses have also replaced the liquid-filled instrument housing with a dampening cup that uses eddy currents to dampen oscillations. Note that a vertical magnetic compass and a directional gyro look very similar and are often in the lower center position of the instrument panel basic T. Both use the nose of an aircraft as the lubber line against which a rotating compass card is read. Vertical magnetic compasses are characterized by the absence of the hand adjustment knob found on DGs, which is used to align the gyro with a magnetic indication. Remote Indicating Compass Magnetic deviation is compensated for by swinging the compass and adjusting compensating magnets in the instrument housing. A better solution to deviation is to remotely locate the magnetic compass in a wing tip or vertical stabilizer where there is very little interference with the earth’s magnetic field. By using a synchro remote indicating system, the magnetic compass float assembly can act as the rotor of the synchro system. As the float mechanism rotates to align with magnetic north in the remotely located compass, a varied electric current can be produced in the transmitter. This alters the magnetic field produced by the coils of the indicator in the cockpit, and a magnetic indication relatively free from deviation is displayed. Many of these systems are of the magnesyn type. Remote Indicating Slaved Gyro Compass (Flux Gate Compass) An elaborate and very accurate method of direction indication has been developed that combines the use of a gyro, a magnetic compass, and a remote indicating system. It is called the slaved gyro compass or flux gate compass system. A study of the gyroscopic instruments section of this chapter assists in understanding this device. A gyroscopic direction indicator is augmented by magnetic direction information from a remotely located compass. The type of compass used is called a flux valve or flux gate compass. It consists of a very magnetically permeable circular segmented core frame or spider. The earth’s magnetic field flows through this iron core and varies its distribution through segments of the core as the flux valve is rotated via the movement of the aircraft. Pickup coil windings are located on each of the core’s spider legs that are positioned 120° apart. [Figure 10-82] The distribution of earth’s magnetic field flowing through the legs is unique for every directional orientation of the aircraft. A coil is placed in the center of the core and is energized by AC current. As the AC flow passes through zero while changing direction, the earth’s magnetic field is allowed to flow through the core. Then, it is blocked or gated as the magnetic field of the core current flow builds to its peak again. The cycle is repeated at the frequency of 10-46 Universal joint Exciter coil Pickup coils Mounting flange Sealed outer case Sealed inner case Damping fluid Figure 10-82. As the aircraft turns in the earth’s magnetic field, the lines of flux flow lines vary through the permeable core of flux gate, creating variable voltages at the three pickoffs. the AC supplied to the excitation coil. The result is repeated flow and nonflow of the earth’s flux across the pickup coils. During each cycle, a unique voltage is induced in each of the pickup coils reflecting the orientation of the aircraft in the earth’s magnetic field. The electricity that flows from each of the pickup coils is transmitted out of the flux valve via wires into a second unit. It contains an autosyn transmitter, directional gyro, an amplifier, and a triple wound stator that is similar to that found in the indicator of a synchro system. Unique voltage is induced in the center rotor of this stator which reflects the voltage received from the flux valve pickup coils sent through the stator coils. It is amplified and used to augment the position of the DG. The gyro is wired to be the rotor of an autosyn synchro system, which transmits the position of the gyro into an indicator unit located in the cockpit. In the indicator, a vertical compass card is rotated against a small airplane type lubber line like that in a vertical magnetic compass. [Figure 10-83 and 84] Further enhancements to direction finding systems of this type involving the integration of radio navigation aids are common. The radio magnetic indicator (RMI) is one such variation. [Figure 10-85] In addition to the rotating direction indicator of the slaved gyro compass, it contains two pointers. One indicates the bearing to a very high frequency (VHF) omnidirectional range (VOR) station and the other indicates the bearing to a nondirectional automatic direction finder (ADF) beacon. These and other radio navigation aids are discussed further in the communications and navigation chapter of this handbook. It should also be noted that integration of slaved gyro direction indicating system information into auto-pilot systems is also possible. Solid State Magnetometers Solid state magnetometers are used on many modern aircraft. They have no moving parts and are extremely accurate. Tiny layered structures react to magnetism on a molecular level resulting in variations in electron activity. These low power consuming devices can sense not only the direction to the earth’s magnetic poles, but also the angle of the flux field. They are free from oscillation that plagues a standard magnetic compass. They feature integrated processing algorithms and easy integration with digital systems. [Figure 10-86] Sources of Power for Gyroscopic Instruments Gyroscopic instruments are essential instruments used on all aircraft. They provide the pilot with critical attitude and directional information and are particularly important while flying under IFR. The sources of power for these instruments can vary. The main requirement is to spin the gyroscopes at a high rate of speed. Originally, gyroscopic instruments were strictly vacuum driven. A vacuum source pulled air across the gyro inside the instruments to make the gyros spin. Later, electricity was added as a source of power. The turning armature of an electric motor doubles as the gyro rotor. In some aircraft, pressure, rather than vacuum, is used to induce the gyro to spin. Various systems and powering configurations have been developed to provide reliable operation of the gyroscopic instruments. 10-47 400 Hz AC 400 Hz AC Amplifier Dial of remote indicator Fixed phase Variable phase Slaving torque motor Autosyn indicator Autosyn transmitter Directional gyro Flux valve Slaved gyro control 400 Hz AC Rotor Stator Earth’s magnetic field Figure 10-83. A simplified schematic of a flux gate, or slaved gyro, compass system. X HMR2300 Com:@232#486 ID: no: Y Z Figure 10-84. Solid state magnetometer units. 10-48 6 N 33 30 24 2I I5 I2 3 S W E Figure 3-26. IThe compass card in this RMI is driven by signals from a flux valve and it indicates the heading of the aircraft opposite the upper center index mark. Figure 10-85. A radio magnetic indicator (RMI) combines a slaved gyro heading indication (red triangle at top of gauge) with magnetic bearing information to a VOR station (solid pointer) and an ADF station (hollow pointer). 33 30 24 2I I5 I2 6 3 GS GS DC NOV HDG Flux valve or flux gate DG/Amplifier or slaved gyro Direction indicator Figure 10-86. Solid state magnetometers. Vacuum Systems Vacuum systems are very common for driving gyro instruments. In a vacuum system, a stream of air directed against the rotor vanes turns the rotor at high speed. The action is similar to a water wheel. Air at atmospheric pressure is first drawn through a filter(s). It is then routed into the instrument and directed at vanes on the gyro rotor. A suction line leads from the instrument case to the vacuum source. From there, the air is vented overboard. Either a venturi or a vacuum pump can be used to provide the vacuum required to spin the rotors of the gyro instruments. The vacuum value required for instrument operation is usually between 3½ inches to 4½ inches of mercury. It is usually adjusted by a vacuum relief valve located in the supply line. Some turn-and-bank indicators require a lower vacuum setting. This can be obtained through the use of an additional regulating valve in the turn and bank vacuum supply line. Venturi Tube Systems The velocity of the air rushing through a venturi can create sufficient suction to spin instrument gyros. A line is run from the gyro instruments to the throat of the venturi mounted on the outside of the airframe. The low pressure in the venturi tube pulls air through the instruments, spins the gyros, and expels the air overboard through the venturi. This source of gyro power is used on many simple, early aircraft. A light, single-engine aircraft can be equipped with a 2-inch venturi (2 inches of mercury vacuum capacity) to operate the turn and bank indicator. It can also have a larger 8-inch venturi to power the attitude and heading indicators. Simplified illustrations of these venturi vacuum systems are shown in Figure 10-87. Normally, air going into the instruments is filtered. The advantages of a venturi as a suction source are its relatively low cost and its simplicity of installation and operation. It also requires no electric power. But there are serious limitations. A venturi is designed to produce the desired vacuum at approximately 100 mph at standard sea level conditions. Wide variations in airspeed or air density cause the suction developed to fluctuate. Airflow can also be hampered by ice that can form on the venturi tube. Additionally, since the rotor does not reach normal operating speed until after takeoff, preflight operational checks of venturi powered gyro instruments cannot be made. For these reasons, alternate sources of vacuum power were developed. Engine-Driven Vacuum Pump The vane-type engine-driven pump is the most common source of vacuum for gyros installed in general aviation, 10-49 2 MIN TURN DC ELEC L R 20 20 I0 I0 2 0 I 0 8 4 6 2 0 I0 PRESSURE Turn-and-bank indicator Pressure gage Altitude indicator Heading indicator Figure 10-87. Simple venturi tube systems for powering gyroscopic instruments. Shaft Rotor Case Inlet Outlet Vane Figure 10-88. Cutaway view of a vane-type engine-driven vacuum pump used to power gyroscopic instruments. light aircraft. One type of engine-driven pump is geared to the engine and is connected to the lubricating system to seal, cool, and lubricate the pump. Another commonly used pump is a dry vacuum pump. It operates without external lubrication and installation requires no connection to the engine oil supply. It also does not need the air oil separator or gate check valve found in wet pump systems. In many other respects, the dry pump system and oil lubricated system are the same. [Figure 10-88] When a vacuum pump develops a vacuum (negative pressure), it also creates a positive pressure at the outlet of the pump. This pressure is compressed air. Sometimes, it is utilized to operate pressure gyro instruments. The components for pressure systems are much the same as those for a vacuum system as listed below. Other times, the pressure developed by the vacuum pump is used to inflate de-ice boots or inflatable seals or it is vented overboard. An advantage of engine-driven pumps is their consistent performance on the ground and in flight. Even at low engine rpm, they can produce more than enough vacuum so that a regulator in the system is needed to continuously provide the correct suction to the vacuum instruments. As long as the engine operates, the relatively simple vacuum system adequately spins the instrument gyros for accurate indications. However, engine failure, especially on singleengine aircraft, could leave the pilot without attitude and directional information at a critical time. To thwart this shortcoming, often the turn and bank indicator operates with an electrically driven gyro that can be driven by the battery for a short time. Thus, when combined with the aircraft’s magnetic compass, sufficient attitude and directional information is still available. Multiengine aircraft typically contain independent vacuum systems for the pilot and copilot instruments driven by separate vacuum pumps on each of the engines. Should an engine fail, the vacuum system driven by the still operating engine supplies a full complement of gyro instruments. An interconnect valve may also be installed to connect the failed instruments to the still operational pump. Typical Pump-Driven System The following components are found in a typical vacuum system for gyroscopic power supply. A brief description is given of each. Refer to the figures for detailed illustrations. 10-50 Figure 10-89. A vacuum regulator, also known as a suction relief valve, includes a foam filter. To relieve vacuum, outside air of a higher pressure must be drawn into the system. This air must be clean to prevent damage to the pump. Air flow to vacuum pump Air flow from vacuum pump Spring fro Gate check valve open Gate check valve closed Figure 10-90. Gate check valve used to prevent vacuum system damage from engine backfire. Air-oil separator—oil and air in the vacuum pump are exhausted through the separator, which separates the oil from the air; the air is vented overboard and the oil is returned to the engine sump. This component is not present when a dry-type vacuum pump is used. The self-lubricating nature of the pump vanes requires no oil. Vacuum regulator or suction relief valve—since the system capacity is more than is needed for operation of the instruments, the adjustable vacuum regulator is set for the vacuum desired for the instruments. Excess suction in the instrument lines is reduced when the spring-loaded valve opens to atmospheric pressure. [Figure 10-89] Gate check valve—prevents possible damage to the instruments by engine backfire that would reverse the flow of air and oil from the pump. [Figure 10-90] Pressure relief valve—since a reverse flow of air from the pump would close both the gate check valve and the suction relief valve, the resulting pressure could rupture the lines. The pressure relief valve vents positive pressure into the atmosphere. Selector valve—In twin-engine aircraft having vacuum pumps driven by both engines, the alternate pump can be selected to provide vacuum in the event of either engine or pump failure, with a check valve incorporated to seal off the failed pump. Restrictor valve—Since the turn needle of the turn and bank indicator operates on less vacuum than that required by the other instruments, the vacuum in the main line must be reduced for use by this instrument. An in-line restrictor valve performs this function. This valve is either a needle valve or a spring-loaded regulating valve that maintains a constant, reduced vacuum for the turn-and-bank indicator. Air filter—A master air filter screens foreign matter from the air flowing through all the gyro instruments. It is an extremely import filter requiring regular maintenance. Clogging of the master filter reduces airflow and causes a lower reading on the suction gauge. Each instrument is also provided with individual filters. In systems with no master filter that rely only upon individual filters, clogging of a filter does not necessarily show on the suction gauge. Suction gauge—a pressure gauge which indicates the difference between the pressure inside the system and atmospheric or cockpit pressure. It is usually calibrated in inches of mercury. The desired vacuum and the minimum and maximum limits vary with gyro system design. If the desired vacuum for the attitude and heading indicators is 5 inches and the minimum is 4.6 inches, a reading below the latter value indicates that the airflow is not spinning the gyros fast enough for reliable operation. In many aircraft, the system provides a suction gauge selector valve permitting the pilot to check the vacuum at several points in the system. Suction/vacuum pressures discussed in conjunction with the operation of vacuum systems are actually negative pressures, indicated as inches of mercury below that of atmospheric pressure. The minus sign is usually not presented, as the importance is placed on the magnitude of the vacuum developed. In relation to an absolute vacuum (0 psi or 0 "Hg), instrument vacuum systems have positive pressure. Figure 10-91 shows a typical engine-driven pump vacuum system containing the above components. A pump capacity of approximately 10"Hg at engine speeds above 1,000 rpm is normal. Pump capacity and pump size vary in different aircraft, depending on the number of gyros to be operated. Twin-Engine Aircraft Vacuum System Operation Twin-engine aircraft vacuum systems are more complicated. They contain an engine-driven vacuum pump on each engine. The associated lines and components for each pump are 10-51 Heading indicator Attitude indicator Turn-and-bank indicator Suction gauge Air Oil Air/oil separator Vacuum pump Air filter Restrictor valve Suction relief valve Pressure relief valve Selector valve Gate check valve Figure 10-91. A typical pump-driven vacuum system for powering gyroscopic instruments. Pressure-Driven Gyroscopic Instrument Systems Gyroscopic instruments are finely balanced devices with jeweled bearings that must be kept clean to perform properly. When early vacuum systems were developed, only oillubricated pumps were available. Even with the use of air-oil separators, the pressure outputs of these pumps contain traces of oil and dirt. As a result, it was preferred to draw clean air through the gyro instruments with a vacuum system, rather than using pump output pressure that presented the risk of contamination. The development of self-lubricated dry pumps greatly reduced pressure output contaminates. This made pressure gyro systems possible. At high altitudes, the use of pressure-driven gyros is more efficient. Pressure systems are similar to vacuum systems and make use of the same components, but they are designed for pressure instead of vacuum. Thus, a pressure regulator is used instead of a suction relief valve. Filters are still extremely important to prevent damage to the gyros. Normally, air is filtered at the inlet and outlet of the pump in a pressure gyro system. Electrically-Driven Gyroscopic Instrument Systems A spinning motor armature can act as a gyroscope. This is the basis for electrically driven gyroscopic instruments in which the gyro rotor spin is powered by an electric motor. isolated from each other and act as two independent vacuum systems. The vacuum lines are routed from each vacuum pump through a vacuum relief valve and through a check valve to the vacuum four-way selector valve. The four-way valve permits either pump to supply a vacuum manifold. From the manifold, flexible hoses connect the vacuumoperated instruments into the system. To reduce the vacuum for the turn and bank indicators, needle valves are included in both lines to these units. Lines to the artificial horizons and the directional gyro receive full vacuum. From the instruments, lines are routed to the vacuum gauge through a turn and bank selector valve. This valve has three positions: main, left turn and bank (T&B), and right T&B. In the main position, the vacuum gauge indicates the vacuum in the lines of the artificial horizons and directional gyro. In the other positions, the lower value of vacuum for the turn and bank indicators can be read. A schematic of this twin-engine aircraft vacuum system is shown in Figure 10-92. Note the following components: two engine-driven pumps, two vacuum relief valves, two flapper type check valves, a vacuum manifold, a vacuum restrictor for each turn and bank indicator, an engine four-way selector valve, one vacuum gauge, and a turn-and-bank selector valve. Not shown are system and individual instrument filters. A drain line may also be installed at the low point in the system. 10-52 Vacuum manifold Copilot’s turn and bank Pilot’s turn and bank Copilot’s artificial horizon Pilot’s directional gyro Pilot’s artificial horizon Needle valve Needle valve Check valve Check valve Vacuum gauge Relief valve Relief valve Right engine vacuum pump Left engine vacuum pump Vacuum 4-way selector valve Turn-and-bank selector valve Right turn and bank Left turn and bank Main Figure 10-92. An example of a twin-engine instrument vacuum system. Electric gyros have the advantage of being powered by battery for a limited time if a generator fails or an engine is lost. Since air is not sent through the gyro to spin the rotor, contamination worries are also reduced. Also, elimination of vacuum pumps, plumbing, and vacuum system components saves weight. On many small, single-engine aircraft, electric turn-and-bank or turn coordinators are combined with vacuum-powered attitude and directional gyro instruments as a means for redundancy. The reverse is also possible. By combining both types of instruments in the instrument panel, the pilot has more options. On more complex multiengine aircraft, reliable, redundant electrical systems make use of all electricpowered gyro instruments possible. It should be noted that electric gyro instruments have some sort of indicator on the face of the dial to show when the instrument is not receiving power. Usually, this is in the form of a red flag or mark of some sort often with the word “off” written on it. Principles of Gyroscopic Instruments Mechanical Gyros Three of the most common flight instruments, the attitude indicator, heading indicator, and turn needle of the turn-andbank indicator, are controlled by gyroscopes. To understand how these instruments operate, knowledge of gyroscopic principles and instrument power systems is required. 10-53 A A A B C D Figure 10-93. Gyroscopes. North Pole Equator Figure 10-94. Once spinning, a free gyro rotor stays oriented in the same position in space despite the position or location of its base. A mechanical gyroscope, or gyro, is comprised of a wheel or rotor with its mass concentrated around its perimeter. The rotor has bearings to enable it to spin at high speeds. [Figure 10-93A] Different mounting configurations are available for the rotor and axle, which allow the rotor assembly to rotate about one or two axes perpendicular to its axis of spin. To suspend the rotor for rotation, the axle is first mounted in a supporting ring. [Figure 10-93B] If brackets are attached 90° around the supporting ring from where the spin axle attached, the supporting ring and rotor can both move freely 360°. When in this configuration, the gyro is said to be a captive gyro. It can rotate about only one axis that is perpendicular to the axis of spin. [Figure 10-93C] The supporting ring can also be mounted inside an outer ring. The bearing points are the same as the bracket just described, 90° around the supporting ring from where the spin axle attached. Attachment of a bracket to this outer ring allows the rotor to rotate in two planes while spinning. Both of these are perpendicular to the spin axis of the rotor. The plane that the rotor spins in due to its rotation about its axle is not counted as a plane of rotation. A gyroscope with this configuration, two rings plus the mounting bracket, is said to be a free gyro because it is free to rotate about two axes that are both perpendicular to the rotor’s spin axis. [Figure 10-93D] As a result, the supporting ring with spinning gyro mounted inside is free to turn 360° inside the outer ring. Unless the rotor of a gyro is spinning, it has no unusual properties; it is simply a wheel universally mounted. When the rotor is rotated at a high speed, the gyro exhibits a couple of unique characteristics. The first is called gyroscopic rigidity, or rigidity in space. This means that the rotor of a free gyro always points in the same direction no matter which way the base of the gyro is positioned. [Figure 10-94] Gyroscopic rigidity depends upon several design factors: 1. Weight—for a given size, a heavy mass is more resistant to disturbing forces than a light mass. 2. Angular velocity—the higher the rotational speed, the greater the rigidity or resistance is to deflection. 3. Radius at which the weight is concentrated— maximum effect is obtained from a mass when its principal weight is concentrated near the rim, rotating at high speed. 4. Bearing friction—any friction applies a deflecting force to a gyro. Minimum bearing friction keeps deflecting forces at a minimum. 10-54 Beam 1 Beam 2 Start and finish—nonrotationg path Start Finish when path rotates Direction of path rotation Frequency difference A ring laser gyro functions due to the Sagnac Effect Figure 10-96. Light traveling in opposite directions around a nonrotating path arrives at the end of the loop at the same time (top). When the path rotates, light traveling with the rotation must travel farther to complete one loop. Light traveling against the rotation completes the loop sooner (bottom). Plane of Precession Plane of Force Plane of Rotation Applied force Resulting motion (precession) Applied force Figure 10-95. When a force is applied to a spinning gyroscope, it reacts as though the force came from 90° further around the rotor in the direction it is spinning. The plane of the applied force, the plane of the rotation, and the plane in which the gyro responds (known as the plane of precession), are all perpendicular to each other. This characteristic of gyros to remain rigid in space is exploited in the attitude-indicating instruments and the directional indicators that use gyros. Precession is a second important characteristic of gyroscopes. By applying a force to the horizontal axis of the gyro, a unique phenomenon occurs. The applied force is resisted. Instead of responding to the force by moving about the horizontal axis, the gyro moves in response about its vertical axis. Stated another way, an applied force to the axis of the spinning gyro does not cause the axis to tilt. Rather, the gyro responds as though the force was applied 90° around in the direction of rotation of the gyro rotor. The gyro rotates rather than tilts. [Figure 10-95] This predictable controlled precession of a gyroscope is utilized in a turn and bank instrument. Solid State Gyros and Related Systems Improved attitude and direction information is always a goal in aviation. Modern aircraft make use of highly accurate solidstate attitude and directional devices with no moving parts. This results in very high reliability and low maintenance. Ring Laser Gyros (RLG) The ring laser gyro (RLG) is widely used in commercial aviation. The basis for RLG operation is that it takes time for light to travel around a stationary, nonrotating circular path. Light takes longer to complete the journey if the path is rotating in the same direction as the light is traveling. And, it takes less time for the light to complete the loop if the path is rotating in the direction opposite to that of the light. Essentially, the path is made longer or shorter by the rotation of the path. [Figure 10-96] This is known as the Sagnac effect. A laser is light amplification by stimulated emission of radiation. A laser operates by exciting atoms in plasma to release electromagnetic energy, or photons. A ring laser gyro produces laser beams that travel in opposite directions around a closed triangular cavity. The wavelength of the light traveling around the loop is fixed. As the loop rotates, the path the lasers must travel lengthens or shortens. The light wavelengths compress or expand to complete travel around the loop as the loop changes its effective length. As the wavelengths change, the frequencies also change. 