and correcting for insufficient vacuum, or removing and replacing the instruments. A list of the most common malfunctions, together with their correction, is included in figure 12-67.
ELECTRIC ATTITUDE INDICATOR In the past, suction-driven gyros have been favored over the electric type for light aircraft because of their comparative simplicity and lower cost. However, the increasing importance of the attitude indicator has stimulated development of improved electric-driven gyros suited for light plane installation. Improvements relating to basic gyro design factors, easier readability, erection characteristics, reduction of induced errors, and instrument limitations are reflected in several available types. Depending upon the particular design improvements, the details for the instrument display and cockpit controls will vary among different instruments. All of them present, to a varying degree, the essential pitch-and-bank information for attitude reference.
The typical attitude indicator, or gyro horizon as it is sometimes referred to, has a vertical-seeking gyro, the axis of rotation tending to point toward the center of the earth. The gyro is linked with a horizon bar and stabilizes a kidney-shaped sphere having pitch attitude markings. The sphere, horizon bar, and bank index pointer move with changes of aircraft attitude. Combined readings of these presentations give a continuous pictorial presentation of the aircraft attitude in pitch and roll with respect to the earth's surface.
magnet attached to the top of the gyro shaft spins at approximately 21,000 r.p.m. Around this magnet, but not attached, is a sleeve that is rotated by magnetic attraction at approximately 44 to 48 r.p.m.
As illustrated in figure 12-69, the 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 will fall to the low side. The resulting precession re-aligns the axis of rotation vertically.
The gyro 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.
Gyro stabilized horizon bar Power failure warning flag Vertical index Bank index Bezel ' mask The gyroscope motor is driven by 115 v., 400 Hz alternating current. The gyro, turning at approximately 21,000 r.p.m., is supported by the yoke and pivot assembly (gimbals). Attached to the yoke and pivot assembly is the horizon bar, which moves up and down through an arc of approximately 27°.
The kidney-shaped sphere provides a background for the horizon bar and has the words climb and dive and a bull's-eye painted on it. Climb and dive represent about 60° of pitch. Attached to the yoke and pivot assembly is the bank index pointer, which is free to rotate 360°. The dial face of the attitude indicator is marked with 0°, 10°, 20°, 30°, and 60° of bank, and is used with the bank index pointer to indicate the degree of bank left or right.
The face of one type of gyro-horizon is shown in figure 12-68.
The function of the erection mechanism is to keep the gyro axis vertical to the surface of the earth. A Trim knob Caging knob FIcuRE 12-68. Gyro-horizon indicator.
Magnetic Compass The magnetic compass is a simple, self-contained instrument which operates on the principle of magnetic attraction.
If a bar magnet is mounted on a pivot to be free to rotate in a horizontal plane, it will assume a position with one of its ends pointing toward the earth's north magnetic pole. This end of the magnet is called the north-seeking end of the magnet.
Erection Magnet mechanism 44-48 г.р.m.
Reaction to precession forces
• 21,000 r.p.m.
Gyro universally mounted Caging mechanism FreuRE 12-69. Erecting and caging mechanisms of an electric attitude indicator.
The magnetic compass consists of = liquid-filled bowl containing a pivoted float element to which one or more bar magnets, called needles, are fastened. The liquid in the bowl dampens the oscillations of the float and decreases the friction of the pivot. A diaphragm and vent provide for expansion and contraction of the liquid as altitude and/or temperatures change.
If more than one magnet is used in a compass, the magnets are mounted parallel to each other, with like poles pointing in the same direction. The element on which the magnets are mounted is so suspended that the magnets are free to align themselves with the earth's north and south magnetic poles.
A compass card, usually graduated in 5° increments, is attached to the float element of the compass. A fixed reference marker, called a lubber line, is attached to the compass bowl. The lubber line and the graduations on the card are visible through a glass window. The magnetic heading of the aircraft is read by noting the graduation on which the lubber line falls. The two views of a magnetic compass in figure 12-70 show the face and the internal components of a magnetic compass, A compensating device containing small permanent magnets is incorporated in the compass to correct for deviations of the compass which result from the magnetic influences of the aircraft structure and electrical system. Two screws on the face of the instrument are used to move the magnets and thus counterbalance the local magnetic influences acting on the main compass magnets. The two setscrews are labelled IN-S and F-W.
Magnetic variation is the angular difference in degrees between the geographic north pole and the magnetic north pole. This variation is caused by the earth's magnetic field, which is constantly changing.
Since variation differs according to geographic location, its effect on the compass cannot be removed by any type of compensation. Variation is called west variation when the earth's magnetic field draws the compass needle to the left of the geographic north pole and east variation when the needle is drawn to the right of the geographic north pole.
