CHAPTER 10 Position and Warning Systems

POSITION AND WARNING SYSTEMS

INTRODUCTION Pilots of today's complex aircraft can no longer fly by the seats of their pants. The pilot receives indications of what the aircraft is doing through instruments and warning systems. These include airspeed indicators, unsafe system warnings, and remote position indicators. Some systems, such as antiskid brake systems, allow the pilot to obtain maximum performance, which may be impossible without mechanical assistance. This section covers some of these systems and the hardware necessary to operate them.

ANTISKID BRAKE CONTROL SYSTEMS method only works well when the control valves are capable of operating very quickly. [Figure 10-1]

It is important that a pilot avoid excessive braking to prevent skidding and loss of control. With a tail-wheel-type airplane, too much braking could result in a nose-over or ground loop. With large-diameter tires on small wheels, heavy braking could cause the tire to slip on the rim and pull the valve out of the tube. Modern high-speed jet aircraft usually have more than one wheel on each side, and all of the brakes on one side are controlled with one pedal. With this arrangement, the pilot has no way of knowing when one of these wheels begins to skid. Without prompt corrective action to release a locked-up wheel, the tire is likely to blow out and damage the aircraft, or in severe cases, result in loss of control. Friction created by the brakes reduces the wheel rotation rate, and friction between the tire and the runway slows the aircraft. If the tire rotation slows too rapidly, the tire will begin to slip on the runway instead of gripping it. Once the tire begins to slip, a skid soon develops and braking effectiveness decreases rapidly to near zero. For maximum brake effectiveness, only enough brake pressure should be applied to cause the tire to reach the point where it just begins to slip. This produces the maximum deceleration rate. Maintaining this optimum friction is not easy. As the airplane slows, less brake pressure is needed to maintain the correct balance. Contamination such as water, snow or ice on the runway reduces the coefficient of friction between the tire and the runway. This, too, complicates the problem of maintaining the right amount of brake pressure to achieve maximum braking without excessive tire slippage.

SYSTEM OPERATION You use a simple form of manual antiskid control when driving on ice. For the most effective stopping, you pump the brakes. They are applied only enough to slow the wheel, then released before the wheel decelerates enough to lock up. This same on-and-off type of operation was employed in some of the early aircraft antiskid systems. However, this

Figure 10-1. This graph shows the wheel speed relative to the amount of brake pressure applied manually by the pilot of an aircraft.

In figure 10-1, the brakes are applied and the pressure rises until the wheel starts to slip, but not skid, at point A. This is the ideal condition, but the pilot, having no indication that a slip has been reached, continues to increase the force on the brake pedal. Sufficient pressure is soon reached to produce enough friction in the brake to cause the tire to start to skid on the runway, as shown at point B. The wheel now decelerates fast enough to be felt, so the pilot reduces pressure on the pedal. Since the braking force that is needed lessens as the wheel slows, the wheel continues to decelerate even though the brake pressure decreases. At point C, the wheel has completely locked up, even though the pressure continues to drop. At point D, the pressure is low enough for the friction between the tire and the runway surface to start the wheel rotating again, and soon after, the brake pressure drops to zero. The wheel then comes back up to speed. A successful antiskid system requires two features that early on-and-off systems did not have. There must be some form of wheel-speed sensor that can detect a change in the rate of deceleration and send a signal for the pressure to be released before the wheel