10-55 Anode Anode Cathode Mirror Piezoelectric dithering motor Readout detector Corner prism Light beams Fringe pattern Gas discharge region Figure 10-97. The ring laser gyro is rugged, accurate, and free of friction. Figure 10-98. The relative scale size of a MEMS gyro. By examining the difference in the frequencies of the two counterrotating beams of light, the rate at which the path is rotating can be measured. A piezoelectric dithering motor in the center of the unit vibrates to prevent lock-in of the output signal at low rotational speeds. It causes units installed on aircraft to hum when operating. [Figure 10-97] An RLG is remotely mounted so the cavity path rotates around one of the axes of flight. The rate of frequency phase shift detected between the counterrotating lasers is proportional to the rate that the aircraft is moving about that axis. On aircraft, an RLG is installed for each axis of flight. Output can be used in analog instrumentation and autopilot systems. It is also easily made compatible for use by digital display computers and for digital autopilot computers. RLGs are very rugged and have a long service life with virtually no maintenance due to their lack of moving parts. They measure movement about an axis extremely quickly and provide continuous output. They are extremely accurate and generally are considered superior to mechanical gyroscopes. Microelectromechanical Based Attitude and Directional Systems On aircraft, microelectromechanical systems (MEMS) devices save space and weight. Through the use of solid-state MEMS devices, reliability is increased primarily due to the lack of moving parts. The development of MEMS technology for use in aviation instrumentation integrates with the use of ADCs. This newest improvement in technology is low cost and promises to proliferate through all forms of aviation. MEMS for gyroscopic applications are used in small, general aviation aircraft, as well as larger commercial aircraft. Tiny vibration-based units with resistance and capacitance measuring pick-offs are accurate and reliable and only a few millimeters in length and width. They are normally integrated into a complete micro-electronic solid-state chip designed to yield an output after various conditioning processes are performed. The chips, which are analogous to tiny circuit boards, can be packaged for installation inside a dedicated computer or module that is installed on the aircraft. While a large mechanical gyroscope spins in a plane, its rigidity in space is used to observe and measure the movement of the aircraft. The basis of operation of many MEMS gyroscopes is the same despite their tiny size. The difference is that a vibrating or oscillating piezoelectric device replaces the spinning, weighted ring of the mechanical gyro. Still, once set in motion, any out-of-plane motion is detectable by varying microvoltages or capacitances detected through geometrically arranged pickups. Since piezoelectric substances have a relationship between movement and electricity, microelectrical stimulation can set a piezoelectric gyro in motion and the tiny voltages produced via the movement in the piezo can be extracted. They can be input as the required variables needed to compute attitude or direction information. [Figure 10-98] Other Attitude and Directional Systems In modern aircraft, attitude heading and reference systems (AHRS) have taken the place of the gyroscope and other individual instruments. While MEMS devices provide part of the attitude information for the system, GPS, solid state magnetometers, solid state accelerometers, and digital air data signals are all combined in an AHRS to compute and output highly reliable information for display on a cockpit panel. [Figure 10-99] 10-56 XPDR 5537 IDNT LCL23:00:34 VOR 1 270° 2 1 1 2 4300 4200 4100 3900 3900 3800 4300 20 80 4000 4000 130 120 110 90 80 70 1 100 9 TAS 100KT OAT 7°C ALERTS NAV1 117.60 117.90 NAV2 117.90 117.60 132.675 120.000 COM1 118.525 132.900 COM2 WPT _ _ _ _ _ _ DIS _ _ ._ NM DTK _ _ _° TRK 360° N-S E-W VOLTS 27.3 2090 NAV1 117.60 117.90 NAV2 117.90 117.60 132.675 120.000 COM1 118.525 132.900 COM2 WPT _ _ _ _ _ _ DIS _ _ ._ NM DTK _ _ _° TRK 360° MAP - NAVIGATION MAP Figure 10-99. Instrumentation displayed within a glass cockpit using an attitude heading and reference system (AHRS) computer. Common Gyroscopic Instruments Vacuum-Driven Attitude Gyros The attitude indicator, or artificial horizon, is one of the most essential flight instruments. It gives the pilot pitch and roll information that is especially important when flying without outside visual references. The attitude indicator operates with a gyroscope rotating in the horizontal plane. Thus, it mimics the actual horizon through its rigidity in space. As the aircraft pitches and rolls in relation to the actual horizon, the gyro gimbals allow the aircraft and instrument housing to pitch and roll around the gyro rotor that remains parallel to the ground. A horizontal representation of the airplane in miniature is fixed to the instrument housing. A painted semisphere simulating the horizon, the sky, and the ground is attached to the gyro gimbals. The sky and ground meet at what is called the horizon bar. The relationship between the horizon bar and the miniature airplane are the same as those of the aircraft and the actual horizon. Graduated scales reference the degrees of pitch and roll. Often, an adjustment knob allows pilots of varying heights to place the horizon bar at an appropriate level. [Figure 10-100] In a typical vacuum-driven attitude gyro system, air is sucked through a filter and then through the attitude indicator in a manner that spins the gyro rotor inside. An erecting mechanism is built into the instrument to assist in keeping the gyro rotor rotating in the intended plane. Precession caused by bearing friction makes this necessary. After air engages the scalloped drive on the rotor, it flows from the instrument to the vacuum pump through four ports. These ports all exhaust the same amount of air when the gyro is rotating in plane. When the gyro rotates out of plane, air tends to port out of one side more than another. Vanes close to prevent this, causing more air to flow out of the opposite side. The force from this unequal venting of the air re-erects the gyro rotor. [Figure 10-101] Early vacuum-driven attitude indicators were limited in how far the aircraft could pitch or roll before the gyro gimbals contacted stops, causing abrupt precession and tumbling of the gyro. Many of these gyros include a caging device. It is used to erect the rotor to its normal operating position prior to flight or after tumbling. A flag indicates that the gyro must be uncaged before use. More modern gyroscopic instruments are built so they do not tumble, regardless of the angular movement of the aircraft about its axes. In addition to the contamination potential introduced by the air-drive system, other shortcomings exist in the performance of vacuum-driven attitude indicators. Some are induced by the erection mechanism. The pendulous vanes that move to direct airflow out of the gyro respond not only to forces caused by a deviation from the intended plane of rotation, but centrifugal force experienced during turns also causes the vanes to allow asymmetric porting of the gyro vacuum air. The result is inaccurate display of the aircraft’s attitude, especially in skids and steep banked turns. Also, abrupt acceleration and deceleration imposes forces on the gyro rotor. Suspended in its gimbals, it acts similar to an accelerometer, resulting in a false nose-up or nose-down indication. Pilots must learn to recognize these errors and adjust accordingly. 10-57 I0 I0 2 0 I 0 20 20 20 I0 I0 2 0 I 0 0 20 Figure 10-100. A typical vacuum-driven attitude indicator shown with the aircraft in level flight (left) and in a climbing right turn (right). Precession Applied force Precession Port A Port A Exhaust air equal in all directions gyro erect Gyro precesses, increasing exhaust from port A Precessing force at port A erects gyro, exhaust air again equal at all ports Figure 10-101. The erecting mechanism of a vacuum-driven attitude indicator. Electric Attitude Indicators Electric attitude indicators are very similar to vacuumdriven gyro indicators. The main difference is in the drive mechanism. Inside the gimbals of an electric gyro, a small squirrel cage electric motor is the rotor. It is typically driven by 115-volt, 400-cycle AC. It turns at approximately 21,000 rpm. Other characteristics of the vacuum-driven gyro are shared by the electric gyro. The rotor is still oriented in the horizontal plane. The free gyro gimbals allow the aircraft and instrument case to rotate around the gyro rotor that remains rigid in space. A miniature airplane fixed to the instrument case indicates the aircraft’s attitude against the moving horizon bar behind it. Electric attitude indicators address some of the shortcomings of vacuum-driven attitude indicators. Since there is no air flowing through an electric attitude indicator, air filters, regulators, plumbing lines and vacuum pump(s) are not needed. Contamination from dirt in the air is not an issue, resulting in the potential for longer bearing life and less precession. Erection mechanism ports are not employed, so pendulous vanes responsive to centrifugal forces are eliminated. It is still possible that the gyro may experience precession and need to be erected. This is done with magnets rather than vent ports. A magnet attached to the top of the gyro shaft spins at approximately 21,000 rpm. Around this magnet, but not attached to it, is a sleeve that is rotated by magnetic attraction at approximately 44 to 48 rpm. Steel balls are free to move around the sleeve. If the pull of gravity is not aligned with the axis of the gyro, the balls fall to the low side. The resulting precession re-aligns the axis of rotation vertically. 10-58 30 60 90 Figure 10-103. A typical vacuum-powered gyroscopic direction indicator, also known as a directional gyro. PUSH TO CAGE Erection mechanism Magnet 44–48 rpm Gyro universally mounted 21,000 rpm Reaction to precession forces Caging mechanism OFF Figure 10-102. Erecting and caging mechanisms of an electric attitude indicator. Typically, electric attitude indicator gyros can be caged manually by a lever and cam mechanism to provide rapid erection. When the instrument is not getting sufficient power for normal operation, an off flag appears in the upper right hand face of the instrument. [Figure 10-102] Gyroscopic Direction Indicator or Directional Gyro (DG) The gyroscopic direction indicator or directional gyro (DG) is often the primary instrument for direction. Because a magnetic compass fluctuates so much, a gyro aligned with the magnetic compass gives a much more stable heading indication. Gyroscopic direction indicators are located at the center base of the instrument panel basic T. A vacuum-powered DG is common on many light aircraft. Its basis for operation is the gyro’s rigidity in space. The gyro rotor spins in the vertical plane and stays aligned with the direction to which it is set. The aircraft and instrument case moves around the rigid gyro. This causes a vertical compass card that is geared to the rotor gimbal to move. It is calibrated in degrees, usually with every 30 degrees labeled. The nose of a small, fixed airplane on the instrument glass indicates the aircraft’s heading. [Figure 10-103] Vacuum-driven direction indicators have many of the same basic gyroscopic instrument issues as attitude indicators. Built-in compensation for precession varies and a caging device is usually found. Periodic manual realignment with the magnetic compass by the pilot is required during flight. Turn Coordinators Many aircraft make use of a turn coordinator. The rotor of the gyro in a turn coordinator is canted upwards 30°. As such, it responds not only to movement about the vertical axis, but also to roll movements about the longitudinal axis. This is useful because it is necessary to roll an aircraft to turn it about the vertical axis. Instrument indication of roll, therefore, is the earliest possible warning of a departure from straight-and-level flight. Typically, the face of the turn coordinator has a small airplane symbol. The wing tips of the airplane provide the indication of level flight and the rate at which the aircraft is turning. [Figure 10-104] Turn-and-Slip Indicator The turn-and-slip indicator may also be referred to as the turnand- bank indicator, or needle-and-ball indicator. Regardless, it shows the correct execution of a turn while banking the aircraft and indicates movement about the vertical axis of the aircraft (yaw). Most turn-and-slip indicators are located below the airspeed indicator of the instrument panel basic T, just to the left of the direction indicator. 10-59 L R Figure 10-105. Turn-and-slip indicator. Figure 10-104. A turn coordinator senses and indicates the rate of both roll and yaw. The turn-and-slip indicator is actually two separate devices built into the same instrument housing: a turn indicator pointer and slip indicator ball. The turn pointer is operated by a gyro that can be driven by a vacuum, air pressure, or by electricity. The ball is a completely independent device. It is a round agate, or steel ball, in a glass tube filled with dampening fluid. It moves in response to gravity and centrifugal force experienced in a turn. Turn indicators vary. They all indicate the rate at which the aircraft is turning. Three degrees of turn per second cause an aircraft to turn 360° in 2 minutes. This is considered a standard turn. This rate can be indicated with marks right and left of the pointer, which normally rests in the vertical position. Sometimes, no marks are present and the width of the pointer is used as the calibration device. In this case, one pointer width deflection from vertical is equal to the 3° per second standard 2-minute turn rate. Faster aircraft tend to turn more slowly and have graduations or labels that indicate 4-minute turns. In other words, a pointer’s width or alignment with a graduation mark on this instrument indicates that the aircraft is turning a 11⁄2° per second and completes a 360° turn in 4 minutes. It is customary to placard the instrument face with words indicating whether it is a 2-or 4-minute turn indicator. [Figure 10-105] The turn pointer indicates the rate at which an aircraft is turning about its vertical axis. It does so by using the precession of a gyro to tilt a pointer. The gyro spins in a vertical plane aligned with the longitudinal axis of the aircraft. When the aircraft rotates about its vertical axis during a turn, the force experienced by the spinning gyro is exerted about the vertical axis. Due to precession, the reaction of the gyro rotor is 90° further around the gyro in the direction of spin. This means the reaction to the force around the vertical axis is movement around the longitudinal axis of the aircraft. This causes the top of the rotor to tilt to the left or right. The pointer is attached with linkage that makes the pointer deflect in the opposite direction, which matches the direction of turn. So, the aircraft’s turn around the vertical axis is indicated around the longitudinal axis on the gauge. This is intuitive to the pilot when regarding the instrument, since the pointer indicates in the same direction as the turn. [Figure 10-106] The slip indicator (ball) part of the instrument is an inclinometer. The ball responds only to gravity during coordinated straight-and-level flight. Thus, it rests in the lowest part of the curved glass between the reference wires. When a turn is initiated and the aircraft is banked, both gravity and the centrifugal force of the turn act upon the ball. If the turn is coordinated, the ball remains in place. Should a skidding turn exist, the centrifugal force exceeds the force of gravity on the ball and it moves in the direction of the outside of the turn. During a slipping turn, there is more bank than needed, and gravity is greater than the centrifugal force acting on the ball. The ball moves in the curved glass toward the inside of the turn. As mentioned previously, often power for the turn-andslip indicator gyro is electrical if the attitude and direction indicators are vacuum powered. This allows limited operation off battery power should the vacuum system and the electric generator fail. The directional and attitude information from the turn-and-slip indicator, combined with information from the pitot static instruments, allow continued safe emergency operation of the aircraft. Electrically powered turn-and-slip indicators are usually DC powered. Vacuum-powered turn-and-slip indicators are usually run on less vacuum (approximately 2 "Hg) than fully gimbaled attitude and direction indicators. Regardless, proper vacuum must be maintained for accurate turn rate information to be displayed. 10-60 Gyro rotation Resultant force on gyro Yaw force Aircraft yaw rotation Gimbal rotation Figure 10-106. The turn-and-slip indicator’s gyro reaction to the turning force in a right hand turn. The yaw force results in a force on the gyro 90° around the rotor in the direction it is turning due to precession. This causes the top of the rotor to tilt to the left. Through connecting linkage, the pointer tilts to the right. Autopilot Systems An aircraft automatic pilot system controls the aircraft without the pilot directly maneuvering the controls. The autopilot maintains the aircraft’s attitude and/or direction and returns the aircraft to that condition when it is displaced from it. Automatic pilot systems are capable of keeping aircraft stabilized laterally, vertically, and longitudinally. The primary purpose of an autopilot system is to reduce the work strain and fatigue of controlling the aircraft during long flights. Most autopilots have both manual and automatic modes of operation. In the manual mode, the pilot selects each maneuver and makes small inputs into an autopilot controller. The autopilot system moves the control surfaces of the aircraft to perform the maneuver. In automatic mode, the pilot selects the attitude and direction desired for a flight segment. The autopilot then moves the control surfaces to attain and maintain these parameters. Autopilot systems provide for one-, two-, or three-axis control of an aircraft. Those that manage the aircraft around only one axis control the ailerons. They are single-axis autopilots, known as wing leveler systems, usually found on light aircraft. [Figure 10-107] Other autopilots are two-axis systems that control the ailerons and elevators. Three-axis autopilots control the ailerons, elevators, and the rudder. Two-and threeaxis autopilot systems can be found on aircraft of all sizes. There are many autopilot systems available. They feature a wide range of capabilities and complexity. Light aircraft typically have autopilots with fewer capabilities than highperformance and transport category aircraft. Integration of navigation functions is common, even on light aircraft autopilots. As autopilots increase in complexity, they not only manipulate the flight control surfaces, but other flight parameters as well. Some modern small aircraft, high-performance, and transport category aircraft have very elaborate autopilot systems known as automatic flight control systems (AFCS). These three-axis systems go far beyond steering the airplane. They control the aircraft during climbs, descents, cruise, and approach to landing. Some even integrate an auto-throttle function that automatically controls engine thrust that makes autolandings possible. For further automatic control, flight management systems have been developed. Through the use of computers, an entire flight profile can be programmed ahead of time allowing the pilot to supervise its execution. An FMS computer coordinates nearly every aspect of a flight, including the autopilot and auto throttle systems, navigation route selection, fuel management schemes, and more. Basis for Autopilot Operation The basis for autopilot system operation is error correction. When an aircraft fails to meet the conditions selected, an error is said to have occurred. The autopilot system automatically corrects that error and restores the aircraft to the flight attitude desired by the pilot. There are two basic ways modern autopilot systems do this. One is position based and the other is rate based. A position based autopilot manipulates the aircraft’s controls so that any deviation from the desired attitude of the aircraft is corrected. This is done by memorizing 10-61 Engine-driven vacuum pump Suction from controller Diaphragm Cable clamp WWiinngg lleevveelleerr vvaaccuuuum lines Turn coordinator Figure 10-107. The wing leveler system on a small aircraft is a vacuum-operated single-axis autopilot. Only the ailerons are controlled. The aircraft’s turn coordinator is the sensing element. Vacuum from the instrument vacuum system is metered to the diaphragm cable actuators to move the ailerons when the turn coordinator senses roll. the desired aircraft attitude and moving the control surfaces so that the aircraft returns to that attitude. Rate based autopilots use information about the rate of movement of the aircraft, and move control surfaces to counter the rate of change that causes the error. Most large aircraft use rate-based autopilot systems. Small aircraft may use either. Autopilot Components Most autopilot systems consist of four basic components, plus various switches and auxiliary units. The four basic components are: sensing elements, computing element, output elements, and command elements. Many advanced autopilot systems contain a fifth element: feedback or follow-up. This refers to signals sent as corrections are being made by the output elements to advise the autopilot of the progress being made. [Figure 10-108] Sensing Elements The attitude and directional gyros, the turn coordinator, and an altitude control are the autopilot sensing elements. These units sense the movements of the aircraft. They generate electric signals that are used by the autopilot to automatically take the required corrective action needed to keep the aircraft flying as intended. The sensing gyros can be located in the cockpit mounted instruments. They can also be remotely mounted. Remote gyro sensors drive the servo displays in the cockpit panel, as well as provide the input signals to the autopilot computer. Modern digital autopilots may use a variety of different sensors. MEMS gyros may be used or accompanied by the use solid state accelerometers and magnetometers. Rate based systems may not use gyros at all. Various input sensors may be located within the same unit or in separate units that transfer information via digital data bus. Navigation information is also integrated via digital data bus connection to avionics computers. Computer and Amplifier The computing element of an autopilot may be analog or digital. Its function is to interpret the sensing element data, integrate commands and navigational input, and send signals to the output elements to move the flight controls as required to control the aircraft. An amplifier is used to strengthen the signal for processing, if needed, and for use by the output devices, such as servo motors. The amplifier and associated circuitry is the computer of an analog autopilot system. Information is handled in channels corresponding to the axis of control for which the signals are intended (i.e., pitch channel, roll channel, or yaw channel). Digital systems use solid state microprocessor computer technology and typically only amplify signals sent to the output elements. Output Elements The output elements of an autopilot system are the servos that cause actuation of the flight control surfaces. They are independent devices for each of the control channels that 10-62 PITCH UP DOWN TURN R L ON OFF ON OFF NAV ALT MASTER TEST STBY PWR 20 20 I0 I0 I0 I0 20 20 33 30 24 2I I5 I2 6 3 Selected radio navigation input Altitude control Altitude indicator Turn-and-bank indicator gyro Directional gyro indicator Aileron servo actuator Rudder servo actuator Elevator servo actuator Computer Flight controller Sensing Elements Command Elements Output Elements Feedback Feedback Feedback Command Command Command e ato se o actuato ck ck and ck Figure 10-108. Typical analog autopilot system components. integrate into the regular flight control system. Autopilot servo designs vary widely depending on the method of actuation of the flight controls. Cable-actuated systems typically utilize electric servo motors or electro-pneumatic servos. Hydraulic actuated flight control systems use electrohydraulic autopilot servos. Digital fly-by-wire aircraft utilize the same actuators for carrying out manual and autopilot maneuvers. When the autopilot is engaged, the actuators respond to commands from the autopilot rather than exclusively from the pilot. Regardless, autopilot servos must allow unimpeded control surface movement when the autopilot is not operating. Aircraft with cable actuated control surfaces use two basic types of electric motor-operated servos. In one, a motor is connected to the servo output shaft through reduction gears. The motor starts, stops, and reverses direction in response to the commands of autopilot computer. The other type of 10-63 DC motor Clutches Capstan Capstan Reversible motor Bridle cable Control cable Clutch activation signal from amplifier Figure 10-109. A reversible motor with capstan and bridle cable (left), and a single-direction constant motor with clutches that drive the output shafts and control cable in opposite directions (right). PITCH UP DOWN TURN R L ON OFF ON OFF NAV ALT MASTER Figure 10-110. An autopilot controller of a simple autopilot system. electric servo uses a constantly running motor geared to the output shaft through two magnetic clutches. The clutches are arranged so that energizing one clutch transmits motor torque to turn the output shaft in one direction; energizing the other clutch turns the shaft in the opposite direction. [Figure 10-109] Electropneumatic servos can also be used to drive cable flight controls in some autopilot systems. They are controlled by electrical signals from the autopilot amplifier and actuated by an appropriate air pressure source. The source may be a vacuum system pump or turbine engine bleed air. Each servo consists of an electromagnetic valve assembly and an output linkage assembly. Aircraft with hydraulically actuated flight control systems have autopilot servos that are electro-hydraulic. They are control valves that direct fluid pressure as needed to move the control surfaces via the control surface actuators. They are powered by signals from the autopilot computer. When the autopilot is not engaged, the servos allow hydraulic fluid to flow unrestricted in the flight control system for normal operation. The servo valves can incorporate feedback transducers to update the autopilot of progress during error correction. Command Elements The command unit, called a flight controller, is the human interface of the autopilot. It allows the pilot to tell the autopilot what to do. Flight controllers vary with the complexity of the autopilot system. By pressing the desired function buttons, the pilot causes the controller to send instruction signals to the autopilot computer, enabling it to activate the proper servos to carry out the command(s). Level flight, climbs, descents, turning to a heading, or flying a desired heading are some of the choices available on most autopilots. Many aircraft make use of a multitude of radio navigational aids. These can be selected to issue commands directly to the autopilot computer. [Figure 10-110] In addition to an on/off switch on the autopilot controller, most autopilots have a disconnect switch located on the control wheel(s). This switch, operated by thumb pressure, can be used to disengage the autopilot system should a malfunction occur in the system or any time the pilot wishes to take manual control of the aircraft. Feedback or Follow-up Element As an autopilot maneuvers the flight controls to attain a desired flight attitude, it must reduce control surface correction as the desired attitude is nearly attained so the controls and aircraft come to rest on course. Without doing so, the system would continuously overcorrect. Surface deflection would occur until the desired attitude is attained. But movement would still occur as the surface(s) returned to pre-error position. The attitude sensor would once again detect an error and begin the correction process all over again. 10-64 Gyro input signal Control surface feedback signal Feedback circuit Amplifier B A Servo Figure 10-111. Basic function of an analog autopilot system including follow-up or feedback signal. Various electric feedback, or follow-up signals, are generated to progressively reduce the error message in the autopilot so that continuous over correction does not take place. This is typically done with transducers on the surface actuators or in the autopilot servo units. Feedback completes a loop as illustrated in Figure 10-111. A rate system receives error signals from a rate gyro that are of a certain polarity and magnitude that cause the control surfaces to be moved. As the control surfaces counteract the error and move to correct it, follow-up signals of opposite polarity and increasing magnitude counter the error signal until the aircraft’s correct attitude is restored. A displacement follow-up system uses control surface pickups to cancel the error message when the surface has been moved to the correct position. Autopilot Functions The following autopilot system description is presented to show the function of a simple analog autopilot. Most autopilots are far more sophisticated; however, many of the operating fundamentals are similar. The automatic pilot system flies the aircraft by using electrical signals developed in gyro-sensing units. These units are connected to flight instruments that indicate direction, rate of turn, bank, or pitch. If the flight attitude or magnetic heading is changed, electrical signals are developed in the gyros. These signals are sent to the autopilot computer/amplifier and are used to control the operation of servo units. A servo for each of the three control channels converts electrical signals into mechanical force, which moves the control surface in response to corrective signals or pilot commands. The rudder channel receives two signals that determine when and how much the rudder moves. The first signal is a course signal derived from a compass system. As long as the aircraft remains on the magnetic heading it was on when the autopilot was engaged, no signal develops. But, any deviation causes the compass system to send a signal to the rudder channel that is proportional to the angular displacement of the aircraft from the preset heading. The second signal received by the rudder channel is the rate signal that provides information anytime the aircraft is turning about the vertical axis. This information is provided by the turn-and-bank indicator gyro. When the aircraft attempts to turn off course, the rate gyro develops a signal proportional to the rate of turn, and the course gyro develops a signal proportional to the amount of displacement. The two signals are sent to the rudder channel of the amplifier, where they are combined and their strength is increased. The amplified signal is then sent to the rudder servo. The servo turns the rudder in the proper direction to return the aircraft to the selected magnetic heading. As the rudder surface moves, a follow-up signal is developed that opposes the input signal. When the two signals are equal in magnitude, the servo stops moving. As the aircraft arrives on course, the course signal reaches a zero value, and the rudder is returned to the streamline position by the follow-up signal. The aileron channel receives its input signal from a transmitter located in the gyro horizon indicator. Any movement of the aircraft about its longitudinal axis causes the gyro-sensing unit to develop a signal to correct for the movement. This signal is amplified, phase detected, and sent to the aileron servo, which moves the aileron control surfaces to correct for the error. As the aileron surfaces move, a follow-up signal builds up in opposition to the input signal. When the two signals are equal in magnitude, the servo stops moving. Since the ailerons are displaced from the streamline, the aircraft now starts moving back toward level flight with the input signal becoming smaller and the follow-up signal driving the control surfaces back toward the streamline position. When the aircraft has returned to level flight roll attitude, the input signal is again zero. At the same time, the control surfaces are streamlined, and the follow-up signal is zero. The elevator channel circuits are similar to those of the aileron channel, with the exception that the elevator channel detects and corrects changes in pitch attitude of the aircraft. For altitude control, a remotely mounted unit containing an altitude pressure diaphragm is used. Similar to the attitude and directional gyros, the altitude unit generates error signals when the aircraft has moved from a preselected altitude. This is known as an altitude hold function. The signals control the pitch servos, which move to correct the error. An altitude select function causes the signals to continuously be sent to the pitch servos until a preselected altitude has been reached. The aircraft then maintains the preselected altitude using altitude hold signals. 10-65 MACH HOLD IAS HOLD IAS ACO INS BACK BEAM MAX CLIMB MAX CRUISE TRK HDG HDG HOLD TURN GLIDE PITCH HOLD MACH HOLD VOR LOG LAND IAS HOLD VERT SPEED ALT HOLD ALT AUTOLAND ACQ CAT 3 GO GROUND 1 7 0 0 0 11 11 00 4 0 0 2 2 4 1 3 0 2 2 0 AP1 AP2 AP2 FD2 FD1 AT1 AT2 AP1 AUTOTHROTTLE SPEED SELECT KNOTS ALTITUDE SELECT FEET Figure 10-112. The AFCS control panel commands several integrated systems from a single panel including: flight directors, autopilots, autothrottles, autoland, and navigational aids. Mode selections for many features are made from this single interface. Yaw Dampening Many aircraft have a tendency to oscillate around their vertical axis while flying a fixed heading. Near continuous rudder input is needed to counteract this effect. A yaw damper is used to correct this motion. It can be part of an autopilot system or a completely independent unit. A yaw damper receives error signals from the turn coordinator rate gyro. Oscillating yaw motion is counteracted by rudder movement, which is made automatically by the rudder servo(s) in response to the polarity and magnitude of the error signal. Automatic Flight Control System (AFCS) An aircraft autopilot with many features and various autopilot related systems integrated into a single system is called an automatic flight control system (AFCS). These were formerly found only on high-performance aircraft. Currently, due to advances in digital technology for aircraft, modern aircraft of any size may have AFCS. AFCS capabilities vary from system to system. Some of the advances beyond ordinary autopilot systems are the extent of programmability, the level of integration of navigational aids, the integration of flight director and autothrottle systems, and combining of the command elements of these various systems into a single integrated flight control human interface. [Figure 10-112] It is at the AFCS level of integration that an autothrottle system is integrated into the flight director and autopilot systems with glide scope modes so that auto landings are possible. Small general aviation aircraft being produced with AFCS may lack the throttle-dependent features. Modern general aviation AFCS are fully integrated with digital attitude heading and reference systems (AHRS) and navigational aids including glideslope. They also contain modern computer architecture for the autopilot (and flight director systems) that is slightly different than described above for analog autopilot systems. Functionality is distributed across a number of interrelated computers and includes the use of intelligent servos that handle some of the error correction calculations. The servos communicate with dedicated avionics computers and display unit computers through a control panel, while no central autopilot computer exists. [Figure 10-113] Flight Director Systems A flight director system is an instrument system consisting of electronic components that compute and indicate the aircraft attitude required to attain and maintain a preselected flight condition. A command bar on the aircraft’s attitude indicator shows the pilot how much and in what direction the attitude of the aircraft must be changed to achieve the desired result. The computed command indications relieve the pilot of many of the mental calculations required for instrument flights, such as interception angles, wind drift correction, and rates of climb and descent. Essentially, a flight director system is an autopilot system without the servos. All of the same sensing and computations are made, but the pilot controls the airplane and makes maneuvers by following the commands displayed on the instrument panel. Flight director systems can be part of an autopilot system or exist on aircraft that do not possess full autopilot systems. Many autopilot systems allow for the option of engaging or disengaging a flight director display. Flight director information is displayed on the instrument that displays the aircraft’s attitude. The process is accomplished with a visual reference technique. A symbol representing the aircraft is fit into a command bar positioned by the flight director in the proper location for a maneuver to be accomplished. The symbols used to represent the aircraft and the command bar vary by manufacturer. Regardless, the object is always to fly the aircraft symbol into the command bar symbol. [Figure 10-114] 10-66 PFD MFD Mode controller Go-around switch Integrated avionics unit 1 Integrated avionics unit 2 A/P disc 4 way trim Pitch trim adapter Pitch trim cartridge Roll servo Pitch servo Yaw servo (optional) AHRS Figure 10-113. Automatic flight control system (AFCS) of a Garmin G1000 glass cockpit instrument system for a general aviation aircraft. 10-67 20 20 I0 I0 I0 I0 20 20 20 20 I0 I0 I0 I0 20 20 Flight director command bars Airplane symbol Figure 10-114. The fight director command bar signals the pilot how to steer the aircraft for a maneuver. By flying the aircraft so the triangular airplane symbol fits into the command bar, the pilot performs the maneuver calculated by the flight director. The instrument shown on the left is commanding a climb while the airplane is flying straight and level. The instrument on the right shows that the pilot has accomplished the maneuver. The instrument that displays the flight director commands is known as a flight director indicator (FDI), attitude director indicator (ADI), or electronic attitude director indicator (EADI). It may even be referred to as an artificial horizon with flight director. This display element combines with the other primary components of the flight director system. Like an autopilot, these consist of the sensing elements, a computer, and an interface panel. Integration of navigation features into the attitude indicator is highly useful. The flight director contributes to this usefulness by indicating to the pilot how to maneuver the airplane to navigate a desired course. Selection of the VOR function on the flight director control panel links the computer to the omnirange receiver. The pilot selects a desired course and the flight director displays the bank attitude necessary to intercept and maintain this course. Allocations for wind drift and calculation of the intercept angle is performed automatically. Flight director systems vary in complexity and features. Many have altitude hold, altitude select, pitch hold, and other features. But flight director systems are designed to offer the greatest assistance during the instrument approach phase of flight. ILS localizer and glideslope signals are transmitted through the receivers to the computer and are presented as command indications. This allows the pilot to fly the airplane down the optimum approach path to the runway using the flight director system. With the altitude hold function engaged, level flight can be maintained during the maneuvering and procedure turn phase of an approach. Altitude hold automatically disengages when the glideslope is intercepted. Once inbound on the localizer, the command signals of the flight director are maintained in a centered or zero condition. Interception of the glideslope causes a downward indication of the command pitch indicator. Any deviation from the proper glideslope path causes a fly-up or fly-down command indication. The pilot needs only to keep the airplane symbol fit into the command bar. Electronic Instruments Electronic Attitude Director Indicator (EADI) The EADI is an advanced version of attitude and electric attitude indicators previously discussed. In addition to displaying the aircraft’s attitude, numerous other situational flight parameters are displayed. Most notable are those that relate to instrument approaches and the flight director command bars. Annunciation of active systems, such as the AFCS and navigation systems, is typical. The concept behind an EADI is to put all data related to the flight situation in close proximity for easy observation by the pilot. [Figure 10-115] Most EADIs can be switched between different display screens depending on the preference of the pilot and the phase of flight. EADIs vary from manufacturer to manufacturer and aircraft to aircraft. However, most of the same information is displayed. EADIs can be housed in a single instrument housing or can be part of an electronic instrument display system. One such system, the electronic flight instrument system (EFIS), uses a cathode ray tube EADI display driven by a signal generator. Large-screen glass cockpit displays use LCD technology to display EADI information as part of an entire situational display directly in front of the pilot in the middle of the instrument panel. Regardless, the EADI 10-68 F S 20 10 10 20 20 10 10 20 ALT GS 264 DH 200 1060 M Roll scale Roll pointer Selected decision height Director command bar Radio altitude Flight director pitch and roll command bars Glideslope deviation scale Glideslope deviation pointer Marker beacon Localizer deviation scale Localizer deviation pointer Slip indicator Speed error pointer Speed error scale Altitude sphere Pitch scale markers Groundspeed Altitude alert Aircraft symbol Figure 10-115. Some of the many parameters and features of an electronic attitude director indicator (EADI). is the primary flight instrument used for aircraft attitude information during instrument flying and especially during instrument approaches. It is almost always accompanied by an electronic horizontal situation indicator (EHSI) located just below it in the display panel. Electronic Horizontal Situation Indicators (EHSI) The EHSI is an evolved version of the horizontal situation indicator (HSI), which was born from the gyroscopic direction indicator or directional gyro. The HSI incorporates directional information to two different navigational aids, as well as the heading of the aircraft. The EHSI does this and more. Its primary purpose is to display as much useful navigational information as possible. In conjunction with a flight management computer and a display controller, an EHSI can display information in PLAN, MAP, VOR, and ILS modes. The PLAN mode shows a fixed map of the input flight plan. This usually includes all selected navigational aids for each flight segment and the destination airport. The MAP mode shows the aircraft against a detailed moving map background. Active and inactive navigational aids are shown, as well as other airports and waypoints. Weather radar information may be selected to be shown in scale as a background. Some HSIs can depict other air traffic when integrated with the TCAS system. Unlike a standard HSI, an EHSI may show only the pertinent portion of the compass rose. Annunciation of active mode and selected features appear with other pertinent information, such as distance and arrival time to the next waypoint, airport designators, wind direction and speed, and more. [Figure 10-116] There are many different displays that vary by manufacturer. The VOR view of an EHSI presents a more traditional focus on a selected VOR, or other navigational station being used, during a particular flight segment. The entire compass rose, the traditional lateral deviation pointer, to/from information, heading, and distance information are standard. Other information may also be displayed. [Figure 10-117] The ILS mode of an EHSI shows the aircraft in relation to the ILS approach aids and selected runway with varying degrees of details. With this information displayed, the pilot need not consult printed airport approach information, allowing full attention to flying the aircraft. Electronic Flight Information Systems In an effort to increase the safety of operating complicated aircraft, computers and computer systems have been incorporated. Flight instrumentation and engine and airframe monitoring are areas particularly well suited to gain advantages from the use of computers. They contribute by helping to reduce instrument panel clutter and focusing the pilot’s attention only on matters of imminent importance. “Glass cockpit” is a term that refers to the use of flat-panel display screens in cockpit instrumentation. In reality, it also refers to the use of computer-produced images that have replaced individual mechanical gauges. Moreover, computers and computer systems monitor the processes and components of an operating aircraft beyond human ability while relieving the pilot of the stress from having to do so. Computerized electronic flight instrument systems have additional benefits. The solid-state nature of the components increases reliability. Also, microprocessors, data buses, and 10-69 24 27 30NM TRK 252 M 0835.4z ALT 9R 360 FLT WPT08 260 80 CSP SCARR 20 Airplane track to heading ETA Windspeed Airplane symbol Present heading WXR display Procedure turn Runway centerline Marker beacon Waypoint and ID VOR and ID Active flight plan path Wind direction Curved trend vector Range to selected altitude Runway and ID Tuned NAVAID radial Holding pattern Intersection and ID Selected heading vector Selected heading cursor Distance to go VORTAC and ID Vertical deviation pointer Figure 10-116. An EHSI presents navigational information for the entire flight. The pilot selects the mode most useful for a particular phase of flight, ranging from navigational planning to instrument approach to landing. The MAP mode is used during most of the flight. 33 30 24 2I I5 I2 6 3 CMP 1 GS 148 K HDG 315 NAV 2 3.7 NM CRS 300 Selected course Heading data source Heading select bug Forward lubber line Navigation data source Distance Course select pointer To/from indicator Selected heading Aircraft symbol Aft lubber line Reciprocal course pointer Groundspeed Glideslope scale Glideslope pointer Lateral deviation bar Figure 10-117. Approach and VOR mode presentation of an electronic horizontal situation indicator. 10-70 Pilot’s display system Copilot’s display system Display controller Display controller EADI EHIS EADI EHIS Pilot’s symbol generator Center symbol generator Copilot’s symbol generator Data buses Display drive signals Figure 10-118. A simplified diagram of an EFIS system. The EADI and EHSI displays are CRT units in earlier systems. Modern systems use digital displays, sometimes with only one multifunctional display unit replacing the two shown. Independent digital processors can also be located in a single unit to replace the three separate symbol generators. 