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gets deep into a skid. A valve is also needed that acts quickly enough to prevent all of the pressure from being released before the next application of the brake. This controlled amount of retained pressure prevents the brake-return system from pulling the pressure plate all of the way back, and allows the brakes to reapply almost immediately. The modern modulated antiskid system provides the fastest wheel-speed recovery and produces the shortest stopping distance on any kind of runway surface. When the pilot wants to stop the aircraft in the shortest distance possible, it is necessary to depress the brake pedals all the way to induce maximum braking. All of the brakes receive the maximum pressure. If any wheel should decelerate at a rate indicating an impending skid, some of the pressure to that brake is dumped into the system-return manifold. The control circuit then measures the amount of time required for the wheel to spin back up and applies a slightly reduced pressure to the brake. This reduced pressure is determined by the time required for the spin-up. If this reduced pressure again causes a skid to develop, the cycle is repeated. Some pressure is maintained in the wheel cylinders to prevent the pressure plate from moving all of the way back. This application and release process continues with progressively decreasing pressure until the wheel is held in the slip area, but not allowed to decelerate fast enough to produce a skid. It produces the proper amount of braking for any runway surface condition, with the pilot having only to apply a hard, steady pressure to the brake pedal. When the airplane slows down to approximately 20 miles per hour (m.p.h.) and there is no further danger of skidding, the antiskid system automatically deactivates. This gives the pilot full control of the brakes for maneuvering and parking. As with most auxiliary systems in modern aircraft, the antiskid systems have built-in test circuits, and may be deactivated in the event of a malfunction to give the pilot normal braking but no antiskid protection.

Figure 10-2. A typical antiskid brake system consists of wheel-speed sensors on each main wheel, a control unit, and control valves for each brake.

WHEEL-SPEED SENSORS There are two types of systems in use, an AC system and a DC system. They are essentially alike except for the wheel-speed sensors and one circuit in the control unit. The AC sensor is a variable-reluctance AC generator in the axle of the landing gear that uses a permanent magnet surrounded by a pickup coil. The outside of this sensor has four equally spaced poles with teeth cut into their periphery. A soft iron exciter ring with internal teeth is mounted in the hubcap of the wheel so that it rotates around the sensor. The two sets of teeth are separated by a small gap, and as the exciter ring rotates, the teeth approach each other and then move apart. As the distance between the teeth changes, the reluctance of the magnetic circuit is alternately increased and decreased. This causes the amount of magnetic flux cutting across the pickup coil to change and induces an alternating current in the coil. The faster the wheel turns, the higher the frequency of the induced current. [Figure 10-3]

Many large jet-transports have an auto-brake feature that works in conjunction with the antiskid system. When the system senses weight on the main wheels, it automatically applies the brakes to produce one of several pilot-selected levels of deceleration. This results in a more immediate application of the wheel brakes and maximizes the use of the antiskid system. The pilot can override and disarm the auto-brake system by applying manual brakes.

SYSTEM COMPONENTS An antiskid system consists of three basic components: wheel-speed sensors, an antiskid computer, and control valves. [Figure 10-2]

Figure 10-3. The AC wheel-speed sensor creates a variable frequency AC current. The control unit converts the varying frequency AC into a DC signal voltage that is proportional to the frequency of the AC current.

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The DC sensor is essentially a small, permanent-magnet direct-current generator, which produces a voltage output directly proportional to the rotational speed of its armature. With this type of sensor, there is no need for a converter in the control unit. There also is less danger of interference with the brakes due to the induction of stray voltage into the sensing system. [Figure 10-4]

Figure 10-5. The antiskid control unit operates a brake control valve.

Figure 10-4. The DC wheel-speed sensor does not require an AC-DC converter in the control unit because it generates a direct current proportional to wheel speed. The shaft of the armature is fitted with a blade driven by a bracket in the wheel hubcap and rotates with the wheel. The generator output usually is in the range of one volt for each ten m.p.h. of wheel speed.

CONTROL VALVES A three-port antiskid control valve is located in the pressure line between the brake valve and the brake cylinder, with the third line connecting the control valve to the system-return manifold. During normal operation of the brakes, with no indication of a skid, the valve serves only as a passage and allows the brake fluid to flow into and out of the brake. When a wheel begins to decelerate fast enough to cause a skid, the control unit detects the changing output voltage of the wheel-speed sensor. The control unit sends a DC signal to the control valve, which closes off the pressure port and opens the passage between the brake and the system return. This rapidly operating valve maintains an output pressure that is directly proportional to the amount of signal current from the control unit. [Figure 10-5] The DC signal from the control unit flows through a coil around the armature of the flapper valve. This armature is free to pivot and is centered between two permanent magnets. [Figure 10-6] When the signal from the control unit indicates that no skid is impending, and the braking action should be normal, the magnetic field of the coil reacts with the fields of the permanent magnets and holds the flapper centered between the nozzles. [Figure 10-7]

Figure 10-6. A direct-current signal from the control unit energizes the coil on the armature of the flapper valve, and the movement of the flapper changes the pressure drop across the fixed orifices.