15 15000 FMS1 DTK 042 44.6NM BCE 3 N 33 30 6 E 12 15 200 4 2 1 1 2 4 28 000 900 800 29.92 IN 100 279 300 280 260 M.708 240 20 10 10 18000 LNV1 VALT Airspeed scale Digital airspeed readout Altitude scale Vertical speed scale Slip indicator Figure 10-119. An EFIS EADI displays an airspeed scale to the left of the horizon sphere and an altimeter and vertical speed scale to the right. The slip indicator is the small rectangle under the direction triangles at the top. This EFIS display presents all of the flight information in the conventional cockpit basic T. LCDs all save space and weight. The following systems have been developed and utilized on aircraft for a number of years. New systems and computer architecture are sure to come in the future. Electronic Flight Instrument System (EFIS) The flight instruments were the first to adopt computer technology and utilize flat screen, multifunctional displays (MFD). EFIS uses dedicated signal generators to drive two independent displays in the center of the basic T. The attitude indicator and directional gyro are replaced by cathode ray tubes (CRT) used to display EADI and EHSI presentations. These enhanced instruments operate alongside ordinary mechanic and electric instruments with limited integration. Still, EADI and EHSI technology is very desirable, reducing workload and panel scan with the added safety provided by integration of navigation information as described. Early EFIS systems have analog technology, while newer models may be digital systems. The signal generators receive information from attitude and navigation equipment. Through a display controller, the pilot can select the various mode or screen features wishing to be displayed. Independent dedicated pilot and copilot systems are normal. A third, backup symbol generator is available to assume operation should one of the two primary units fail. [Figure 10-118] Electronic depiction of ADI and HSI information is the core purpose of an EFIS system. Its expanded size and capabilities over traditional gauges allow for integration of even more flight instrument data. A vertical airspeed scale is typically displayed just left of the attitude field. This is in the same relative position as the airspeed indicator in an analog basic T instrument panel. To the right of the attitude field, many EFIS systems display an altitude and vertical speed scale. Since most EFIS EADI depictions include the inclinometer, normally part of the turn coordinator, all of the basic flight instruments are depicted by the EFIS display. [Figure 10-119] 10-71 Left hand display unit (CRT) Right hand display unit (CRT) Warning light display Flight warning computer Flight warning computer Symbol generator Symbol generator ECAM control panel System data analog converter Aircraft data inputs Aircraft data inputs Aural warnings Aural warnings Discrete analog inputs Figure 10-120. An electronic centralized aircraft monitor (ECAM) system displays aircraft system status, checklists, advisories, and warnings on a pair of controllable monitors. Electronic Centralized Aircraft Monitor (ECAM) The pilot’s workload on all aircraft includes continuous monitoring of the flight instruments and the sky outside of the aircraft. It also includes vigilant scrutiny for proper operation of the engine and airframe systems. On transport category aircraft, this can mean monitoring numerous gauges in addition to maneuvering the aircraft. The electronic centralized aircraft monitoring (ECAM) system is designed to assist with this duty. The basic concept behind ECAM (and other monitoring systems) is automatic performance of monitoring duties for the pilot. When a problem is detected or a failure occurs, the primary display, along with an aural and visual cue, alerts the pilot. Corrective action that needs to be taken is displayed, as well as suggested action due to the failure. By performing system monitoring automatically, the pilot is free to fly the aircraft until a problem occurs. Early ECAM systems only monitor airframe systems. Engine parameters are displayed on traditional full-time cockpit gauges. Later model ECAM systems incorporate engine displays, as well as airframe. An ECAM system has two CRT monitors. In newer aircraft, these may be LCD. The left or upper monitor, depending on the aircraft panel layout, displays information on system status and any warnings associated corrective actions. This is done in a checklist format. The right or lower monitor displays accompanying system information in a pictorial form, such as a diagram of the system being referred to on the primary monitor. The ECAM monitors are typically powered by separate signal generators. Aircraft data inputs are fed into two flight warning computers. Analog inputs are first fed through a system data analog converter and then into the warning computers. The warning computers process the information and forward information to the signal generators to illuminate the monitors. [Figure 10-120] 10-72 TO CONFIG CLR APU ENG DOOR PRESS STS WHEEL ELEC RCL F/CTL HYD EMER CANC ALL FUEL CLR COND BLEED BRT LOWER DISPLAY ECAM UPPER DISPLAY BRT Figure 10-121. An ECAM display control panel. There are four basic modes to the ECAM system: flight phase, advisory, failure related, and manual. The flight phase mode is normally used. The phases are: preflight, takeoff, climb, cruise, descent, approach, and post landing. Advisory and failure–related modes will appear automatically as the situation requires. When an advisory is shown on the primary monitor, the secondary monitor will automatically display the system schematic with numerical values. The same is true for the failure-related mode, which takes precedent over all other modes regardless of which mode is selected at the time of the failure. Color coding is used on the displays to draw attention to matters in order of importance. Display modes are selected via a separate ECAM control panel shown in Figure 10-121. The manual mode of an ECAM is set by pressing one of the synoptic display buttons on the control panel. This allows the display of system diagrams. A failure warning or advisory event will cancel this view. [Figure 10-122] ECAM flight warning computers self-test upon startup. The signal generators are also tested. A maintenance panel allows for testing annunciation and further testing upon demand. BITE stands for built-in test equipment. It is standard for monitoring systems to monitor themselves as well as the aircraft systems. All of the system inputs to the flight warning computers can also be tested for continuity from this panel, as well as inputs and outputs of the system data analog converter. Any individual system faults will be listed on the primary display as normal. Faults in the flight warning computers and signal generators will annunciate on the maintenance panel. [Figure 10-123] Follow the manufacturer’s guidelines when testing ECAM and related systems. Engine Indicating and Crew Alerting System (EICAS) An engine indicating and crew alerting system (EICAS) performs many of the same functions as an ECAM system. The objective is still to monitor the aircraft systems for the pilot. All EICAS display engine, as well as airframe, parameters. Traditional gauges are not utilized, other than a standby combination engine gauge in case of total system failure. EICAS is also a two-monitor, two-computer system with a display select panel. Both monitors receive information from the same computer. The second computer serves as a standby. Digital and analog inputs from the engine and airframe systems are continuously monitored. Caution and warning lights, as well as aural tones, are incorporated. [Figure 10-124] EICAS provides full time primary engine parameters (EPR, N1, EGT) on the top, primary monitor. Advisories and warning are also shown there. Secondary engine parameters and nonengine system status are displayed on the bottom screen. The lower screen is also used for maintenance diagnosis when the aircraft is on the ground. Color coding is used, as well as message prioritizing. The display select panel allows the pilot to choose which computer is actively supplying information. It also controls the display of secondary engine information and system status displays on the lower monitor. EICAS has a unique feature that automatically records the parameters of a failure event to be regarded afterwards by maintenance personnel. Pilots that suspect a problem may be occurring during flight can press the event record button on the display select panel. This also records the parameters for that flight period to be studied later by maintenance. Hydraulic, electrical, environmental, performance, and APU data are examples of what may be recorded. 10-73 HYD BLUE 3000 GREEN 3000 YELLOW 3000 SENRO CTL PSI ELEC PSI ELEC AC TAI 100.1 GEN CYRD APU GEM 100.2 GEM1 OFF GEM2 OFF 115 V 400 Hz TA2 TSS TM ESS BUS BUS 1 BUS 1 EMER BUS 0 1 0 1 0 1 MCNM BUS ELEC DC ESS BUS TA1 BAT 1 BAT 2 BAT 3 ESSTA 28 V 27 V OFF 27 V 28 V 28 V TA2 40 A 0 40 A 0 AIR BLEED RAM AIR APU LP HP HP LP ANTI ICE ANTI ICE PSI °C PSI °C C F LO HI 15 20 C F LO HI 0 0 FUEL OUTER DW: CG: INNER INNER TOTAL: 120.10 CTR LEF: 1000 THIN OUTER ENG 1 JPU ENG 2 47 51 24 54 24 54 6 41 6 41 CAB PRESS 0.0 P PSI FWD BATTERY SFTY GYSO 10 0 2 0 UP DW2 10 0 AFT SYS1 SYS2 CAS ALT FT V/S FT/MIN 0 0 FLT CTL SPLA SPD MAX AQL AQL 7 6 6 7 5 4 4 5 3 2 2 3 1 1 L AIL 0 AUD R AIL 0 NOSE UP DW ELEV STAD 34.3 APU 85 N% 0 122 200 662 EST DFF APU DRM EST °C 15 15 CAPT 15 FWD 15 MID 15 AFT H C 15 H C 15 H C 15 H C AIR COND 11 11 Figure 10-122. Nine of the 12 available system diagrams from the ECAM manual mode. EICAS uses BITE for systems and components. A maintenance panel is included for technicians. From this panel, when the aircraft is on the ground, push-button switches display information pertinent to various systems for analysis. [Figure 10-125] Flight Management System (FMS) The highest level of automated flight system is the flight FMS. Companies flying aircraft for hire have special results they wish to achieve. On-time performance, fuel conservation, and long engine and component life all contribute to profitability. An FMS helps achieve these results by operating the aircraft with greater precision than possible by a human pilot alone. A FMS can be thought of as a master computer system that has control over all other systems, computerized and otherwise. As such, it coordinates the adjustment of flight, engine, and airframe parameters either automatically or by instructing the pilot how to do so. Literally, all aspects of the flight are considered, from preflight planning to pulling up to the jet-way upon landing, including in-flight amendments to planned courses of action. The main component of an FMS is the flight management computer (FMC). It communicates with the EICAS or ECAM, the ADC, the thrust management computer that controls the autothrottle functions, the EIFIS symbol generators, the automatic flight control system, the inertial reference system, collision avoidance systems, and all of the radio navigational aids via data busses. [Figure 10-126] 10-74 WLDP FAULT LOAD DSPL TEST LOAD DSPL FWC 1 FAULT FWC 2 FAULT TEST TEST RE BITE DISPLAY INHIB OVER BITE INPUTS INPUTS ECAM EFIS Aircraft systems inputs test button Inhibited warnings override switch Displays switch Signal generator test button Flight warning computer test buttons and annunciation lights Figure 10-123. An ECAM maintenance panel used for testing and annunciating faults in the ECAM system. The interface to the system is a control display unit (CDU) that is normally located forward on the center pedestal in the cockpit. It contains a full alphanumeric keypad, a CRT or LCD display/work screen, status and condition annunciators, and specialized function keys. [Figure 10-127] The typical FMS uses two FMS FMCs that operate independently as the pilot’s unit and the copilot’s unit. However, they do crosstalk through the data busses. In normal operation, the pilot and copilot divide the workload, with the pilot’s CDU set to supervise and interface with operational parameters and the copilot’s CDU handling navigational chores. This is optional at the flightcrew’s discretion. If a main component fails (e.g., an FMC or a CDU), the remaining operational units continue to operate with full control and without system compromise. Each flight of an aircraft has vertical, horizontal, and navigational components, which are maintained by manipulating the engine and airframe controls. While doing so, numerous options are available to the pilot. Rate of climb, thrust settings, EPR levels, airspeed, descent rates, and other terms can be varied. Commercial air carriers use the FMC to establish guidelines by which flights can be flown. Usually, these promote the company’s goals for fuel and equipment conservation. The pilot need only enter variables as requested and respond to suggested alternatives as the FMC presents them. The FMC has stored in its database literally hundreds of flight plans with predetermined operational parameters that can be selected and implemented. Integration with NAV-COM aids allows the FMS to change radio frequencies as the flight plan is enacted. Internal computations using direct input from fuel flow and fuel quantity systems allow the FMC to carry out lean operations or pursue other objectives, such as high performance operations if making up time is paramount on a particular flight. Weather and traffic considerations are also integrated. The FMS can handle all variables automatically, but communicates via the CDU screen to present its planned action, gain consensus, or ask for an input or decision. As with the monitoring systems, FMS includes BITE. The FMC continuously monitors its entire systems and inputs for faults during operation. Maintenance personal can retrieve system generated and pilot recorded fault messages. They may also access maintenance pages that call out line replaceable units (LRUs) to which faults have been traced by the BITE system. Follow manufacturers’ procedures for interfacing with maintenance data information. Warnings and Cautions Annunciator Systems Instruments are installed for two purposes: to display current conditions and to notify of unsatisfactory conditions. Standardized colors are used to differentiate between visual messages. For example, the color green indicates a satisfactory condition. Yellow is used to caution of a serious condition that requires further monitoring. Red is the color for an unsatisfactory condition. Whether part of the instrument face or of a visual warning system, these colors give quickreference information to the pilot. Most aircraft include annunciator lights that illuminate when an event demanding attention occurs. These use the aforementioned colors in a variety of presentations. Individual lights near the associated cockpit instrument or a collective display of lights for various systems in a central location are common. Words label each light or are part of the light itself to identify any problem quickly and plainly. On complex aircraft, the status of numerous systems and components must be known and maintained. Centralized warning systems have been developed to annunciate critical messages concerning a multitude of systems and components in a simplified, organized manner. Often, this will be done by locating a single annunciator panel somewhere on the instrument panel. These analog aircraft warning systems may look different in various aircraft, and depend on manufacturer preference and the systems installed. [Figure 10-128] EFIS provide for annunciation of advisory and warning messages as part of its flight control and monitoring capabilities, as previously described. Usually, the primary display unit is designated as the location to display annunciations. 10-75 Standy engine indicators Aural warning Upper DU Lower DU Warning & cautions Engine primary displays Engine secondary or status display or maintenance display Discrete caution & warning lights L computer R computer Display switching Maintenance panel Display select panel Other system discretes FCC MCDP TMC interface FEC interface FMC interface RAD Altitude interface ADC interface Engine sensors N1 Oil press N2 Oil quantity N3 Oil temperature EPR Vibration EGT FF System sensors Hydraulic quantity & press ADC hydraulic system temperature Control surface positions Electical system: volts amps freq Generator drive temperature ECS temps APU EGT. RPM Brake temperature Data buses Figure 10-124. Schematic of an engine indicating and crew alerting system (EICAS). Master caution lights are used to draw the attention of the crew to a critical situation in addition to an annunciator that describes the problem. These master caution lights are centrally wired and illuminate whenever any of the participating systems or components require attention. Once notified, the pilot may cancel the master caution, but a dedicated system or component annunciator light stays illuminated until the situation that caused the warning is rectified. Cancelling resets the master caution lights to warn of a subsequent fault event even before the initial fault is corrected. [Figure 10-129] Press to test is available for the entire annunciator system, which energizes all warning circuitry and lights to confirm readiness. Often, this test exposes the need to replace the tiny light bulbs that are used in the system. Aural Warning Systems Aircraft aural warning systems work in conjunction with illuminated annunciator systems. They audibly inform the pilot of a situation requiring attention. Various tones and phrases sound in the cockpit to alert the crew when certain conditions exist. For example, an aircraft with retractable landing gear uses an aural warning system to alert the crew to an unsafe condition. A bell sounds if the throttle is retarded and the landing gear is not in a down and locked condition. A typical transport category aircraft has an aural warning system that alerts the pilot with audio signals for the following: abnormal takeoff, landing, pressurization, mach airspeed conditions, an engine or wheel well fire, calls from 10-76 ECS MSG ELEC HYD PERF APU CONF MCDP ENG EXCD AUTO MANUAL REC ERASE TEST DESPLAY SELECT EICAS MAINT EVENT READ Environmental control systems and maintenance message formats Engine exceedances BITE test switch for self-test routine Configuration and maintenance control/display panel Electrical and hydraulic systems formal Performance and auxiliary power unit formats Selects data from auto or manual event in memory Erases stored data currently displayed Records real-time data currently displayed (in manual event) Figure 10-125. The EICAS maintenance control panel is for the exclusive use of technicians. the crew call system, collision avoidance recommendations, and more. Figure 10-130 shows some of the problems that trigger aural warnings and the action to be taken to correct the situation. Clocks Whether called a clock or a chronometer, an FAA-approved time indicator is required in the cockpit of IFR-certified aircraft. Pilots use a clock during flight to time maneuvers and for navigational purposes. The clock is usually mounted near the flight instrument group, often near the turn coordinator. It indicates hours, minutes, and seconds. For many years, the mechanical 8-day clock was the standard aircraft timekeeping device largely because it continues to run without electrical power as long as it has been hand wound. The mechanical 8-day clock is reliable and accurate enough for its intended use. Some mechanical aircraft clocks feature a push-button elapsed time feature. [Figure 10-131] As electrical systems developed into the reliable, highly redundant systems that exist today, use of an electric clock to replace the mechanical clock began. An electric clock is an analog devise that may also have an elapsed time feature. It can be wired to the battery or battery bus. Thus, it continues to operate in the event of a power failure. Electric aircraft clocks are often used in multiengine aircraft where complete loss of electrical power is unlikely. Many modern aircraft have a digital electronic clock with LED readout. This device comes with the advantages of low power consumption and high reliability due to the lack of moving parts. It is also very accurate. Solid-state electronics allow for expanded features, such as elapsed time, flight time that starts automatically upon takeoff, a stop watch, and memories for all functions. Some even have temperature and date readouts. Although wired into the aircraft’s electrical system, electronic digital clocks may include a small independent battery inside the unit that operates the device should aircraft electrical power fail. [Figure 10-132] 10-77 Flight crew FMC CDU Controls and indicators Data buses Engine control & monitoring (indicating & alerting) Radio NAV VOR ADF DME ILS RAD. ALT ATC transponder WXR GPWS ADC TMC IRS AFCS FMC Navigation database Operational program EFIS symbol generators Control panels AFCS mode IRS mode ILS EFIS ATC VOR/DME RDMI ADF WXR EFIS ADI HSI Maintenance control and display Antennae Control surface servos Database loading Autothrottle servos Sensing probes Discrete inputs from aircraft systems Fuel quantity & fuel flow data Figure 10-126. A flight management system (FMS) integrates numerous engine, aircraft, and navigational systems to provide overall management of the flight. On aircraft with fully digital computerized instrument systems utilizing flat panel displays, the computer’s internal clock, or a GPS clock, can be used with a digital time readout usually located somewhere on the primary flight display. Instrument Housings and Handling Various materials are used to protect the inner workings of aircraft instruments, as well as to enhance the performance of the instrument and other equipment mounted in the immediate vicinity. Instrument cases can be one piece or multipiece. Aluminum alloy, magnesium alloy, steel, iron, and plastic are all common materials for case construction. Electric instruments usually have a steel or iron alloy case to contain electromagnetic flux caused by current flow inside. Despite their rugged outward appearance, all instruments, especially analog mechanical instruments, should be handled with special care and should never be dropped. A crack in an airtight instrument case renders it unairworthy. Ports should never be blown into and should be plugged until the instrument is installed. Cage all gyro instruments until mounted in the instrument panel. Observe all cautions written on the instrument housing and follow the manufacturer’s instruction for proper handling and shipping, as well as installation. 10-78 3 6 0 3 6 0 2 9 9 9 11 M I 0 0 0 1 8 2 9 9 1 9 1 8 2 7 3 6 5 4 0 0 0 0 0 0 50 100 120 140 160 200 400 350 250 1 2 4 .5 6 Up Vert Speed .5 Down 6 4 2 1 0 1 2 3 6 5 4 7 8 9 2 9 9 9 1 AP AT ALT NAV AUX 0 10 20 30 40 50 60 70 80 90 100 0 2 5 0 10 20 30 40 50 60 70 80 90 100 0 2 5 0 10 20 30 40 50 60 70 80 90 100 0 0 2 5 10 20 30 40 50 60 70 80 90 100 0 2 5 10 0 30 20 40 50 60 70 80 90 100 110 10 0 30 20 40 50 60 70 80 90 100 110 10 0 30 20 40 50 60 70 80 90 100 110 10 0 30 20 40 50 60 70 80 90 100 110 10 0 20 30 40 50 60 70 0 3 1 10 0 20 30 40 50 60 70 0 3 1 10 0 20 30 40 50 60 70 0 3 1 10 0 20 30 40 50 60 70 0 3 1 1 4 7 10 13 16 19 3 0 3 1 4 7 10 13 16 19 3 0 3 1 4 7 10 13 16 19 3 0 3 1 4 7 10 13 16 19 3 0 3 14.2 NM 10 20 30 40 50 60 70 10 20 30 40 50 60 70 10 20 30 40 50 60 70 0 10 20 30 0 10 20 30 20 20 10 10 20 20 10 F S 10 ----- 14.2 NAV Master warning and master caution Annunciator panel LEFT GEAR Figure 10-128. The centralized analog annunciator panel has indicator lights from systems and components throughout the aircraft. It is supported by the master caution system. Figure 10-129. A master caution switch removed from the instrument panel. Figure 10-127. The control display unit (CDU) of an FMS. Instrument Installations and Markings Instrument Panels Instrument panels are usually made from sheet aluminum alloy and are painted a dark, nonglare color. They sometimes contain subpanels for easier access to the backs of instruments during maintenance. Instrument panels are usually shockmounted to absorb low-frequency, high-amplitude shocks. The mounts absorb most of the vertical and horizontal vibration, but permit the instruments to operate under conditions of minor vibration. Bonding straps are used to ensure electrical continuity from the panel to the airframe. [Figure 10-133] The type and number of shock mounts to be used for instrument panels are determined by the weight of the unit. Shock-mounted instrument panels should be free to move in all directions and have sufficient clearance to avoid striking the supporting structure. When a panel does not have 10-79 + °C UP D SET B 1 hr up DIM F. T. TIME E.T. DAVTRON ZERO M811B st op RUN Figure 10-132. A typical aircraft electronic clock. Examples of Aircraft Aural Warnings Stage of Operation Warning System Warning Signal Cause of Warning Signal Activation Corrective action Takeoff Flight control Intermittent Throttles are advanced and any of the Correct the aircraft to horn following conditions exist: proper takeoff conditions 1. Speed brakes are not down 2. Flaps are not in takeoff range 3. Auxiliary power exhaust door is open 4. Stabilizer is not in the takeoff setting In flight Mach warning Clacker Equivalent airspeed or mach number Decrease aircraft speed exceeds limits In flight Pressurization Intermittent If cabin pressure becomes equal to Correct the condition horn atmospheric pressure at the specific altitude (altitude at time of occurrence) Landing Landing gear Continuous Landing gear is not down and locked when Raise flaps; advance horn flaps are less than full up and throttle is throttle retarded to idle Any stage Fire warning Continuous Any overheat condition or fire in any engine or 1. Lower the heat in the bell nacelle, or main wheel or nose wheel well, the area where in the APU engine, or any compartment having fire F/W was activated warning system installed 2. Signal may be silenced Whenever the fire warning system is tested pushing the F/W bell cutout switch or the APU cutout switch Any stage Communications High chime Any time captain’s call button is pressed at Release button; if button ATA 2300 external power panel forward or rearward remains locked in, pull cabin attendant’s panel button out Figure 10-130. Aircraft aural warnings. 12 6 8 DAYS 9 3 Figure 10-131. A typical mechanical 8-day aircraft clock. adequate clearance, inspect the shock mounts for looseness, cracks, or deterioration. Instrument panel layout is seemingly random on older aircraft. The advent of instrument flight made the flight instruments of critical importance when flying without outside reference to the horizon or ground. As a result, the basic T arrangement for flight instruments was adopted, as mentioned in the beginning of this chapter. [Figure 10-4] Electronic flight instrument systems and digital cockpit displays have kept the same basic T arrangement for flight instrument and data presentations. The flight instruments and basic T are located directly in front of the pilot and copilot’s seats. Some light 10-80 Airframe structure Instrument panel Bonding strap Figure 10-133. Instrument panel shock mounts. Engine instruments Navigation instruments Flight instruments Basic “T” layout Figure 10-134. Flight instruments directly in front of the pilot, engine instruments to the left and right, and navigation instruments and radios primarily to the right, which is the center of the instrument panel. This arrangement is commonly on light aircraft to be flown by a single pilot. aircraft have only one full set of flight instruments that are located in front of the left seat. The location of engine instruments and navigation instruments varies. Ideally, they should be accessible to both the pilot and copilot. Numerous variations exist to utilize the limited space in the center of the instrument panel and still provide accessibility by the flight crew to all pertinent instruments. On large aircraft, a center pedestal and overhead panels help create more space. On small aircraft, the engine instruments are often moved to allow navigation instruments and radios to occupy the center of the instrument panel. [Figure 10-134] On modern aircraft, EFIS and digital flight information systems reduce panel clutter and allow easier access to all instruments by both crewmembers. Controllable display panels provide the ability to select from pages of information that, when not displayed, are completely gone from view and use no instrument panel space. Instrument Mounting The method of mounting instruments in their respective panels depends on the design of the instrument case. In one design, the bezel is flanged in such a manner that the instrument can be flush mounted in its cutout from the rear of the panel. Integral, self-locking nuts are provided at the rear faces of the flange corners to receive mounting screws from the front of the panel. The flanged-type instrument can also be mounted to the front of the panel. In this case, nut-plates are usually installed in the panel itself. Nonferrous screws are usually used to mount the instruments. There are also instrument mounting systems where the instruments are flangeless. A special clamp, shaped and dimensioned to fit the instrument case, is permanently secured to the rear face of the panel. The instrument is slid into the panel from the front and into the clamp. The clamp’s tightening screw is accessible from the front side of the panel. [Figure 10-135] Regardless of how an instrument is mounted, it should not be touching or be so close as to touch another instrument during the shock of landing. 