Fluid from the brake valve flows through the filter and discharges equally from each nozzle. Since the amount of flow is the same through each orifice, the pressure drop across the orifices will be the same, and the second-stage spool valve will assume a position that allows free passage between the brake valve and the brake. When the control unit receives a signal from the wheel-speed sensor indicating an impending skid, it sends current through the coil of the armature to polarize it. This causes the flapper to pivot and unbalance the flow from the nozzles. In figure 10-8, the flapper has moved over, restricting the flow

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Figure 10-7. When the flapper is centered between the nozzles, the pressure-drops across orifices O, and O 2 are equal, resulting in output pressure P 1 equaling P 2.

Figure 10-8. When the armature of the flapper valve is energized, the flapper moves over and restricts the flow through orifice O 1 while increasing it through O 2 . The increased pressure drop across O2 causes P 1 to be greater than P2 .

from the left nozzle and opening the flow from the one on the right. There is now more flow through orifice O 2 and therefore a greater pressure drop across it, leaving P a greater than P 2 . This imbalance of pressures moves the second-stage spool over, shutting off the flow of fluid from the brake valve to the brake, and opening a passage from the brake to the return manifold.

vents the pilot from landing the brakes applied. [Figure 10-9]

with

The extremely fast reaction time of this valve allows it to maintain a pressure at the brake that is directly proportional to the amount of current flowing in the armature coil. CONTROL UNIT The control unit has three main functions: to generate electrical signals usable by the control valve; to regulate brake pressure to prevent a skid during landing deceleration; and to prevent application of brake pressure prior to touchdown. Before the airplane touches down, the locked-wheel detector sends a signal into the amplifier, which causes the control valve to open the passage between the brakes and the system-return manifold. This pre-

Figure 10-9. The locked-wheel detector receives a signal from the squat switch, which indicates whether the aircraft is airborne or on the ground. If airborne, the circuitry prevents the brakes from being applied before touchdown.

As soon as the airplane touches down, the squat switch registers that weight is on the wheels. The wheels start to spin up, and at approximately 20 m.p.h., generates enough voltage in the wheel-speed sensor to signal the locked-wheel detector. The

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detector then removes the touchdown control signal from the amplifier. This allows the control valve to apply full pressure to the brakes. [Figure 10-10]

Position and Warning Systems

this pressure to increase slowly until another skid starts to occur, repeating the cycle. When the aircraft is on a wet or icy runway, the antiskid system holds the wheels in the slip region. However, the locked-wheel detector activates whenever one wheel hydroplanes or hits ice and slows down to less than ten m.p.h. while its mated reference wheel still rotates faster than 20 m.p.h. A timer measures the duration of the skid detector signal. If it is more than one-tenth of a second, it sends a "full dump" signal that holds the valve in the full-dump position until the wheel spins back up above ten m.p.h.

Figure 10-10. On touchdown, the squat switch removes the ground from the locked-wheel arming circuit, and the wheel-speed sensor generates a signal which allows the control valve to send full pressure to the brakes.

When the airplane is on the ground and the wheels are rotating at more than 20 m.p.h., the skid detector and modulator provide almost all of the antiskid control. [Figure 10-11]

When all of the wheels are turning at less than 20 m.p.h., the locked-wheel arming circuit disarms, giving the pilot full braking action for low-speed taxiing and parking. [Figure 10-12]

Figure 10-12. When the airplane is on the ground and all three wheels are rotating less than 20 miles per hour, the locked-wheel arming circuit is inoperative and the pilot has full brake control for low speed taxiing and parking.

Figure 10-11. When the airplane is on the ground and all wheels are rotating more than 20 miles per hour, the skid detector and the modulator provide signals for the amplifier.