10-81 Front mounted Rear mounted Clamp mounted Nut plates mounted in instrument Nut plates mounted in panel Strap tightened by clamp Clamp mounted on instrument panel Figure 10-135. Instrument mounts— flanged (top and middle) and flangeless (bottom). new component may require a load check be performed. This is essentially an on the ground operational check to ensure the electrical system can supply all of the electricity consuming devices installed on the aircraft. Follow the manufacturer’s instructions on how to perform this check. Instrument Range Markings Many instruments contain colored markings on the dial face to indicate, at a glance, whether a particular system or component is within a range of operation that is safe and desirable or if an undesirable condition exists. These markings are put on the instrument by the original equipment manufacturer in accordance with the Aircraft Specifications in the Type Certificate Data Sheet. Data describing these limitations can also sometimes be found in the aircraft manufacturer’s operating and maintenance manuals. Occasionally, the aircraft technician may find it necessary to apply these marking to an approved replacement instrument on which they do not appear. It is crucial that the instrument be marked correctly and only in accordance with approved data. The marking may be placed on the cover glass of the instrument with paint or decals. A white slippage mark is made to extend from the glass to the instrument case. Should the glass rotate in the bezel, the marking will no longer be aligned properly with the calibrated instrument dial. The broken slippage mark indicates this to the pilot or technician. The colors used as range markings are red, yellow, green, blue, or white. The markings can be in the form of an arc or a radial line. Red is used to indicate maximum and minimum ranges; operations beyond these markings are dangerous and should be avoided. Green indicates the normal operating range. Yellow is used to indicate caution. Blue and white are used on airspeed indicators to define specific conditions. [Figures 10-136 and 10-137] Maintenance of Instruments and Instrument Systems An FAA airframe and powerplant (A&P) technician is not qualified to do internal maintenance on instruments and related line replaceable units discussed in this chapter. This must be carried out at facilities equipped with the specialized equipment needed to perform the maintenance properly. Qualified technicians with specialized training and intimate knowledge of instruments perform this type of work, usually under repair station certification. However, licensed airframe technicians and A&P technicians are charged with a wide variety of maintenance functions related to instruments and instrument systems. Installation, removal, inspection, troubleshooting, and functional checks Instrument Power Requirements Many aircraft instruments require electric power for operation. Even nonelectric instruments may include electric lighting. Only a limited amount of electricity is produced by the aircraft’s electric generator(s). It is imperative that the electric load of the instruments, radios, and other equipment on board the aircraft does not exceed this amount. Electric devices, including instruments, have power ratings. These show what voltage is required to correctly operate the unit and the amount of amperage it draws when operating to capacity. The rating must be checked before installing any component. Replacement of a component with one that has the same power rating is recommended to ensure the potential electric load of the installed equipment remains within the limits the aircraft manufacturer intended. Adding a component with a different rating or installing a completely 10-82 Instrument Range marking Airspeed indicator White arc bottom Top Green arc bottom Top Blue radial line Yellow arc bottom top Red radial line Carburetor air temperature Green arc Yellow arc Red radial line Cylinder head temperature Green arc Yellow arc Red radial line Manifold pressure gauge Green arc Yellow arc Red radial line Fuel pressure gauge Green arc Yellow arc Red radial line Oil pressure gauge Green arc Yellow arc Red radial line Flap operating range Flaps-down stall speed Maximum airspeed for flaps-down flight Normal operating range Flaps-up stall speed Maximum airspeed for rough air Best single-engine rate-of-climb airspeed Structural warning area Maximum airspeed for rough air Never-exceed airspeed Never-exceed airspeed Normal operating range Range in which carburetor ice is most likely to form Maximum allowable inlet air temperature Normal operating range Operation approved for limited time Never-exceed temperature Normal operating range Precautionary range Maximum permissible manifold absolute pressure Normal operating range Precautionary range Maximum and/or minimum permissible fuel pressure Normal operating range Precautionary range Maximum and/or minimum permissible oil pressure Instrument Range marking Oil temperature gauge Green arc Yellow arc Red radial line Tachometer (reciprocating engine) Green arc Yellow arc Red arc Red radial line Tachometer (turbine engine) Green arc Yellow arc Red radial line Tachometer (helicopter) Engine tachometer Green arc Yellow arc Red radial line Rotor tachometer Green arc Red radial line Torque indicator Green arc Yellow arc Red radial line Exhaust gas temperature indicator (turbine engine) Green arc Yellow arc Red radial line Gas producer N1 tachometer (turboshaft helicopter) Green arc Yellow arc Red radial line Normal operating range Precautionary range Maximum and/or minimum permissible oil temperature Normal operating range Precautionary range Restricted operating range Maximum permissible rotational speed Normal operating range Precautionary range Maximum permissible rotational speed Normal operating range Precautionary range Maximum permissible rotational speed Normal operating range Maximum and minimum rotor speed for power-off operational conditions Normal operating range Precautionary range Maximum permissible torque pressure Normal operating range Precautionary range Maximum permissible gas temperature Normal operating range Precautionary range Maximum permissible rotational speed Figure 10-136. Instrument range markings. are all performed in the field by licensed personnel. It is also a responsibility of the licensed technician holding an airframe rating to know what maintenance is required and to access the approved procedures for meeting those requirements. In the following paragraphs, various maintenance and servicing procedures and suggestions are given. The discussion follows the order in which the various instruments and instrument systems were presented throughout this chapter. This is not meant to represent all of the maintenance required by any of the instruments or instruments systems. The aircraft manufacturer’s and instrument manufacturer’s approved maintenance documents should always be consulted for required maintenance and servicing instructions. FAA regulations must also be observed. Altimeter Tests When an aircraft is to be operated under IFR, an altimeter test must have been performed within the previous 24 months. Title 14 of the Code of Federal Regulations (14 CFR) part 91, section 91.411, requires this test, as well as tests on the pitot static system and on the automatic pressure altitude reporting system. The licensed airframe or A&P mechanic is not qualified to perform the altimeter inspections. They must be conducted by either the manufacturer or a certified repair station. 14 CFR part 43, Appendixes E and F detail the requirements for these tests. 10-83 Figure 10-138. An analog pitot-static system test unit (left) and a digital pitot static test unit (right). AIR SPEED KNOTS 40 60 80 100 120 140 160 180 200 220 240 PEE NOT Figure 10-137. An airspeed indicator makes extensive use of range markings. Pitot-Static System Maintenance and Tests Water trapped in a pitot static system may cause inaccurate or intermittent indications on the pitot-static flight instruments. This is especially a problem if the water freezes in flight. Many systems are fitted with drains at the low points in the system to remove any moisture during maintenance. Lacking this, dry compressed air or nitrogen may be blown through the lines of the system. Always disconnect all pitot-static instruments before doing so and always blow from the instrument end of the system towards the pitot and static ports. This procedure must be followed by a leak check described below. Systems with drains can be drained without requiring a leak check. Upon completion, the technician must ensure that the drains are closed and made secure in accordance with approved maintenance procedures. Aircraft pitot-static systems must be tested for leaks after the installation of any component parts or when system malfunction is suspected. It must also be tested every 24 months if on an IFR certified aircraft intended to be flown as such as called out in 14 CFR section 91.411. Licensed airframe and A&P technicians may perform this test. The method of leak testing depends on the type of aircraft, its pitot-static system, and the testing equipment available. [Figure 10-138] Essentially, a testing device is connected into the static system at the static vent end, and pressure is reduced in the system by the amount required to indicate 1,000 feet on the altimeter. Then, the system is sealed and observed for 1 minute. A loss of altitude of more than 100 feet is not permissible. If a leak exists, a systematic check of portions of the system is conducted until the leak is isolated. Most leaks occur at fittings. The pitot portion of the pitotstatic system is checked in a similar fashion. Follow the manufacturer’s instructions when performing all pitot-static system checks. In all cases, pressure and suction must be applied and released slowly to avoid damage to the aircraft instruments. Pitot-static system leak check units usually have their own built-in altimeters. This allows a functional cross-check of the aircraft’s altimeter with the calibrated test unit’s altimeter while performing the static system check. However, this does not meet the requirements of 14 CFR section 91.411 for altimeter tests. Upon completion of the leak test, be sure that the system is returned to the normal flight configuration. If it is necessary to block off various portions of a system, check to be sure that all blanking plugs, adaptors, or pieces of adhesive tape have been removed. 10-84 Figure 10-139. A magnetic compass with a deviation correction card attached, on which the results of swinging the compass should be recorded. Tachometer Maintenance Tachometer indicators should be checked for loose glass, chipped scale markings, or loose pointers. The difference in indications between readings taken before and after lightly tapping the instrument should not exceed approximately 15 rpm. This value may vary, depending on the tolerance established by the indicator manufacturer. Both tachometer generator and indicator should be inspected for tightness of mechanical and electrical connections, security of mounting, and general condition. For detailed maintenance procedures, the manufacturer’s instructions should always be consulted. When an engine equipped with an electrical tachometer is running at idle rpm, the tachometer indicator pointers may fluctuate and read low. This is an indication that the synchronous motor is not synchronized with the generator output. As the engine speed is increased, the motor should synchronize and register the rpm correctly. The rpm at which synchronization occurs varies with the design of the tachometer system. If the instrument pointer(s) oscillate(s) at speeds above the synchronizing value, determine that the total oscillation does not exceed the allowable tolerance. Pointer oscillation can also occur with a mechanical indication system if the flexible drive is permitted to whip. The drive shaft should be secured at frequent intervals to prevent it from whipping. When installing mechanical type indicators, be sure that the flexible drive has adequate clearance behind the panel. Any bends necessary to route the drive should not cause strain on the instrument when it is secured to the panel. Avoid sharp bends in the drive. An improperly installed drive can cause the indicator to fail to read or to read incorrectly. Magnetic Compass Maintenance and Compensation The magnetic compass is a simple instrument that does not require setting or a source of power. A minimum of maintenance is necessary, but the instrument is delicate and should be handled carefully during inspection. The following items are usually included in an inspection: 1. The compass indicator should be checked for correct readings on various cardinal headings and recompensated if necessary. 2. Moving parts of the compass should work easily. 3. The compass bowl should be correctly suspended on an antivibration device and should not touch any part of the metal container. 4. The compass bowl should be filled with liquid. The liquid should not contain any bubbles or have any discoloration. 5. The scale should be readable and be well lit. Compass magnetic deviation is caused by electromagnetic interference from ferrous materials and operating electrical components in the cockpit. Deviation can be reduced by swinging the compass and adjusting its compensating magnets. An example of how to perform this calibration process is given below. The results are recorded on a compass correction card which is placed near the compass in the cockpit. [Figure 10-139] There are various ways to swing a compass. The following is meant as a representative method. Follow the aircraft manufacturer’s instructions for method and frequency of swinging the magnetic compass. This is usually accomplished at flight hour or calendar intervals. Compass calibration is also performed when a new electric component is added to the cockpit, such as a new radio. A complete list of conditions requiring a compass swing and procedure can be found in FAA Advisory Circular (AC) 43.13-1 (as revised), Chapter 12-37. To swing a compass, a compass rose is required. Most airports have one painted on the tarmac in a low-traffic area where maintenance personnel can work. One can also be made with chalk and a good compass. The area where the compass rose is laid out should be far from any possible electromagnetic disturbances, including those underground, and should remain clear of any ferrous vehicles or large equipment while the procedure takes place. [Figure 10-140] 10-85 Figure 10-140. The compass rose on this airport ramp can be used to swing an aircraft magnetic compass. The aircraft should be in level flight attitude for the compass swing procedure. Tail draggers need to have the aft end of the fuselage propped up, preferably with wood, aluminum, or some other nonferrous material. The aircraft interior and baggage compartments should be free from miscellaneous items that might interfere with the compass. All normal equipment should be on board and turned on to simulate a flight condition. The engine(s) should be running. The basic idea when swinging a compass is to note the deviation along the north-south radial and the east-west radial. Then, adjust the compensating magnets of the compass to eliminate as much deviation as possible. Begin by centering or zeroing the compass’ compensating magnets with a non-ferrous screw driver. Align the longitudinal axis of the aircraft on the N-S radial facing north. Adjust the N-S compensating screw so the indication is 0°. Next, align the longitudinal axis of the aircraft on the E-W radial facing east. Adjust the E-W compensating screw so that the compass indicates 90°. Now, move the aircraft to be aligned with the N-S radial facing south. If the compass indicates 180°, there is no deviation while the aircraft is heading due north or due south. However, this is unlikely. Whatever the southfacing indication is, adjust the N-S compensating screw to eliminate half of the deviation from 180°. Continue around to face the aircraft west on the E-W radial, and use the E-W compensating screw to eliminate half of the west-facing deviation from 270°. Once this is done, return the aircraft to alignment with the N-S radial facing north and record the indication. Up to 10° deviation is allowed. Align the aircraft with the radials every 30° around the compass rose and record each indication on the compass compensation card. Date and sign the card and place it in full view of the pilot near the compass in the cockpit. Vacuum System Maintenance Errors in the indication presented on a vacuum gyroscopic instrument could be the result of any factor that prevents the vacuum system from operating within the design suction limits. Errors can also be caused by problems within the instrument, such as friction, worn parts, or broken parts. Any source that disturbs the free rotation of the gyro at design speed is undesirable resulting in excessive precession and failure of the instruments to maintain accurate indication. The aircraft technician is responsible for the prevention or correction of vacuum system malfunctions. Usually this consists of cleaning or replacing filters, checking and correcting insufficient vacuum, or removing and replacing the vacuum pump or instruments. A list of the most common malfunctions, together with their correction, is included in Figure 10-141. Autopilot System Maintenance The information in this section does not apply to any particular autopilot system, but gives general information that relates to all autopilot systems. Maintenance of an autopilot system consists of visual inspection, replacement of components, cleaning, lubrication, and an operational checkout of the system. Consult the manufacturer’s maintenance manual for all of these procedures. With the autopilot disengaged, the flight controls should function smoothly. The resistance offered by the autopilot servos should not affect the control of the aircraft. The interconnecting mechanisms between the autopilot system and the flight control system should be correctly aligned and smooth in operation. When applicable, the operating cables should be checked for tension. An operational check is important to assure that every circuit is functioning properly. An autopilot operational check should be performed on new installations, after replacement of an autopilot component, or whenever a malfunction in the autopilot is suspected. After the aircraft’s main power switch has been turned on, allow the gyros to come up to speed and the amplifier to warm up before engaging the autopilot. Some systems are designed with safeguards that prevent premature autopilot engagement. While holding the control column in the normal flight position, engage the autopilot system using the switch on the autopilot controller. 10-86 1. No vacuum pressure or insufficient pressure 2. Excessive vacuum 3. Gyro horizon bar fails to respond 4. Turn-and-bank indicator fails to respond 5. Turn-and-bank pointer vibrates Problem and Potential Causes Isolation Procedure Correction Defective vacuum gauge Check opposite engine system on the gauge Replace faulty vacuum gauge Vacuum relief valve incorrectly adjusted Change valve adjustment Make final adjustment to correct setting Vacuum relief valve installed backward Visually inspect Install lines properly Broken lines Visually inspect Replace line Lines crossed Visually inspect Install lines properly Obstruction in vacuum line Check for collapsed line Clean & test line; replace defective part(s) Vacuum pump failure Remove and inspect Replace faulty pump Vacuum regulator valve incorrectly adjusted Make valve adjustment and note pressure Adjust to proper pressure Vacuum relief valve dirty Clean and adjust relief valve Replace valve if adjustment fails Relief valve improperly adjusted Adjust relief valve to proper setting Inaccurate vacuum gauge Check calibration of gauge Replace faulty gauge Instrument caged Visually inspect Uncage instrument Instrument filter dirty Check filter Replace or clean as necessary Insufficient vacuum Check vacuum setting Adjust relief valve to proper setting Instrument assembly worn or dirty Replace instrument No vacuum supplied to instrument Check lines and vacuum system Clean and replace lines and components Instrument filter clogged Visually inspect Replace filter Defective instrument Test with properly functioning instrument Replace faulty instrument Defective instrument Test with properly functioning instrument Replace defective instrument Figure 10-141. Vacuum system troubleshooting guide. After the system is engaged, perform the operational checks specified for the particular aircraft. In general, the checks are as follows: 1. Rotate the turn knob to the left; the left rudder pedal should move forward, and the control column wheel should move to the left and slightly aft. 2. Rotate the turn knob to the right; the right rudder pedal should move forward, and the control column wheel should move to the right and slightly aft. Return the turn knob to the center position; the flight controls should return to the level-flight position. 3. Rotate the pitch-trim knob forward; the control column should move forward. 4. Rotate the pitch-trim knob aft; the control column should move aft. If the aircraft has a pitch-trim system installed, it should function to add down-trim as the control column moves forward and add up-trim as the column moves aft. Many pitch-trim systems have an automatic and a manual mode of operation. The above action occurs only in the automatic mode. Check to see if it is possible to manually override or overpower the autopilot system in all control positions. Center all the controls when the operational checks have been completed. Disengage the autopilot system and check for freedom of the control surfaces by moving the control columns and rudder pedals. Then, reengage the system and check the emergency disconnect release circuit. The autopilot should disengage each time the release button on the control yoke is actuated. When performing maintenance and operational checks on a specific autopilot system, always follow the procedure recommended by the aircraft or equipment manufacturer. LCD Display Screens Electronic and digital instrument systems utilizing LCD technology may have special considerations for the care of the display screens. Antireflective coatings are sometimes used to reduce glare and make the displays more visible. These treatments can be degraded by human skin oils and certain cleaning agents, such as those containing ammonia. It is very important to clean the display lens using a clean, lint-free cloth and a cleaner that is specified as safe for antireflective coatings, preferable one recommended by the aircraft manufacturer. 11-1 Introduction With the mechanics of flight secured, early aviators began the tasks of improving operational safety and functionality of flight. These were developed in large part through the use of reliable communication and navigation systems. Today, with thousands of aircraft aloft at any one time, communication and navigation systems are essential to safe, successful flight. Continuing development is occurring. Smaller, lighter, and more powerful communication and navigation devices increase situational awareness on the flight deck. Coupled with improved displays and management control systems, the advancement of aviation electronics is relied upon to increase aviation safety. Clear radio voice communication was one of the first developments in the use of electronics in aviation. Navigational radios soon followed. Today, numerous electronic navigation and landing aids exist. Electronic devices also exist to assist with weather, collision avoidance, automatic flight control, flight recording, flight management, public address, and entertainment systems. Communication and Navigation Chapter 11 11-2 Figure 11-1. Early voice communication radio tests in 1917. Courtesy of AT&T Archives and History Center. Avionics in Aviation Maintenance Avionics is a conjunction of the words aviation and electronics. It is used to describe the electronic equipment found in modern aircraft. The term “avionics” was not used until the 1970s. For many years, aircraft had electrical devices, but true solid-state electronic devices were only introduced in large numbers in the 1960s. Airframe and engine maintenance is required on all aircraft and is not likely to ever go away. Aircraft instrument maintenance and repair also has an inevitable part in aviation maintenance. The increased use of avionics in aircraft over the past 50 years has increased the role of avionics maintenance in aviation. However, modern, solid-state, digital avionics are highly reliable. Mean times between failures are high, and maintenance rates of avionics systems compared to mechanical systems are likely to be lower. The first decade of avionics proliferation saw a greater increase in the percent of cost of avionics compared to the overall cost of an aircraft. In some military aircraft with highly refined navigation, weapons targeting, and monitoring systems, it hit a high estimate of 80 percent of the total cost of the aircraft. Currently, the ratio of the cost of avionics to the cost of the total aircraft is beginning to decline. This is due to advances in digital electronics and numerous manufacturers offering highly refined instrumentation, communication, and navigation systems that can be fitted to nearly any aircraft. New aircraft of all sizes are manufactured with digital glass cockpits, and many owners of older aircraft are retrofitting digital avionics to replace analog instrumentation and radio navigation equipment. The airframe and powerplant (A&P) maintenance technician needs to be familiar with the general workings of various avionics. Maintenance of the actual avionics devices is often reserved for the avionics manufacturers or certified repair stations. However, the installation and proper operation of these devices and systems remains the responsibility of the field technician. This chapter discusses some internal components used in avionics devices. It also discusses a wide range of common communication and navigational aids found on aircraft. The breadth of avionics is so wide that discussion of all avionics devices is not possible. History of Avionics The history of avionics is the history of the use of electronics in aviation. Both military and civil aviation requirements contributed to the development. The First World War brought about an urgent need for communications. Voice communications from ground-to-air and from aircraft to aircraft were established. [Figure 11-1] The development of aircraft reliability and use for civilian purposes in the 1920s led to increased instrumentation and set in motion the need to conquer blind flight—flight without the ground being visible. Radio beacon direction finding was developed for en route navigation. Toward the end of the decade, instrument navigation combined with rudimentary radio use to produce the first safe blind landing of an aircraft. In the 1930s, the first all radio-controlled blind-landing was accomplished. At the same time, radio navigation using ground-based beacons expanded. Instrument navigation certification for airline pilots began. Low and medium frequency radio waves were found to be problematic at night and in weather. By the end of the decade, use of highfrequency radio waves was explored and included the advent of high-frequency radar. In the 1940s, after two decades of development driven by mail carrier and passenger airline requirements, World War II injected urgency into the development of aircraft radio communication and navigation. Communication radios, despite their size, were essential on board aircraft. [Figure 11-2] Very high frequencies were developed for communication and navigational purposes. Installation of the first instrument landing systems for blind landings began mid-decade and, by the end of the decade, the very high frequency omni-directional range (VOR) navigational network was instituted. It was also in the 1940s that the first transistor was developed, paving the way for modern, solidstate electronics. Civilian air transportation increased over the ensuing decades. Communication and navigation equipment was refined. Solid-state radio development, especially in the 11-3 Figure 11-2. Bomber onboard radio station. 1960s, produced a wide range of small, rugged radio and navigational equipment for aircraft. The space program began and added a higher level of communication and navigational necessity. Communication satellites were also launched. The Cold War military build-up caused developments in guidance and navigation and gave birth to the concept of using satellites for positioning. In the 1970s, concept-validation of satellite navigation was introduced for the military and Block I global positioning system (GPS) satellites were launched well into the 1980s. Back on earth, the long range navigation system (LORAN) was constructed. Block II GPS satellites were commissioned in the mid-80s and GPS became operational in 1990 with the full 24-satellite system operational in 1994. In the new millennium, the Federal Aviation Administration (FAA) assessed the national airspace system (NAS) and traffic projections for the future. Gridlock is predicted by 2022. Therefore, a complete overhaul of the NAS, including communication and navigational systems, has been developed and undertaken. The program is called NextGen. It uses the latest technologies to provide a more efficient and effective system of air traffic management. Heavily reliant on global satellite positioning of aircraft in flight and on the ground, NextGen combines GPS technology with automatic dependant surveillance broadcast technology (ADS-B) for traffic separation. A large increase in air system capacity is the planned result. Overhauled ground facilities accompany the technology upgrades mandated for aircraft. NextGen implementation has started and is currently scheduled through the year 2025. For the past few decades, avionics development has increased at a faster pace than that of airframe and powerplant development. This is likely to continue in the near future. Improvements to solid-state electronics in the form of microand nano-technologies continue to this day. Trends are toward lighter, smaller devices with remarkable capability and reliability. Integration of the wide range of communication and navigational aids is a focus. Fundamentals of Electronics Analog Versus Digital Electronics Electronic devices represent and manipulate real world phenomenon through the use of electrical signals. Electronic circuits are designed to perform a wide array of manipulations. Analog representations are continuous. Some aspect of an electric signal is modified proportionally to the real world item that is being represented. For example, a microphone has electricity flowing through it that is altered when sound is applied. The type and strength of the modification to the electric signal is characteristic of the sound that is made into the microphone. The result is that sound, a real world phenomenon, is represented electronically. It can then be moved, amplified, and reconverted from an electrical signal back into sound and broadcast from a speaker across the room or across the globe. Since the flow of electricity through the microphone is continuous, the sound continuously modifies the electric signal. On an oscilloscope, an analog signal is a continuous curve. [Figure 11-3] An analog electric signal can be modified by changing the signal’s amplitude, frequency, or phase. A digital electronic representation of a real world event is discontinuous. The essential characteristics of the continuous event are captured as a series of discrete incremental values. Electronically, these representative samplings are successive chains of voltage and non-voltage signals. They can be transported and manipulated in electronic circuits. When the samples are sufficiently small and occur with high frequency, real world phenomenon can be represented to appear continuous. 11-4 Volts Time Figure 11-3. An analog signal displayed on an oscilloscope is a continuous curve. Analog signal Digital signal Figure 11-4. Analog signals are continuous voltage modified by all external events including those that are not desired called noise. Digital signals are a series of voltage or no voltage that represent a desired event. Noise A significant advantage of digital electronics over analog electronics is the control of noise. Noise is any alteration of the represented real world phenomenon that is not intended or desired. Consider the operation of a microphone when understanding noise. A continuous analog voltage is modified by a voice signal that results in the continuous voltage varying in proportion to the volume and tone of the input sound. However, the voltage responds and modifies to any input. Thus, background sounds also modify the continuous voltage as will electrostatic activity and circuitry imperfections. This alteration by phenomenon that are not the intended modifier is noise. During the processing of digitized data, there is little or no signal degradation. The real world phenomenon is represented in a string of binary code. A series of ones and zeros are electronically created as a sequence of voltage or no voltage and carried through processing stages. It is relatively immune to outside alteration once established. If a signal is close to the set value of the voltage, it is considered to be that voltage. If the signal is close to zero, it is considered to be no voltage. Small variations or modifications from undesired phenomenon are ignored. Figure 11-4 illustrates an analog sine wave and a digital sine wave. Any unwanted voltage will modify the analog curve. The digital steps are not modified by small foreign inputs. There is either voltage or no voltage. Analog Electronics Early aircraft were equipped with radio communication and navigational devices that were constructed with analog electronic circuits. They used vacuum tubes that functioned as electron control valves. These were later replaced by solidstate devices. Today, digital electronic circuits dominate modern avionics. A brief look at various electron control valves used on aircraft follows. Electron Control Valves Electron control valves are an essential part of an electronic circuit. Control of electron flow enables the circuit to produce the desired outcome. Early aircraft made use of vacuum tubes to control electron flow. Later, transistors replaced vacuum tubes. Semiconductors used in transistors and integrated circuits have enabled the solid-state digital electronics found in aircraft today. Vacuum Tubes Electron control valves found in the analog circuits of early aircraft electronics are constructed of vacuum tubes. Only antique aircraft retain radios with these devices due to their size and inability to withstand the harsh vibration and shock of the aircraft operating environment. However, they do function, and a description is included here as a foundation for the study of more modern electronic circuits and components. Diodes A diode acts as a check valve in an alternating current (AC) circuit. It allows current to flow during half of the AC cycle but not the other half. In this manner, it creates a pulsating direct current (DC) with current that drops to zero in between pulses. A diode tube has two active electrodes: the cathode and the plate. It also contains a heater. All of this is housed in a vacuum environment inside the tube. [Figure 11-5] The 11-5 Plate Plate Heater Cathode Heater Cathode Figure 11-5. A vacuum tube diode contains a cathode, heater, and plate. Note that the arrow formed in the symbol for the heater points to the direction of electron flow. Output waveform Current flow Figure 11-6. A vacuum tube diode in a circuit allows current to flow in one direction only. The output waveform illustrates the lack of current flow as the AC cycles. Anode Cathode Anode Cathode Diode symbol Zener diode symbol Anode Cathode Anode Cathode Tunnel diode symbol LED symbol Anode Cathode Anode Cathode Gate Photodiode symbol SCR symbol Anode Cathode Anode Cathode Vericap symbol Schottkey diode symbol Figure 11-7. Diode symbols. Plate Heater Cathode Grid Heater Cathode Grid Plate Figure 11-8. A triode has three elements: the cathode, plate, and a grid. heater glows red hot while heating the cathode. The cathode is coated with a material whose electrons are excited by the heat. The excited electrons expand their orbit when heated. They move close enough to the plate, which is constructed around the cathode and heater arrangement, that they are attracted to the positively-charged plate. When the AC current cycles, the plate becomes negatively charged and the excited cathode electrons do not flow to the plate. In a circuit, this causes a check valve effect that allows current to only flow in one direction, which is the definition of a diode. [Figure 11-6] The various symbols used to depict diodes are shown in Figure 11-7. Triodes A triode is an electron control valve containing three elements. It is often used to control a large amount of current with a smaller current flow. In addition to the cathode, plate, and heater present in the diode, a triode also contains a grid. The grid is composed of fine wire spiraled between the cathode and the plate but closer to the cathode. Applying voltage to the grid can influence the cathode’s electrons, which normally flow to the plate when the cathode is heated. Changes in the relatively small amount of current that flows through the grid can greatly impact the flow of electrons from the cathode to the plate. [Figure 11-8] Figure 11-9 illustrates a triode in a simple circuit. AC voltage input is applied to the grid. A high-resistance resistor is used so that only minimum voltage passes through to the grid. As this small AC input voltage varies, the amount of DC output in the cathode-plate circuit also varies. When the input signal is positive, the grid is positive. This aids in drawing electrons from the cathode to the plate. However, when the AC input signal cycles to negative, the grid becomes negatively 11-6 Plate Cathode Control grid Screen grid Figure 11-10. A tetrode is a four element electron control valve vacuum tube including a cathode, a plate, a control grid, and a screen grid. Plate CSG RSG Cathode Control grid Input Screen grid Load resistor High resistance resistor Battery − + Figure 11-11. To enable a triode to be used at high frequencies, a screen grid is constructed between the plate and the control grid. Plate CSG RSG Cathode Control grid Suppression grid Input Screen grid Load resistor High resistance resistor Battery − + Figure 11-12. A pentode contains a suppression grid that controls secondary electron emissions from the plate at high power. This keeps the current in the screen grid from becoming too high. Output voltage varying DC Load resistor High resistance resistor Battery Plate − Control grid + Varying AC input Voltage Cathode Figure 11-9. Varying AC input voltage to the grid circuit in a triode produces a varying DC output. charged and flow from the cathode to the plate is cut off with the help of the negatively charged grid that repels the electrons on the cathode. Tetrodes A tetrode vacuum tube electron control valve has four elements. In addition to the cathode, plate, and grid found in a tridode, a tetrode also contains a screen grid. The cathode and plate of a vacuum tube electron control valve can act as a capacitor. At high frequencies, the capacitance is so low that feedback occurs. The output in the plate circuit feeds back into the control grid circuit. This causes an oscillation generating AC voltage that is unwanted. By placing a screen grid between the anode and the control grid windings, this feedback and the inter-electrode capacitive effect of the anode and cathode are neutralized. [Figure 11-10] Figure 11-11 illustrates a tetrode in a circuit. The screen grid is powered by positive DC voltage. The inter-electrode capacitance is now between the screen grid and the plate. A capacitor is located between the screen grid and ground. AC feedback generated in the screen grid goes to ground and does not oscillate. This allows use of the tetrode at higher frequencies than a triode. Pentodes The plate in a vacuum tube can have a secondary emission that must be controlled. When electrons flow from the cathode through the control grid and screen grid to the plate, they can arrive at such high velocity that some bounce off. Therefore, the tendency is for those electrons to be attracted to the positively charged screen grid. The screen grid is not capable of handling large amounts of current without burning up. To solve this problem, a third grid is constructed between the plate and the screen grid. Called a suppression grid, it is charged negatively so that secondary electron flow from the plate is repelled by the negative charge back toward the plate and is not allowed to reach the screen grid. The five element pentode is especially useful in high power circuits where secondary emissions from the plate are high. [Figure 11-12] Solid-State Devices Solid-state devices began replacing vacuum tube electron control valves in the late 1950s. Their long life, reliability, and resilience in harsh environments make them ideal for use in avionics. 11-7 Maximum number of electrons Shell or Orbit Number 1 2 3 4 5 2 8 18 32 50 Figure 11-13. Maximum number of electrons in each orbital shell of an atom. Felium Neon Argon Krypton He Ne Ar Kr Figure 11-14. Elements with full valence shells are good insulators. Most insulators used in aviation are compounds of two or more elements that share electrons to fill their valence shells. Al Cu Ag Au Aluminum Copper Silver Gold Figure 11-15. The valence shells of elements that are common conductors have one (or three) electrons. Semiconductors The key to solid-state electronic devices is the electrical behavior of semiconductors. To understand semiconductors, a review of what makes a material an insulator or a conductor follows. Then, an explanation for how materials of limited conductivity are constructed and some of their many uses is explained. Semiconductor devices are the building blocks of modern electronics and avionics. An atom of any material has a characteristic number of electrons orbiting the nucleus of the atom. The arrangement of the electrons occurs in somewhat orderly orbits called rings or shells. The closest shell to the nucleus can only contain two electrons. If the atom has more than two electrons, they are found in the next orbital shell away from the nucleus. This second shell can only hold eight electrons. If the atom has more than eight electrons, they orbit in a third shell farther out from the nucleus. This third shell is filled with eight electrons and then a fourth shell starts to fill if the element still has more electrons. However, when the fourth shell contains eight electrons, the number of electrons in the third shell begins to increase again until a maximum of 18 is reached. [Figure 11-13] The outer most orbital shell of any atom’s electrons is called the valence shell. The number of electrons in the valence shell determines the chemical properties of the material. When the valence shell has the maximum number of electrons, it is complete and the electrons tend to be bound strongly to the nucleus. Materials with this characteristic are chemically stable. It takes a large amount of force to move the electrons in this situation from one atom valence shell to that of another. Since the movement of electrons is called electric current, substances with complete valence shells are known as good insulators because they resist the flow of electrons (electricity). [Figure 11-14] In atoms with an incomplete valence shell, that is, those without the maximum number of electrons in their valence shell, the electrons are bound less strongly to the nucleus. The material is chemically disposed to combine with other materials or other identical atoms to fill in the unstable valence configuration and bring the number of electrons in the valence shell to maximum. Two or more substances may share the electrons in their valence shells and form a covalent bond. A covalent bond is the method by which atoms complete their valence shells by sharing valence electrons with other atoms. Electrons in incomplete valence shells may also move freely from valence shell to valence shell of different atoms or compounds. In this case, these are known as free electrons. As stated, the movement of electrons is known as electric current or current flow. When electrons move freely from atom to atom or compound to compound, the substance is known as a conductor. [Figure 11-15] Not all materials are pure elements, that is, substances made up of one kind of atom. Compounds occur when two or more different types of atoms combine. They create a new substance with different characteristics than any of the component elements. When compounds form, valence shells and their maximum number of electrons remain the rule of physics. The new compound molecule may either share electrons to fill the valence shell or free electrons may exist to make it a good conductor. Silicon is an atomic element that contains four electrons in its valence shell. It tends to combine readily with itself and form a lattice of silicon atoms in which adjacent atoms share electrons to fill out the valance shell of each to the maximum of eight electrons. [Figure 11-16] This unique symmetric alignment of silicon atoms results in a crystalline structure. 11-8 Si Si Si Si Si Si Si Si Si Si Si Si Si Si Si Si Valence electrons Figure 11-16. The silicon atoms with just the valence shell electrons share these valence electrons with each other. By sharing with four other silicon atoms, the number of electrons in each silicon atom valence shell becomes eight, which is the maximum number. This makes the substance stable and it resists any flow of electrons. Si Si Si Si Si As Si Si Si Si Si Si Si Si As Si Si Si Si Si Si Si Si As Si Si Si Si Si Si Si Si As Si Si Si Free electron Figure 11-17. Silicon atoms doped with arsenic form a lattice work of covalent bonds. Free electrons exist in the material from the arsenic atom’s 5th valence electron. These are the electrons that flow when the semiconductor material, known as N-type or donor material, is conducting. Once bound together, the valence shells of each silicon atom are complete. In this state, movement of electrons does not occur easily. There are no free electrons to move to another atom and no space in the valence shells to accept a free electron. Therefore, silicon in this form is a good insulator. Silicon is a primary material used in the manufacture of semiconductors. Germanium and a few other materials are also used. Since silicon is an insulator, it must be modified to become a semiconductor. The process often used is called doping. Starting with ultra-pure silicon crystal, arsenic, phosphorus, or some other element with five valence electrons in each atom is mixed into the silicon. The result is a silicon lattice with flaws. [Figure 11-17] The elements bond, but numerous free electrons are present in the material from the 5th electron that is part of the valence shell of the doping element atoms. These free electrons can now flow under certain conditions. Thus, the silicon becomes semiconductive. The conditions required for electron flow in a semiconductor are discussed in the following paragraphs. When silicon is doped with an element or compound containing five electrons in its valence shell, the result is a negatively charged material due to the excess free electrons, and the fact that electrons are negatively charged. This is known as an N-type semiconductor material. It is also known as a donor material because, when it is used in electronics, it donates the extra electrons to current flow. Doping silicon can also be performed with an element that has only three valence electrons, such as boron, gallium, or indium. Valence electron sharing still occurs, and the silicon atoms with interspersed doping element atoms form a lattice molecular structure. However, in this case, there are many valence shells where there are only seven electrons and not eight. This greatly changes the properties of the material. The absence of the electrons, called holes, encourages electron flow due to the preference to have eight electrons in all valence shells. Therefore, this type of doped silicon is also semiconductive. It is known as P-type material or as an acceptor since it accepts electrons in the holes under certain conditions. [Figure 11-18] 11-9 Si Si Si Si Si B Si Si Si Si Si Si Si Si B Si Si Si Si Si Si Si Si B Si Si Si Si Si Si Si Si B Si Si Si A “hole” exists because there is no electron in the boron to form covalent bond here. Figure 11-18. The lattice of boron doped silicon contains holes where the three boron valence shell electrons fail to fill in the combined valence shells to the maximum of eight electrons. This is known as P-type semiconductor material or acceptor material. Depletion area Diode symbol P N Holes Electrons − + Potential hill Figure 11-19. A potential hill. arrive and fill the holes in the lattice. As this occurs, more room is available for electrons and holes to move into the area. Pushed by the potential of the battery, electrons and holes continue to combine. The depletion area becomes extremely narrow under these conditions. The potential hill or barrier is, therefore, very small. The flow of current in the electrical circuit is in the direction of electron movement shown in Figure 11-20. In similar circuits where the negative battery terminal is attached to the N-type semiconductor material and the positive terminal is attached to the P-type material, current flows from N-type, or donor material, to P-type receptor material. This is known as a forward biased semiconductor. A voltage of approximately 0.7 volts is needed to begin the current flow over the potential hill. Thereafter, current flow is linear with the voltage. However, temperature affects the ease at which electrons and holes combine given a specific voltage. Combining N- and P-type semiconductor material in certain ways can produce very useful results. A look at various semiconductor devices follows. Semiconductor Diodes A diode is an electrical device that allows current to flow in one direction through the device but not the other. A simple device that can be made from N- and P-type semiconductors is a semiconductor diode. When joined, the junction of these two materials exhibits unique properties. Since there are holes in the P-type material, free electrons from the N-type material are attracted to fill these holes. Once combined, the area at the junction of the two materials where this happens is said to be depleted. There are no longer free electrons or holes. However, having given up some electrons, the N-type material next to the junction becomes slightly positively charged, and having received electrons, the P-type material next to the junction becomes slightly negatively charged. The depletion area at the junction of the two semiconductor materials constitutes a barrier or potential hill. The intensity of the potential hill is proportional to the width of the depletion area (where the electrons from the N-type material have filled holes in the P-type material). [Figure 11-19] The two semiconductors joined in this manner form a diode that can be used in an electrical circuit. A voltage source is attached to the diode. When the negative terminal of