A deceleration threshold is designed into the skid detector circuit. The reference normally is set to about 20 feet per second, with a wheel speed that is at least six m.p.h. below the speed of the airplane. When a wheel decelerates at a rate greater than this threshold value, the skid detector signals the amplifier and then the control valve to reduce the brake pressure. It also signals the modulator, which automatically establishes the amount of current that will continue to flow through the valve after the wheel has recovered from the skid. When the amplifier receives its signal from the modulator, it maintains this current, which is just enough to position the flapper to prevent the pressure from being completely released. The applied current maintains a pressure slightly less than that which caused the skid. A timer circuit in the modulator then allows

The control unit for antiskid systems using AC sensors operates in the same way as those using DC generators, the only difference being the addition of a converter circuit. This circuit receives the varying-frequency alternating current and converts it into a varying voltage of direct current. The changes in the DC voltage exactly follow the frequency changes of the AC. [Figure 10-13]

SYSTEM TESTS Because it is vitally important that a pilot know the exact condition of the brake system before using it, antiskid systems include test circuits and control switches. These allow the pilot to test the entire system, and if any faults are found, disable the system without affecting normal braking action. There is an anti-skid warning light in the flight deck to warn pilots whenever the system is off or has failed. GROUND TEST The integrity of the antiskid system can be tested on the ground before flight. The pilot turns on the anti-

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tem before condemning the antiskid system. If the brakes are spongy, remove the air by bleeding them. Carefully check for warped disks, malfunctioning return systems, and any indications of damage.

Figure 10-13. The difference between the control unit of an antiskid system using an AC wheel-speed sensor and one using a DC sensor is in the converter between the sensor and the control circuit.

skid control switch and presses the brake pedal. Both the left and right brake lights should illuminate, indicating that all of the pressure from the brake valves is being routed to the brakes. With the brakes still applied, the pilot presses the test switch and holds it for a few seconds. This sends a signal through the wheel-speed sensors into the control unit to simulate a wheel speed of more than 20 m.p.h. The lights should remain on. When the test switch is released, the two brake lights should go out and stay out for a couple of seconds, then come back on. This simulates a wheel lockup that causes a release of, then restoration of, pressure. This test checks the continuity of all of the wiring and operation of the locked-wheel circuits, amplifiers, and control valves. These procedures vary with aircraft type. Consult the appropriate manuals to determine the correct procedure for your aircraft.

Inspection and maintenance of antiskid systems requires logical troubleshooting to locate faults. Due to the complexity of the components, they are usually returned to the manufacturer or a repair station for any needed repairs. If one of the tests shows a malfunction in the system, the most logical place to start troubleshooting is with the wheel-speed sensor. WHEEL-SPEED SENSOR Some DC ■wheel-speed sensors can be checked on

the airplane by removing the wheel hubcap to expose the blade of the sensor. With your finger, give the blade a sharp spin in its normal direction of rotation with the brakes applied and the antiskid switch on. It ■will not turn more than 180 degrees. It is not the amount of rotation that is important, but the rate at which it is turned. If the system is operating properly, the brakes should momentarily release and then reapply. Watch the brake disk stack for relaxation then tightening, this will confirm proper system operation. If this "tweak" test does not cause the brakes to release, consult the maintenance manual for the specific type of airplane on ■which you are working to determine the correct test procedures. [Figure

IN-FLIGHT TEST The antiskid system is included in the pilot's pre-landing checklist. With the airplane configured for landing, the pilot depresses the brake pedals. The brake lights should remain off, which indicates the control valves are holding the brakes in the fully released position. The pilot then presses the test switch, which should illuminate the brake lights for as long as the switch is held down. The test switch sends a signal through the wheel speed sensors, simulating a wheel speed greater than 20 m.p.h. If the system is operating properly, the control valve will direct normal pressure to the brake. SYSTEM MAINTENANCE If a flight crew reports an antiskid or brake malfunction, verify that there is no air in the brake sys-

Figure 10-14. When the blade of the wheel-speed sensor is flipped, it should cause the brakes to release and then reapply.

10-14]

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CONTROL UNIT

CONTROL VALVE

The control unit, shown in figure 10-15, may be checked using a substitution method. Remove both of the connector plugs from the box and swap them left to right. For example, suppose the trouble indication was originally on the left side of the airplane. If the leads from the box are switched and the indication remains on the left side, the trouble is probably not with the control unit. However, if the indication moves to the right side, the control unit may be defective. Any time you switch the leads, be sure to reinstall them on the proper receptacles and properly secure them before returning the aircraft to service.

If the trouble remains after checking the two devices that were the easiest to access, all that remains in the antiskid system is the control valve. These valves are electrohydraulic, and the trouble could be in either the electrical or hydraulic section.

Figure 10-15. The two leads on the antiskid control unit may be switched as a part of the troubleshooting procedure.

The easiest check is the electrical resistance of the coil. Remove the connector plug and measure the resistance of the coil with an accurate ohmmeter. It should measure within the tolerance specified in the service manual. If the trouble is traced to the control valve and is not electrical, the valve must be removed. The problem is probably in the hydraulic portion of the valve. The extremely close tolerances used in the manufacture of this valve make the use of absolutely clean fluid imperative. A fifteen-micron steel-mesh screen is commonly installed in the line before the orifices to insure that no contaminants reach the inside of the valve. If this screen clogs, the valve may malfunction. Check the manufacturer's service manuals to see if it is possible to replace this filter in the field. If it is allowed, follow the service instructions carefully. If any field servicing is allowed on the valve, it must be done in an area free from contamination. Again, be certain to follow the manufacturer's latest service information.

INDICATING AND WARNING SYSTEMS STALL WARNING INDICATOR A stall is a flight condition where the airflow over the upper surface of the wing separates and becomes turbulent. It occurs when the aircraft reaches a critically high angle-of-attack (AOA). If an airplane does not provide sufficient aerodynamic warning of an impending stall, such as buffeting, the pilot must be warned through some other means. Small general aviation aircraft usually use an audible tone or a red light. Many high-performance aircraft use a stick shaker, which vibrates the control column, or which may even force the column forward to reduce the angle-of-attack. Many stall warning systems, particularly on lower performance aircraft, measure the movement of the stagnation point on the wing. The stagnation point marks the particular location on the leading edge of an airfoil where the air separates, some passing over the top of the surface and the rest passing below it. As the angle-of-attack increases, the stagnation point moves down toward the lower surface. The stagnation point is always in the same location when the airflow over the surface becomes turbulent, indicating the approach to a stall. ELECTRIC STALL WARNING An electrically operated stall warning system uses a small vane mounted near the stagnation point in the leading edge of the wing. At flights above the stall speed, the airflow over the vane is downward and the vane is held down. An electrical switch connected to the vane is open while the vane is down. As the angle-of-attack increases toward an impending stall, the stagnation point moves down until the airflow over the vane is upward. The vane is blown up, closing the switch and illuminating a red light or sounding a warning horn. [Figure 10-16] NON-ELECTRIC STALL WARNING The reed-type stall warning system operates in a manner similar to a musical instrument reed which produces a tone when air travels through it. The inlet of the small reed-type horn is located on the leading edge of the wing near the stagnation point.

Figure 10-16. W hen the wing is nearly stalled, the upward airflow moves the vane to activate the stall warning.

As the angle-of-attack increases, the low-pressure air traveling over the wing moves into an area "where the reed inlet is located, causing it to sound. By listening to the changing pitch of the horn, the pilot can easily identify the point at which the stall will occur. On many high-performance aircraft, the margin between the aerodynamically generated pre-stall buffet and the actual stall is insufficient. Using the

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stagnation point to activate a stall warning system may not provide enough warning. Many corporate jet and transport category aircraft use a stick shaker to provide the pilot with an earlier and more reliable warning of an impending stall. The stick shaker consists of a motor that drives an eccentric weight. This motor is attached to the control column and shakes it to alert the pilot before a stall develops. A stall-warning computer based on airspeed, angle-of-attack, flap configuration, and power setting activates the stick shaker. The system is energized at all times when the aircraft is airborne and is deactivated on the ground by squat switches on the gear.