the battery is attached to the N-type semiconductor material and the positive terminal is attached to the P-type material, electricity can flow in the circuit. The negative potential of the battery forces free electrons in the N-type material toward the junction. The positive potential of the battery forces holes in the P-type material toward the other side of the junction. The holes move by the rebinding of the doping agent ions closer to the junction. At the junction, free electrons continuously 11-10 Holes Electrons Diode symbol Electron flow Decreased potential hill Decreased Depletion area + − P N Figure 11-20. The flow of current and the P-N junction of a semiconductor diode attached to a battery in a circuit. Diode symbol Increased width of depletion area P N Increased depletion area + − Figure 11-21. A reversed biased condition. Pulsating DC output AC power source Load resistor Semiconductor diode Figure 11-22. A semiconductor diode acts as a check valve in an AC circuit resulting in a pulsating DC output. P N Anode Cathode Electron flow Conventional current flow Figure 11-23. Symbols and drawings of semiconductor diodes. electrons do not flow. A simple AC rectifier circuit containing a semiconductor diode and a load resistor is illustrated in Figure 11- 22. Semiconductor diode symbols and examples of semiconductor diodes are shown in Figure 11-23. NOTE: Electron flow is typically discussed in this text. The conventional current flow concept where electricity is thought to flow from the positive terminal of the battery through a circuit to the negative terminal is sometimes used in the field. Semiconductor diodes have limitations. They are rated for a range of current flow. Above a certain level, the diode overheats and burns up. The amount of current that passes through the diode when forward biased is directly proportional to the amount of voltage applied. But, as mentioned, it is affected by temperature. Figure 11-24 indicates the actual behavior of a semiconductor diode. In practice, a small amount of current does flow If the battery terminals are reversed, the semiconductor diode circuit is said to be reversed biased. [Figure 11-21] Attaching the negative terminal of the battery to the P-type material attracts the holes in the P-type material away from the junction in the diode. The positive battery terminal attached to the N-type material attracts the free electrons from the junction in the opposite direction. In this way, the width of the area of depletion at the junction of the two materials increases. The potential hill is greater. Current cannot climb the hill; therefore, no current flows in the circuit. The semiconductors do not conduct. Semiconductor diodes are used often in electronic circuits. When AC current is applied to a semiconductor diode, current flows during one cycle of the AC but not during the other cycle. The diode, therefore, becomes a rectifier. When it is forward biased, electrons flow; when the AC cycles, 11-11 Voltage Reverse bias Forward bias Avalanche voltage Burn-out current 0.7 Volts Forward current (MA) Leakage Reverse current ( A) current Figure 11-24. A semiconductor diode. Anode Cathode Electron flow IZ RV L 2 D1 VA RS Figure 11-25. A zener diode, when reversed biased, will break down and allow a prescribed voltage to flow in the direction normally blocked by the diode. through a semiconductor diode when reversed biased. This is known as leakage current and it is in the micro amperage range. However, at a certain voltage, the blockage of current flow in a reversed biased diode breaks down completely. This voltage is known as the avalanche voltage because the diode can no longer hold back the current and the diode fails. Zener Diodes Diodes can be designed with a zener voltage. This is similar to avalanche flow. When reversed biased, only leakage current flows through the diode. However, as the voltage is increased, the zener voltage is reached. The diode lets current flow freely through the diode in the direction in which it is normally blocked. The diode is constructed to be able to handle the zener voltage and the resulting current, whereas avalanche voltage burns out a diode. A zener diode can be used as means of dropping voltage or voltage regulation. It can be used to step down circuit voltage for a particular application but only when certain input conditions exist. Zener diodes are constructed to handle a wide range of voltages. [Figure 11-25] Transistors While diodes are very useful in electronic circuits, semiconductors can be used to construct true control valves known as transistors. A transistor is little more than a sandwich of N-type semiconductor material between two pieces of P-type semiconductor material or vice versa. However, a transistor exhibits some remarkable properties and is the building block of all things electronic. [Figure 11-26] As with any union of dissimilar types of semiconductor materials, the junctions of the P- and N- materials in a transistor have depletion areas that create potential hills for the flow of electrical charges. Like a vacuum tube triode, the transistor has three electrodes or terminals, one each for the three layers of semiconductor material. The emitter and the collector are on the outside of the sandwiched semiconductor material. The center material is known as the base. A change in a relatively small amount of voltage applied to the base of the transistor allows a relatively large amount of current to flow from the collector to the emitter. In this way, the transistor acts as a switch with a small input voltage controlling a large amount of current. If a transistor is put into a simple battery circuit, such as the one shown in Figure 11-27, voltage from the battery (EB) forces free electrons and holes toward the junction between the base and the emitter just as it does in the junction of a semiconductor diode. The emitter-base depletion area becomes narrow as free electrons combine with the holes at the junction. Current (IB) (solid arrows) flows through the junction in the emitter-base battery circuit. At the same time, an emitter-collector circuit is constructed with a battery (EC) of much higher voltage in its circuit. Because of the narrow depletion area at the emitter-base junction, current IC is able to cross the collector base junction, flow through emitter-base junction, and complete the collector-emitter battery circuit (hollow arrows). To some extent, varying the voltage to the base material can increase or decrease the current flow through the transistor as the emitter-base depletion area changes width in response to the base voltage. If base voltage is removed, the emitter11- 12 Depletion areas Depletion areas P N P Collector Collector Emitter Emitter Base Base N P N Collector Collector Emitter Emitter Base Base Typical Transistors PNP Transistor NPN Transistor Symbols for transistors used in an electronic circuit diagram Figure 11-26. Typical transistors, diagrams of a PNP and NPN transistor, and the symbol for those transistors when depicted in an electronic circuit diagram. Collector-base Emitter-base depletion area depletion area B E C IB IB IC IE = IB + IC IE = IB + IC IC EB EC + − + − P N P EB EC + − + − IC IB IC Figure 11-27. The effect of applying a small voltage to bias the emitter-base junction of a transistor (top). A circuit diagram for this same transistor (bottom). 11-13 P P N N P P N P N N Anode Cathode Four-Layer Diode Transistor Equivalent Equivalent Schematic Schematic Symbol Anode Cathode Figure 11-28. A four-layer semiconductor diode behaves like two transistors. When breakover voltage is reached, the device conducts current until the voltage is removed. base depletion area becomes too wide and all current flow through the transistor ceases. Current in the transistor circuit illustrated has a relationship as follows: IE = IB + IC. It should be remembered that it is the voltage applied to the base that turns the collector-emitter transistor current on or off. Controlling a large amount of current flow with a small independent input voltage is very useful when building electronic circuits. Transistors are the building blocks from which all electronic devices are made, including Boolean gates that are used to create micro processor chips. As production techniques have developed, the size of reliable transistors has shrunk. Now, hundreds of millions and even billions of transistors may be used to construct a single chip such as the one that powers your computer and various avionic devices. Silicon Controlled Rectifiers Combination of semiconductor materials is not limited to a two-type, three-layer sandwich transistor. By creating a fourlayer sandwich of alternating types of semiconductor material (i.e., PNPN or NPNP), a slightly different semiconductor diode is created. As is the case in a two-layer diode, circuit current is either blocked or permitted to flow through the diode in a single direction. Within a four-layer diode, sometimes known as a Shockley diode, there are three junctions. The behavior of the junctions and the entire four-layer diode can be understood by considering it to be two interconnected three-layer transistors. [Figure 11-28] Transistor behavior includes no current flow until the base material receives an applied voltage to narrow the depletion area at the base-emitter junction. The base materials in the four-layer diode transistor model receive charge from the other transistor’s collector. With no other means of reducing any of the depletion areas at the junctions, it appears that current does not flow in either direction in this device. However, if a large voltage is applied to forward bias the anode or cathode, at some point the ability to block flow breaks down. Current flows through whichever transistor is charged. Collector current then charges the base of the other transistor and current flows through the entire device. Some caveats are necessary with this explanation. The transistors that comprise this four-layer diode must be constructed of material similar to that described in a zener diode. That is, it must be able to endure the current flow without burning out. In this case, the voltage that causes the diode to conduct is known as breakover voltage rather than breakdown voltage. Additionally, this diode has the unique characteristic of allowing current flow to continue until the applied voltage is reduced significantly, in most cases, until it is reduced to zero. In AC circuits, this would occur when the AC cycles. While the four-layer, Shockley diode is useful as a switching device, a slight modification to its design creates a silicon controlled rectifier (SCR). To construct a SCR, an additional terminal known as a gate is added. It provides more control and utility. In the four-layer semiconductor construction, there are always two junctions forward biased and one junction reversed biased. The added terminal allows the momentary application of voltage to the reversed biased junction. All three junctions then become forward biased and 11-14 P P N N Gate Gate Gate Gate P P N P N N Anode Cathode Silicon Controlled Rectifier Transistor Equivalent Equivalent Schematic Schematic Symbol Anode Anode Cathode Cathode Figure 11-29. A silicon controlled rectifier (SCR) allows current to pass in one direction when the gate receives a positive pulse to latch the device in the on position. Current ceases to flow when it drops below holding current, such as when AC current reverses cycle. Gate Anode base-plate Mounting stud Cathode Anode (case) N type (cathode) P type (gate) N type P type (anode) Figure 11-30. Cross-section of a medium-power SCR. current at the anode flows through the device. Once voltage is applied to the gate, the SCR become latched or locked on. Current continues to flow through it until the level drops off significantly, usually to zero. Then, another applied voltage through the gate is needed to reactivate the current flow. [Figures 11-29 and 11-30] SCRs are often used in high voltage situations, such as power switching, phase controls, battery chargers, and inverter circuits. They can be used to produce variable DC voltages for motors and are found in welding power supplies. Often, lighting dimmer systems use SCRs to reduce the average voltage applied to the lights by only allowing current flow during part of the AC cycle. This is controlled by controlling the pulses to the SCR gate and eliminating the massive heat dissipation caused when using resistors to reduce voltage. Figure 11-31 graphically depicts the timing of the gate pulse that limits full cycle voltage to the load. By controlling the phase during which time the SCR is latched, a reduced average voltage is applied. Triacs SCRs are limited to allowing current flow in one direction only. In AC circuitry, this means only half of the voltage cycle can be used and controlled. To access the voltage in the reverse cycle from an AC power source, a triac can be used. A triac is also a four-layer semiconductor device. It differs from an SCR in that it allows current flow in both directions. A triac has a gate that works the same way as in a SCR; however, a positive or negative pulse to the gate triggers current flow in a triac. The pulse polarity determines the direction of the current flow through the device. Figure 11-32 illustrates a triac and shows a triac in a simple circuit. It can be triggered with a pulse of either polarity and remains latched until the voltage declines, such as when the AC cycles. Then, it needs to be triggered again. In many ways, the triac acts as though it is two SCRs connected side by side only in opposite directions. Like an SCR, the timing of gate pulses determines the amount of the total voltage that is allowed to pass. The output waveform if triggered at 90° is shown in Figure 11-32. Because a triac allows current to flow in both directions, the reverse cycle of AC voltage can also be used and controlled. When used in actual circuits, triacs do not always maintain the same phase firing point in reverse as they do when fired with a positive pulse. This problem can be regulated 11-15 Conduction Firing angle angle Balance of waveform applied to load Applied anode To cathode voltage 180° SCR blocks until gate voltage is applied SCR blocks this half cycle 30° firing 90° firing Average voltage Average voltage Shaded area represents voltage applied to the load. The earlier the SCR is fired, the higher the output voltage is. D1 R1 R2 R3 D2 D1 A C Controlled DC output G SCR Output waveform Power source Figure 11-31. Phase control is a key application for SCR. By limiting the percentage of a full cycle of AC voltage that is applied to a load, a reduced voltage results. The firing angle or timing of a positive voltage pulse through the SCR’s gate latches the device open allowing current flow until it drops below the holding current, which is usually at or near zero voltage as the AC cycle reverses. MT1 MT2 Base Main terminal 2 Gate Main terminal 1 Output waveform Figure 11-32. A triac is a controlled semiconductor device that allows current flow in both directions. somewhat through the use of a capacitor and a diac in the gate circuit. However, as a result, where precise control is required, two SCRs in reverse of each other are often used instead of the triac. Triacs do perform well in lower voltage circuits. Figure 11-33 illustrates the semiconductor layering in a triac. NOTE: The four layers of N- and P-type materials are not uniform as they were in previously described semiconductor devices. None the less, gate pulses affect the depletion areas at the junctions of the materials in the same way allowing current to flow when the areas are narrowed. Unijunction Transistors (UJT) The behavior of semiconductor materials is exploited through the construction of numerous transistor devices containing various configurations of N-type and P-type materials. The physical arrangement of the materials in relation to each other yields devices with unique behaviors and applications. The transistors described above having two junctions of P-type and N-type materials (PN) are known as bipolar junction transistors. Other more simple transistors can be fashioned with only one junction of the PN semiconductor materials. These are known as unijunction transistors (UJT). [Figure 11-34] 11-16 N-type N-type N-type P-type Terminal Gate Terminal P-type N-type Figure 11-33. The semiconductor layering in a triac. A positive or negative gate pulse with respect to the upper terminal allows current to flow through the devise in either direction. N-type material E P-type material I junction B1 B2 Figure 11-34. A unijunction transistor (UJT). +10V +8V +6V +4V +2V +0V E B1 B2 +10V Figure 11-35. The voltage gradient in a UJT. The UJT contains one base semiconductor material and a different type of emitter semiconductor material. There is no collector material. One electrode is attached to the emitter and two electrodes are attached to the base material at opposite ends. These are known as base 1 (B1) and base 2 (B2). The electrode configuration makes the UJT appear physically the same as a bipolar junction transistor. However, there is only one PN junction in the UJT and it behaves differently. The base material of a UJT behaves like a resistor between the electrodes. With B2 positive with respect to B1, voltage gradually drops as it flows through the base. [Figure 11-35] By placing the emitter at a precise location along the base material gradient, the amount of voltage needed to be applied to the emitter electrode to forward bias the UJT base-emitter junction is determined. When the applied emitter voltage exceeds the voltage at the gradient point where the emitter is attached, the junction is forward biased and current flows freely from the B1 electrode to the E electrode. Otherwise, the junction is reversed biased and no significant current flows although there is some leakage. By selecting a UJT with the correct bias level for a particular circuit, the applied emitter voltage can control current flow through the device. UJTs transistors of a wide variety of designs and characteristics exist. A description of all of them is beyond the scope of this discussion. In general, UJTs have some advantages over bipolar transistors. They are stable in a wide range of temperatures. In some circuits, use of UJTs can reduce the overall number of components used, which saves money and potentially increases reliability. They can be found in switching circuits, oscillators, and wave shaping circuits. However, four-layered semiconductor thyristors that function the same as the UJT just described are less expensive and most often used. Field Effect Transistors (FET) As shown in the triac and the UJT, creative arrangement of semiconductor material types can yield devices with a variety of characteristics. The field effect transistor (FET) is another such device which is commonly used in electronic circuits. Its N- and P-type material configuration is shown in Figure 11-36. A FET contains only one junction of the two types of semiconductor material. It is located at the gate where it contacts the main current carrying portion of the device. Because of this, when an FET has a PN junction, it is known as a junction field effect transistor (JFET). All FETs operate by expanding and contracting the depletion area at the junction of the semiconductor materials. 11-17 P P Source Gate Drain Source Gate Drain Diffused P-type material N-type silicon bar Channel Figure 11-36. The basic structure of a field effect transistor and its electronic symbol. P substrate Source Oxido layer Gate Body Drain Metal contact N N Figure 11-37. A MOSFET has a metal gate and an oxide layer between it and the semiconductor material to prevent current leakage. One of the materials in a FET or JFET is called the channel. It is usually the substrate through which the current needing to be controlled flows from a source terminal to a drain terminal. The other type of material intrudes into the channel and acts as the gate. The polarity and amount of voltage applied to the gate can widen or narrow the channel due to expansion or shrinking of the depletion area at the junction of the semiconductors. This increases or decreases the amount of current that can flow through the channel. Enough reversed biased voltage can be applied to the gate to prevent the flow of current through the channel. This allows the FET to act as a switch. It can also be used as a voltage controlled resistance. FETs are easier to manufacture than bipolar transistors and have the advantage of staying on once current flow begins without continuous gate voltage applied. They have higher impedance than bipolar transistors and operate cooler. This makes their use ideal for integrated circuits where millions of FETs may be in use on the same chip. FETs come in N-channel and P-channel varieties. Metal Oxide Semiconductor Field Effect Transistors (MOSFETs) and Complementary Metal Oxide Semiconductor (CMOS) The basic FET has been modified in numerous ways and continues to be at the center of faster and smaller electronic component development. A version of the FET widely used is the metal oxide semiconductor field effect transistor (MOSFET). The MOSFET uses a metal gate with a thin insulating material between the gate and the semiconductor material. This essentially creates a capacitor at the gate and eliminates current leakage in this area. Modern versions of the MOSFET have a silicon dioxide insulating layer and many have poly-crystalline silicon gates rather than metal, but the MOSFET name remains and the basic behavioral characteristic are the same. [Figure 11-37] As with FETs, MOSFETs come with N-channels or P-channels. They can also be constructed as depletion mode or enhancement mode devices. This is analogous to a switch being normally open or normally closed. Depletion mode MOSFETs have an open channel that is restricted or closed when voltage is applied to the gate (i.e., normally open). Enhancement mode MOSFETs allow no current to flow at zero bias but create a channel for current flow when voltage is applied to the gate (normally closed). No voltage is used when the MOSFETs are at zero bias. Millions of enhancement mode MOSFETs are used in the construction of integrated circuits. They are installed in complimentary pairs such that when one is open, the other is closed. This basic design is known as complementary MOSFET (CMOS), which is the basis for integrated circuit design in nearly all modern electronics. Through the use of these transistors, digital logic gates can be formed and digital circuitry is constructed. Other more specialized FETs exist. Some of their unique characteristics are owed to design alterations and others to material variations. The transistor devices discussed above use silicon-based semiconductors. But the use of other semiconductor materials can yield variations in performance. Metal semiconductor FETs (MESFETS) for example, are often used in microwave applications. They have a combined metal and semiconductor material at the gate and are typically made from gallium arsenide or indium phosphide. MESFETs are used for their quickness when starting and stopping current flows especially in opposite directions. High electron mobility transistors (HEMT) and pseudomorphic high electron mobility transistors (PHEMT) are also constructed from gallium arsenide semiconductor material and are used for high frequency applications. 11-18 + − Photodiode symbol Simple coil circuit Figure 11-38. The symbol for a photodiode and a photodiode in a simple coil circuit. B C E + − B E C Photo transistor symbol Simple coil circuit Figure 11-39. A photo transistor in a simple coil circuit (bottom) and the symbol for a phototransistor (top). Figure 11-40. Phototransistors. Photodiodes and Phototransistors Light contains electromagnetic energy that is carried by photons. The amount of energy depends on the frequency of light of the photon. This energy can be very useful in the operation of electronic devices since all semiconductors are affected by light energy. When a photon strikes a semiconductor atom, it raises the energy level above what is needed to hold its electrons in orbit. The extra energy frees an electron enabling it to flow as current. The vacated position of the electron becomes a hole. In photodiodes, this occurs in the depletion area of the reversed biased PN junction turning on the device and allowing current to flow. Figure 11-38 illustrates a photodiode in a coil circuit. In this case, the light striking the photodiode causes current to flow in the circuit whereas the diode would have otherwise blocked it. The result is the coil energizes and closes another circuit enabling its operation. A photon activated transistor could be used to carry even more current than a photodiode. In this case, the light energy is focused on a collector-base junction. This frees electrons in the depletion area and starts a flow of electrons from the base that turns on the transistor. Once on, heavier current flows from the emitter to the collector. [Figure 11-39] In practice, engineers have developed numerous ways to use the energy in light photons to trigger semiconductor devices in electronic circuits. [Figure 11-40] Light Emitting Diodes Light emitting diodes (LEDs) have become so commonly used in electronics that their importance may tend to be overlooked. Numerous avionics displays and indicators use LEDs for indicator lights, digital readouts, and backlighting of liquid crystal display (LCD) screens. LEDs are simple and reliable. They are constructed of semiconductor material. When a free electron from a semiconductor drops into a semiconductor hole, energy is given off. This is true in all semiconductor materials. However, the energy released when this happens in certain materials is in the frequency range of visible light. Figure 11-41 is a table that illustrates common LED colors and the semiconductor material that is used in the construction of the diode. NOTE: When the diode is reversed biased, no light is given off. When the diode is forward biased, the energy given off is visible in the color characteristic for the material being used. Figure 11-42 illustrates the anatomy of a single LED, the symbol of an LED, and a graphic depiction of the LED process. 