ANGLE-OF-ATTACK INDICATORS All stall warning systems provide an indication of an impending stall that is related to the angle-of-attack. For precision flying, the pilot needs to know the actual angle-of-attack during various stages of the flight. One system for measuring and displaying the angle-of-attack uses a slotted probe sticking out of the side of the aircraft fuselage. The slots carry impact air into the housing of the probe where it moves a set of paddles connected to a variable resistor. The change in resistance moves a pointer around the indicator dial, which is calibrated in percent of the stall-speed angle-of-attack, or color-coded with a qualitative indication of angle-of-attack. [Figure 10-17] Another method of measuring angle-of-attack utilizes a vane-type sensor. A thin, wedge-shaped vane is mounted on a short arm that is free to rotate. In flight, the vane streamlines with the relative wind. As the angle-of-attack changes, the arm pivots and a potentiometer connected to the arm transmits a position signal to the stall warning system. The vane is heated to prevent ice formation. [Figure 10-18] The pilot can set a reference bug to show the desired ratio of the airspeed to the stall airspeed. For example, if the pilot wants to make an approach to landing at an airspeed of 30% over the stall speed, the reference bug would be set on 1.3. The pilot then maintains the attitude needed to center the angle-of-attack needle on the reference bug and the approach speed will automatically be correct. If the angle-of-attack goes above or below the desired value, the indicator will move away from the bug.

REMOTE POSITION INDICATING SYSTEMS A pilot needs to know that a control surface has actually moved when commanded. Remote position indicating systems provide feedback about the status of control surfaces, landing gear, control valves,

Figure 10-17. As the angle-of-attack changes, the amount of air entering the angle-of-attack sensor changes. This causes the paddles inside to change position. These paddles are attached to a potentiometer that varies the current to an indicator that in turn gives an indication of AOA.

and other mechanically actuated devices. DIRECT CURRENT Direct-current remote indicating systems are used in some aircraft to transmit position information so that it can be seen on an instrument dial. The position pickup, or transmitter, is a variable resistor

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Figure 10-20. A variable resistor provides a variable current to a coil that aligns a permanent magnet with the resistor's wiper. Figure 10-18. Many large airplanes utilize a vane-type sensor for angle-of-attack.

AUTOSYN SYSTEMS

One of the more popular remote indicating systems used for all types of mechanical movement is the Autosyn system. Autosyn is a registered trade name for a system that uses a single-phase electromagnet for the rotor and a three-phase delta connected coil for the stator. [Figure 10-21]

Figure 10-19. When the pilot selects the Test position on a Boeing 747 stall warning system, the air/ground relay is bypassed, the stick shaker operates, the black and white test indicator rotates, and the system checks the angle-of-attack vane and flap position sensor.

made of wire wound around an insulating core in the shape of a cylinder. Two wipers contact bare portions of the wire along one edge of the cylinder, and current flows into the circuit through one of the wipers and out through the other. The cylindrical resistor is tapped at each 120-degree position and is connected to a coil in the indicator that is wound on a ring-shaped core. The indicator coil is also tapped at each 120 degrees and connected to form an electrical delta circuit. The current through each of the three portions of the coil varies depending upon the position of the two wipers in the transmitter. As the current changes, so does the magnetic field. Since a small permanent magnet attached to the pointer always aligns with the composite magnetic field, the indicator is always aligned with the wiper arms in the transmitter. [Figure 10-20] ALTERNATING CURRENT Many larger aircraft require greater accuracy than is available from a DC remote position indicating system. For these applications, alternating-current systems of either the Autosyn or Magnesyn -type are used.

Figure 10-21. The Autosyn -type alternating-current remote indicating system employs two delta-wound coils. These coils align with each other; one of them attached to an input shaft and the other to a remote pointer.

The synchronous motors in the indicator and transmitter are identical. The rotors are connected in parallel and supplied with 28-volt, 400-hertz AC. The three-phase stators are also connected in parallel, and in most installations, one side of the rotor is connected to one of the terminals of the stator. Whatever position is being monitored physically moves the rotor of the transmitter. This could be the flap position, landing gear position, or oil or fuel quantity, as well as many of the pressure measurements made with bourdon tubes or pressure capsules. The AC magnetic field in the rotor induces a voltage in the three windings of the stator, and because the two stators are connected in parallel, the magnetic field in the indicator will be exactly the same as that

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in the transmitter. The same AC voltage as the rotor in the transmitter excites the rotor in the indicator so their magnetic fields are identical. Since mechanical load on the indicator rotor is nothing more than a small pointer, the rotor will assume the same position inside the indicator as the rotor inside the transmitter. The rotor in the indicator immediately follows any movement of the transmitter rotor. Many Autosyn systems use dual indicators. The two synchronous motors are stacked, and the shaft of the rear motor sticks through the hollow shaft in the forward motor. One dial serves both indicators, and the two pointers move in the same way the hands of a clock do. MAGNESYN SYSTEMS

Magnesyn is another remote indicating system bearing a registered trade name and operating on AC. The basic difference between an Autosyn and a Magnesyn system is in the rotor. The Magnesyn system uses a permanent magnet for its rotor rather than the electromagnet used in the Autosyn system. The stator of a Magnesyn system is a toroidal coil: a coil wound around a ring-shaped iron core. The transmitter and indicator are not necessarily the same physical size and configuration, but they are alike in their electrical characteristics. The coils in both the transmitter and the indicator are supplied with 28-volt, 400-Hertz AC, are tapped each 120 degrees, and are connected in parallel. The voltage generated in the transmitter coil is carried into the indicator coil where it produces magnetic fields in its three sections. The composite field of these coils pulls the permanent magnet in the indicator into exactly the same alignment as the magnet in the transmitter. Any movement of the transmitter magnet causes the magnet in the indicator to mirror the transmitter position. [Figure 10-22]

CONFIGURATION WARNING SYSTEMS The number and complexity of modern aircraft systems require various warning systems to alert the pilot of malfunctions or incorrect aircraft configuration for a particular flight mode. Most warnings are visual, aural, tactile, or some combination. Warnings alert the aircrew to conditions that require some sort of action to ensure proper and safe operation of the aircraft. The type of signal depends upon the degree of urgency. One type of warning system is the fire warning system, which will be covered in

Figure 10-22. A Magnesyn -type AC remote indicating system uses the paired relationship of two permanent magnets to transfer transmitter position information to an indicator.

depth in Chapter 16. Other types of warning systems include takeoff configuration warning, landing gear configuration warning, Mach/airspeed warning, stall warning, ground proximity warning system (GPWS), and the engine indication and crew alerting system (EICAS). TAKEOFF CONFIGURATION WARNING SYSTEM The takeoff configuration warning system is armed when the aircraft is on the ground and one or more thrust levers are advanced to the takeoff power position. A warning light and/or aural warning will sound if the stabilizer trim is not properly set, trailing edge flaps are not in the correct position, any leading edge devices are not properly set, or the speed brake is not properly stowed. The warning signal stops when all monitored devices are properly set. LANDING GEAR CONFIGURATION WARNING SYSTEM The landing gear indication lights are activated according to signals from each gear and the landing gear lever. The particular gear indications may vary slightly, but the FAA requires positive indication of "up and locked" and "down and locked" gear positions. A typical system might indicate the landing gear down and locked with an illuminated green light for each individual gear. Another may use a single green light for the entire gear configuration "down and locked" indication. If a single green light is used, the switches at each gear are connected in series so that the "down and locked" light only illuminates when all gear are in the proper position.

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When the landing gear is in disagreement with the landing gear lever position, a red light illuminates, meaning that the gear is in transit or in an unsafe condition. When the landing gear is in the proper up position and the gear lever is also in the "UP" position, the gear position lights go out signifying an "up and locked" condition. A technician normally checks the gear warning system during landing gear retraction tests. Problems with the warning system are often caused by the gear position switches. Always consult the manufacturer's service manual for the proper procedures for adjusting the landing gear position switches in addition to any other maintenance performed.

• Excessive descent rate.

On some aircraft, a steady warning horn is provided to alert the pilot that the airplane is in a landing configuration and the gear is not down and locked. The landing gear warning horn is usually dependent on flap and thrust lever position.

When one of these conditions is encountered, the computer flashes warning lights and sounds an alarm or warning. Some warnings are computer-generated directions such as "Pull up" or "Windshear."