11-19 Infrared Red Orange Yellow Green Blue Violet Purple Ultraviolet White V < 1.9 1.63 < V < 2.03 2.03 < V < 2.10 2.10 < V < 2.18 1.9[32] < V < 4.0 2.48 < V < 3.7 2.76 < V < 4.0 2.48 < V < 3.7 3.1 < V < 4.4 V = 3.5 Gallium arsenide (GaAs) Aluminium gallium arsenide (AlGaAs) Aluminium gallium arsenide (AlGaAs) Gallium arsenide phosphide (GaAsP) Aluminium gallium indium phosphide (AlGaInP) Gallium(III) phosphide (GaP) Gallium arsenide phosphide (GaAsP) Aluminium gallium indium phosphide (AlGaInP) Gallium(III) phosphide (GaP) Gallium arsenide phosphide (GaAsP) Aluminium gallium indium phosphide (AlGaInP) Gallium(III) phosphide (GaP) Indium gallium nitride (InGaN) / Gallium(III) nitride (GaN) Gallium(III) phosphide (GaP) Aluminium gallium indium phosphide (AlGaInP) Aluminium gallium phosphide (AlGaP) Zinc selenide (ZnSe) Indium gallium nitride (InGaN) Silicon carbide (SiC) as substrate Silicon (Si) as substrate — (under development) Indium gallium nitride (InGaN) Dual blue/red LEDs, blue with red phosphor, or white with purple plastic diamond (235 nm)[33] Boron nitride (215 nm)[34][35] Aluminium nitride (AlN) (210 nm)[36] Aluminium gallium nitride (AlGaN) Aluminium gallium indium nitride (AlGaInN) — (down to 210 nm)[37] Blue/UV diode with yellow phosphor Color Wavelength (nm) Voltage (V) Semiconductor Material Figure 11-41. LED colors and the materials used to construct them as well as their wavelength and voltages. P-type N-type Hole Electron Conduction band Valence band Band gap Light + − recombi nation Anode Cathode Expoxy lens/case Reflective cavity Semiconductor die Flat spot Leadframe Anvil Post S An Po R Wire bond Figure 11-42. A close up of a single LED (left) and the process of a semi-conductor producing light by electrons dropping into holes and giving off energy (right). The symbol for a light emitting diode is the diode symbol with two arrows pointing away from the junction. 11-20 Load resistor Diode Transformer AC input AC output + − + 0 − Output waveform Positive half wave Figure 11-43. A half wave rectifier uses one diode to produce pulsating DC current from AC. Half of the AC cycle is wasted when the diode blocks the current flow as the AC cycles below zero. Diode 1 Diode 2 + − VRL Diode 1 Diode 2 + − VRL OV OV OV A B Figure 11-44. A full wave rectifier can be built by center tapping the secondary coil of the transformer and using two diodes in separate circuits. This rectifies the entire AC input into a pulsating DC with twice the frequency of a half wave rectifier. D4 D1 D2 D3 + − + 0 − Output waveform Transformer AC Input Figure 11-45. The bridge-type four-diode full wave rectifier circuit is most commonly used to rectify single-phase AC into DC. Basic Analog Circuits The solid-state semiconductor devices described in the previous section of this chapter can be found in both analog and digital electronic circuits. As digital electronics evolve, analog circuitry is being replaced. However, many aircraft still make use of analog electronics in radio and navigation equipment, as well as in other aircraft systems. A brief look at some of the basic analog circuits follows. Rectifiers Rectifier circuits change AC voltage into DC voltage and are one of the most commonly used type of circuits in aircraft electronics. [Figure 11-43] The resulting DC waveform output is also shown. The circuit has a single semiconductor diode and a load resistor. When the AC voltage cycles below zero, the diode shuts off and does not allow current flow until the AC cycles through zero voltage again. The result is pronounced pulsating DC. While this can be useful, half of the original AC voltage is not being used. A full wave rectifier creates pulsating DC from AC while using the full AC cycle. One way to do this is to tap the secondary coil at its midpoint and construct two circuits with the load resistor and a diode in each circuit. [Figure 11-44] The diodes are arranged so that when current is flowing through one, the other blocks current. When the AC cycles so the top of the secondary coil of the transformer is positive, current flows from ground, through the load resistor (VRL), Diode 1, and the upper half of the coil. Current cannot flow through Diode 2 because it is blocked. [Figure 11-44A] As the AC cycles through zero, the polarity of the secondary coil changes. [Figure 11-44B] Current then flows from ground, through the load resistor, Diode 2, and the bottom half of the secondary coil. Current flow through Diode 1 is blocked. This arrangement yields positive DC from cycling AC with no wasted current. Another way to construct a full wave rectifier uses four semiconductor diodes in a bridge circuit. Because the secondary coil of the transformer is not tapped at the center, the resultant DC voltage output is twice that of the two-diode full wave rectifier. [Figure 11-45] During the first half of the AC cycle, the bottom of the secondary coil is negative. 11-21 D1 D4 D5 D6 D3 D2 Rotor (field) Stator + − + 1 2 3 1 2 3 − Output waveform Load resistor Figure 11-46. A six-diode three-phase AC rectifier. Current flows from it through diode (D1), then through the load resistor, and through diode (D2) on its way back to the top of the secondary coil. When the AC reverses its cycle, the polarity of the secondary coil changes. Current flows from the top of the coil through diode (D3), then through the load resistor, and through diode (D4) on its way back to the bottom of the secondary coil. The output waveform reflects the higher voltage achieved by rectifying the full AC cycle through the entire length of the secondary coil. Use and rectification of three-phase AC is also possible on aircraft with a specific benefit. The output DC is very smooth and does not drop to zero. A six diode circuit is built to rectify the typical three-phase AC produced by an aircraft alternator. [Figure 11-46] Each stator coil corresponds to a phase of AC and becomes negative for 120° of rotation of the rotor. When stator 1 or the first phase is negative, current flows from it through diode (D1), then through the load resistor and through diode (D2) on its way back to the third phase coil. Next, the second phase coil becomes negative and current flows through diode (D3). It continues to flow through the load resistor and diode (D4) on its way back to the first phase coil. Finally, the third stage coil becomes negative causing current to flow through diode (D5), then the load resistor and diode (D6) on its way back to the second phase coil. The output waveform of this three-phase rectifier depicts the DC produced. It is a relatively steady, non-pulsing flow equivalent to just the tops of the individual curves. The phase overlap prevents voltage from falling to zero producing smooth DC from AC. Amplifiers An amplifier is a circuit that changes the amplitude of an electric signal. This is done through the use of transistors. As mentioned, a transistor that is forward biased at the base-emitter junction and reversed biased at the collectorbase junction is turned on. It can conduct current from the collector to the emitter. Because a small signal at the base can cause a large current to flow from collector to emitter, a transistor in itself can be said to be an amplifier. However, a transistor properly wired into a circuit with resistors, power sources, and other electronic components, such as capacitors, can precisely control more than signal amplitude. Phase and impedance can also be manipulated. Since the typical bi-polar junction transistor requires a based circuit and a collector-emitter circuit, there should be four terminals, two for each circuit. However, the transistor only has three terminals (i.e., the base, the collector, and the emitter). Therefore, one of the terminals must be common to both transistor circuits. The selection of the common terminal affects the output of the amplifier. Since the typical bipolar junction transistor requires a base circuit and a collector-emitter circuit, there should be four terminals—two for each circuit. However, the transistor only has three terminals: the base, the collector, and the emitter. Therefore, one of the terminals must be common to both transistor circuits. The selection of the common terminal affects the output of the amplifier. The three basic amplifier types, named for which terminal of the transistor is the common terminal to both transistor circuits, include: 1. Common-emitter amplifier 2. Common-collector amplifier 3. Common-base amplifier Common-Emitter Amplifier The common-emitter amplifier controls the amplitude of an electric signal and inverts the phase of the input signal. Figure 11-47 illustrates a common-emitter amplifier for AC using a NPN transistor and its output signal graph. Common emitter circuits are characterized by high current gain and a 180° voltage phase shift from input to output. It is for the amplification of a microphone signal to drive a speaker. As always, adequate voltage of the correct polarity to the base puts the transistor in the active mode, or turns it on. Then, as the base input current fluctuates, the current through the transistor fluctuates proportionally. However, AC cycles through positive and negative polarity. Every 180°, the transistor shuts 11-22 Vinput E Rload B C Voutput Figure 11-48. A basic common-collector amplifier circuit. Both the input and output circuits share a path through the load and the emitter. This causes a direct relationship of the output current to the input current. Q1 R1 1k AC V1 Vbias 2.3V v(1) I(v(1)) V I(v(1)) −40 mA −20 mA Vinput 1.5V 2 kHz Speaker 8 15 V + − 4.0 2.0 0.0 0 0.0 500.0 1000.0 −2.0 −4.0 Units V Input Output Time uS Figure 11-47. A common-emitter amplifier circuit for amplifying an AC microphone signal to drive a speaker (top) and the graph of the output signal showing a 180 degree shift in phase (bottom). off because the polarity to the base-emitter junction of the transistor is not correct to forward bias the junction. To keep the transistor on, a DC biasing voltage of the correct polarity (shown as a 2.3 volts (V) battery) is placed in series with the input signal in the base circuit to hold the transistor in the active mode as the AC polarity changes. This way the transistor stays in the active mode to amplify an entire AC signal. Transistors are rated by ratio of the collector current to the base current, or Beta (β). This is established during the manufacture of the unit and cannot be changed. A 100 β transistor can handle 100 times more current through a collector-emitter circuit than the base input signal. This current in Figure 11-47 is provided from the 15V battery, V1. So, the amplitude of amplification is a factor of the beta of the transistor and any in-line resistors used in the circuits. The fluctuations of the output signal, however, are entirely controlled by the fluctuations of current input to the transistor base. If measurements of input and output voltages are made, it is shown that as the input voltage increases, the output voltage decreases. This accounts for the inverted phase produced by a common-emitter circuit. [Figure 11-47] Common-Collector Amplifier Another basic type of amplifier circuit is the common-collector amplifier. Common-collector circuits are characterized by high current gain, but vertically no voltage gain. The input circuit and the load circuit in this amplifier share the collector terminal of the transistor used. Because the load is in series with the emitter, both the input current and output current run through it. This causes a directly proportional relationship between the input and the output. The current gain in this circuit configuration is high. A small amount of input current can control a large amount of current to flow from the collector to the emitter. A common collector amplifier circuit is illustrated in Figure 11-48. The base current needs to flow through the PN junction of the transistor, which has about a 0.7V threshold to be turned on. The output current of the amplifier is the beta value of the transistor plus 1. During AC amplification, the common-collector amplifier has the same problem that exists in the common-emitter amplifier. The transistor must stay on or in the active mode regardless of input signal polarity. When the AC cycles through zero, the transistor turns off because the minimum amount of current to forward bias the transistor is not available. The addition of a DC biasing source (battery) in series with the AC signal in the input circuit keeps the transistor in the active mode throughout the full AC cycle. [Figure 11-49] A common-collector amplifier can also be built with a PNP transistor. [Figure 11-50] It has the same characteristics as the NPN common-collector amplifier shown in Figure 11-50. When arranged with a high resistance in the input circuit and a small resistance in the load circuit, the common-collector amplifier can be used to step down the impedance of a signal. [Figure 11-51] Common-Base Amplifier A third type of amplifier circuit using a bipolar transistor is the common-base amplifier. In this circuit, the shared transistor terminal is the base terminal. [Figure 11-52] This causes a unique situation in which the base current is actually larger than the collector or emitter current. As such, the common base amplifier does not boost current as the other amplifiers do. It attenuates current but causes a high gain in voltage. A very small fluctuation in base voltage in the input circuit causes a large variation in output voltage. The effect 11-23 V1 15V Rload 5 k v(1) v(3) 4.0 3.0 2.0 0.0 500.0 1000.0 1.0 0.0 Time uS Vinput 1.5V 2 kHz Vbias 2.3V Figure 11-49. A DC biasing current is used to keep the transistor of a common-collector amplifier in the active mode when amplifying AC (top). The output of this amplifier is in phase and directly proportional to the input (bottom). The difference in amplitude between the two is the 0.7V used to bias the PN junction of the transistor in the input circuit. + − − + Rload v(1) v(3) −1.0 −2.0 −3.0 0.0 500.0 1000.0 −4.0 0.0 Time uS Vinput Figure 11-50. A common-collector amplifier circuit with a PNP transistor has the same characteristics as that of a commoncollector amplifier with a NPN transistor except for reversed voltage polarities and current direction. − + + − RE 50 RB 20K EEB ECE Input C Output Input signal Output signal Figure 11-51. This common-collector circuit has high input impedance and low output impedance. − + − + Voutput Vinput E B C Rload Figure 11-52. A common-base amplifier circuit for DC current. on the circuit output is direct, so the output voltage phase is the same as the input signal but much greater in amplitude. As with the other amplifier circuits, when amplifying an AC signal with a common-base amplifier circuit, the input signal to the base must include a DC source to forward bias the transistor’s base-emitter junction. This allows current to flow from the collector to the emitter during both cycles of the AC. A circuit for AC amplification is illustrated in Figure 11-53 with a graph of the output voltage showing the large increase produced. The common-base amplifier is limited in its use since it does not increase current flow. This makes it the least used configuration. However, it is used in radio frequency amplification because of the low input Z. Figure 11-54 summarizes the characteristics of the bipolar amplifier circuits discussed above. NOTE: There are many variations in circuit design. JFETs and MOSFETS are also used in amplifier circuits, usually in small signal amplifiers due to their low noise outputs. 11-24 Common-emitter Common-collector Common-base Input: fairly high Output: fairly high Input: high Output: low Input: low Output: high Relatively large Always less than one Large Large Relatively large Relatively large Inverts phase Output same as input Output same as input Relatively large Relatively large Always less than one Type of Amplifier Impedance Voltage Gain Current Gain Power Gain Phase Figure 11-54. PN junction transistor amplifier characteristics. S N 90° 180° 0° 0° 60° 120° 120° 240°300° 360° 1.0 .5 .5 .87 1.0 .87 270° Figure 11-55. Voltage over time of sine waveform electricity created when a conductor is rotated through a uniform magnetic field, such as in a generator. Square wave + − Figure 11-56. The waveform of pulsing DC is a square wave. − + − + Voutput Vinput Q1 Rload R1100 5.0k V1 15V Vinput 0.12Vpp 0Voffset 2 kHz v(4) I0*V(5.2) 15.0 10.0 5.0 0.0 500.0 1000.0 0.0 −5.0 Time uS Figure 11-53. In a common-base amplification circuit for AC (top), output voltage amplitude is greatly increased in phase with the input signal (bottom). Oscillator Circuits Oscillators function to make AC from DC. They can produce various waveforms as required by electronic circuits. There are many different types of oscillators and oscillator circuits. Some of the most common types are discussed below. A sine wave is produced by generators when a conductor is rotated in a uniform magnetic field. The typical AC sine wave is characterized by a gradual build-up and decline of voltage in one direction, followed by a similar smooth buildup to peak voltage and decline to zero again in the opposite direction. The value of the voltage at any given time in the cycle can be calculated by taking the peak voltage and multiplying it by the sine of the angle through which the conductor has rotated. [Figure 11-55] A square wave is produce when there is a flow of electrons for a set period that stops for a set amount of time and then repeats. In DC current, this is simply pulsing DC. [Figure 11-56] This same wave form can be of opposite polarities when passed through a transformer to produce AC. Certain oscillators produce square waves. An oscillator known as a relaxation oscillator produces another kind of wave form, a sawtooth wave. A slow rise from zero to peak voltage is followed by a rapid drop-off of voltage back to almost zero. Then it repeats. [Figure 11-57] In the circuit, a capacitor slowly charges through a resistor. A neon bulb is wired across the capacitor. When its ignition voltage is reached, the bulb conducts. This short-circuits the charged capacitor, which causes the voltage to drop to nearly zero and the bulb goes out. Then, the voltage rises again as the cycle repeats. 11-25 Ignition Rise rate determined by resistor voltage 0 E + + − C NE Resistor Figure 11-57. A relaxation oscillator produces a sawtooth wave output. A C B + − + 0 − Damped oscillation caused by resistance in the circuit Battery Figure 11-58. A tank circuit alternately charges opposite plates of a capacitor through a coil in a closed circuit. The oscillation is an alternating current that diminishes due to resistance in the circuit. RA RB S RFC RE C2 C1 L2 L1 Figure 11-59. A Hartley oscillator uses a tank circuit and a transistor to maintain oscillation whenever power is applied. Electronic Oscillation Oscillation in electronic circuits is accomplished by combining a transistor and a tank circuit. A tank circuit is comprised of a capacitor and coil parallel to each other. [Figure 11-58] When attached to a power source by closing switch A, the capacitor charges to a voltage equal to the battery voltage. It stays charged, even when the circuit to the battery is open (switch in position B). When the switch is put in position C, the capacitor and coil are in a closed circuit. The capacitor discharges through the coil. While receiving the energy from the capacitor, the coil stores it by building up an electromagnetic field. When the capacitor is fully discharged, the coil stops conducting. The magnetic field collapses, which induces current flow. The current charges the opposite plate of the capacitor. When completely charged, the capacitor discharges into the coil again. The magnetic field builds again and stops when the capacitor is fully discharged. The magnetic field collapses again, which induces current that charges the original plate of the capacitor and the cycle repeats. This oscillation of charging and discharging the capacitor through the coil would continue indefinitely if a circuit could be built with no resistance. This is not possible. However, a circuit can be built using a transistor that restores losses due to resistance. There are various ways to accomplish this. The Hartley oscillator circuit in Figure 11-59 is one. The circuit can oscillate indefinitely as long as it is connected to power. When the switch is closed, current begins to flow in the oscillator circuit. The transistor base is supplied with biasing current through the voltage divider RA and RB. This allows current to flow through the transistor from the collector to the emitter, through RE and through the lower portion of the center tapped coil that is labeled L1. The current increasing through this coil builds a magnetic field that induces current in the upper half of the coil labeled L2 . The current from L2 charges capacitor C2, which increases the forward bias of the transistor. This allows an increasing flow of current through the transistor, RE, and L1 until the transistor is saturated and capacitor C1 is fully charged. Without force to add electrons to capacitor C1, it discharges and begins the oscillation in the tank circuit described in the previous section. As C1 becomes fully charged, current to charge C2 reduces and C2 also discharges. This adds the energy needed to the tank circuit to compensate for resistance losses. As C2 is discharging, it reduces forward biasing and eventually the transistor becomes reversed biased and cuts off. When the opposite plate of capacitor C1 is fully charged, it discharges and the oscillation is in progress. The transistor base becomes forward biased again, allowing for current flow through the resistor RE, coil L1, etc. 11-26 A B A 1 0 B 0 1 The NOT gate Figure 11-61. A NOT logic gate symbol and a NOT gate truth table. RB S RFC C2 C3 C1 L Crystal Figure 11-60. A crystal in an electronic oscillator circuit is used to tune the frequency of oscillation. The frequency of the AC oscillating in the Hartley oscillator circuit depends on the inductance and capacitance values of the components used. Use of a crystal in an oscillator circuit can control the frequency more accurately. A crystal vibrates at a single, consistent frequency. When flexed, a small pulse of current is produced through the piezoelectric effect. Placed in the feedback loop, the pulses from the crystal control the frequency of the oscillator circuit. The tank circuit component values are tuned to match the frequency of the crystal. Oscillation is maintained as long as power is supplied. [Figure 11-60] Other types of oscillator circuits used in electronics and computers have two transistors that alternate being in the active mode. They are called multi-vibrators. The choice of oscillator in an electronic device depends on the exact type of manipulation of electricity required to permit the device to function as desired. Digital Electronics The above discussion of semiconductors, semiconductor devices, and circuitry is only an introduction to the electronics found in communications and navigation avionics. In-depth maintenance of the interior electronics on most avionics devices is performed only by certified repair stations and trained avionics technicians. The airframe technician is responsible for installation, maintenance, inspection, and proper performance of avionics in the aircraft. Modern aircraft increasingly employs digital electronics in avionics rather than analog electronics. Transistors are used in digital electronics to construct circuits that act as digital logic gates. The purpose and task of a device is achieved by manipulating electric signals through the logic gates. Thousands, and even millions, of tiny transistors can be placed on a chip to create the digital logic landscape through which a component’s signals are processed. Digital Building Blocks Digital logic is based on the binary number system. There are two conditions than may exist, 1 or 0. In a digital circuit, these are equivalent to voltage or no voltage. Within the binary system, these two conditions are called Logic 1 and Logic 0. Using just these two conditions, gates can be constructed to manipulate information. There are a handful of common logic gates that are used. By combining any number of these tiny solid-state gates, significant memorization, manipulation, and calculation of information can be performed. The NOT Gate The NOT gate is the simplest of all gates. If the input to the gate is Logic 1, then the output is NOT Logic 1. This means that it is Logic 0, since there are only two conditions in the binary world. In an electronic circuit, a NOT gate would invert the input signal. In other words, if there was voltage at the input to the gate, there would be no output voltage. The gate can be constructed with transistors and resistors to yield this electrical logic every time. (The gate or circuit would also have to invert an input of Logic 0 into an output of Logic 1.) To understand logic gates, truth tables are often used. A truth table gives all of the possibilities in binary terms for each gate containing a characteristic logic function. For example, a truth table for a NOT gate is illustrated in Figure 11-61. Any input (A) is NOT present at the output (B). This is simple, but it defines this logic situation. A tiny NOT gate circuit can be built using transistors that produce these results. In other words, a circuit can be built such that if voltage arrives at the gate, no voltage is output or vice-versa. When using transistors to build logic gates, the primary concern is to operate them within the circuits so the transistors are either OFF (not conducting) or fully ON (saturated). In this manner, reliable logic functions can be performed. The variable voltage and current situations present during the active mode of the transistor are of less importance. 11-27 Input Output D1 D2 Q1 Q2 VCC Q3 R1 R2 R3

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