Generally, when a thrust lever is retarded and any landing gear is not down and locked, the landing gear warning horn will sound, but can be silenced using the warning horn cutout switch. Under certain conditions, the landing gear warning horn cannot be silenced. Although the actual flap settings and thrust lever positions will vary from one aircraft type to another, generally some provision is made to remove the pilot's ability to silence the gear warning when specific conditions occur. For example, the warning horn cutout might be disabled if the radar altimeter indicates less than 1,000 feet above ground with the aircraft in a landing configuration and with an unsafe gear. MACH/AIRSPEED WARNING SYSTEM Some aircraft are equipped with Mach/airspeed warning systems that provide a distinct aural warning any time the maximum operating airspeed is exceeded. Reducing speed below the limiting value is usually the only way to silence the ■warning. The system operates from an internal mechanism inside the Mach/airspeed indicator. Test switches allow an operational check of the system at any time. Maximum operating airspeeds exist primarily due to airplane structural limitations at lower altitudes and airplane handling characteristics at higher altitudes. GROUND PROXIMITY WARNING SYSTEM (GPWS) The ground proximity warning system (GPWS) provides warnings and/or alerts to the flight crew when any of the following conditions exist:

• Excessive terrain closure rate. • Altitude loss after takeoff or go-around. • Unsafe terrain clearance when not in the landing configuration. • Excessive deviation below an ILS (Instrument Landing System) glide slope. • Descent below the selected minimum radio alti tude. • Windshear condition encountered.

ENGINE INDICATION AND CREW ALERTING SYSTEM (EICAS) Older commercial airplanes utilize electromechanical system indicators that employ multiple visual and aural cautions and warnings to alert of hazardous conditions such as engine problems or open cabin doors. Most of these systems use an annunciator that provides a master warning light along with an aural indication to alert the crew that a malfunction has occurred and that corrective action may be required. These indicators do not offer the versatility and redundancy available with modern digital technology. New generation aircraft use electronic displays and a full-time monitoring system known as EICAS, Engine Indication and Crew Alerting. The use of EICAS requires very little monitoring by the crew and promotes quick, accurate identification and recording of problems. EICAS reduces flight crew workload by automatically monitoring and recording engine parameters for later review. EICAS also alerts the aircrew of problems when necessary. It is operative through all phases of flight, from power-up through post-flight maintenance. Parameters used to set and monitor engine thrust are displayed full time. The system automatically displays any out-of-tolerance values on a cathode-ray-tube (CRT) or liquid-crystal display (LCD) in an appropriate color. The colored messages are designed to alert the aircrew to any failure and convey the urgency in which to respond. By utilizing electronic displays, EICAS provides accurate, timely information on a single screen rather than multiple engine instruments scattered throughout the panel.

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Position and Warning Systems

Figure 10-23. A simplified system diagram of the EICAS installed in the Boeing 757 shows its typically required components.

EICAS provides an improved level of maintenance data for the ground crew without causing the flight crew any extra workload. This has been achieved by designing a system that will automatically record subsystem parameters when malfunctions are detected. The system also provides the flight crew with the capability for manual data recording with the push of a single button. This eliminates the need for extensive hand recording of systems and performance data. These features increase the accuracy of maintenance data recordings and improve the communication between the aircrew and ground maintenance crews.

upper display unit shows primary engine parameters and crew alerting messages, and the lower display unit shows secondary engine parameters. [Figure 10-24] EICAS monitors inputs from airplane subsystems and sensors. When an abnormal condition is detected, EICAS will generate and display an alert, status, or maintenance message.

EICAS usually includes two multicolor display units, two computers, and two control panels. These components, together with two display-switching modules, cancel/recall switches, and captain's and first officer's master caution lights, jointly perform the various EICAS functions. [Figure 10-23] The EICAS computer processes and displays all engine and aircraft system information required by the crew. One computer is used at a time for displaying the data on both display units. Computer selection is done on the display select panel. The

Figure 10-24. EICAS operational mode displays and engine parameters are presented on two displays. A pilot can select status and maintenance readouts on the secondary display using the EICAS maintenance panel.