The compass needle is affected not only by the earth's magnetic field, but also by the magnetic fields generated when aircraft electrical equipment is operated, and by metal components in the aircraft. These magnetic disturbances within the aircraft, called deviation, deflect the compass needle from alignment with magnetic north.
To reduce this deviation, each compass in an aircraft is checked and compensated periodically by adjustment of the N-S and E-W magnets. The errors remaining after "swinging" the compass are recorded on a compass correction card mounted near the compass.
Filler plug Magnetic compensator assembly Spring suspension Liquid chamber Lubber line E 6
• Card Siphon expansion chamber Float Pivot assembly FICURE 12-70. Magnetic Compass.
The "swinging" (calibration) of a compass can be accomplished in flight or on the ground. Ground swinging of a compass is usually done with the aircraft at rest on a compass "rose." A compass rose (figure 12-71) is a circle laid out or painted on a level surface and graduated in degrees. The directions marked on the compass rose are magnetic directions, although true north is also marked on some compass roses.
Compass compensation procedures vary, depending on the type of aircraft. Requirements are often set up on a flight-hour and calendar basis. Most facilities perform compass checks anytime that 360° 330° 30° * 300° 80°1
• 270° 90° 240° 120° 210° 150° 180° FICURE 12-71. Typical compass rose.
equipment replacement, modification or relocation might cause compass deviation.
An example of compass compensation is outlined in the following paragraphs. These procedures are general in nature and do not have specific application.
(1) The compensator should be set either to zero or in a position where it has no effect on the main compass magnets.
(2) The aircraft is placed directly on a south magnetic heading on the compass rose.
The tail of tailwheel aircraft should be raised to level-flying position.
(3) Note the compass reading and record it.
The deviation is the algebraic difference between the magnetic heading and the compass reading.
EXAMPLE:
On the south (180°) heading, the compass reading is 175.5°. This would be recorded as a deviation of +4.5° (180° - 175.5° = 4.5°). If the compass reading is too low, the deviation is plus; if the reading is too high, the deviation is minus.
(4) Align the aircraft on a magnetic north heading. Record the compass reading and compute the deviation.
EXAMPLE:
On the north (000°) heading, the compass reads 006.5°. Since the deviation is
6.5° too high, it is recorded as a minus deviation (-6.5°).
(5) The coefficient of north-south deviation is determined by subtracting, algebraically, the south deviation from the north deviation and dividing the remainder by 2:
Coefficient = (-6.5°) - (4.5°) N
- 11°
(10) Leaving the aircraft on an east magnetic heading, compute the coefficient of overall deviation. This coefficient is equal to the algebraic sum of the compass deviations on all four cardinal headings (north, east, south, and west), divided by 4:
(-65°) +0° +45° +(-6°) Coefficient = = - 5.5° The coefficient of north-south deviation, which is the average of the deviation on the two headings, is -5.5°. The northsouth compensator is adjusted by this amount, and the reading heading will now be 001°. This adjustment also corrects the south deviation by the same amount, so that on a south heading the compass will now read 181°.
--8° 4 = -2°.
If the coefficient is greater than 1°, further compensation is usually accomplished. The compensation is not done with the magnetic compensation device. It is accomplished by re-aligning the compass, so that it is mounted parallel to the longitudinal axis of the aircraft.
(6) Align the aircraft on a magnetic west (270°) heading on the compass rose.
Record the compass reading and compute the deviation. Suppose the compass reads 276°, a deviation of -6°.
(7) Align the aircraft on a magnetic east (090°) heading. Record the compass reading and compute the deviation. Suppose the compass reading is exactly 90° on the magnetic east heading, a deviation of 0°.
(8) Compute the coefficient of east-west deviation:
Coefficient = 0°- (-6°) + 6° = = + 3°
(9) While the aircraft is on the east heading, adjust the east-west compensator to add 3° to the compass reading. This reading then becomes 93° on the east heading and 273° on a west heading.
(11) After the initial compensation is completed, the aircraft will be compensated again on headings of 30°, 60°, 120°, 150°, 210°, 240°, 300°, and 330°. The compass readings for each heading are recorded on a compass correction card.
This card is then mounted as close as possible, to the instrument for ready reference. An example of a correction card is shown in figure 12-72.
The procedure described is a basic compensation procedure. Additional circuits around the compass rose should be made with the engine(s) and electrical and ratio equipment operating to verify the accuracy of the basic compensations.
Jacks, lifts, hoists, or any dolly needed to move and align the aircraft on the various headings of a compass rose should preferably be made of nonmagnetic material. When this is impossible, devices can be tested for their effect on the compass by moving them about the aircraft in a circle at the same distance that would separate them from the compass when they are being used. Equipment that causes a change in compass readings of more than onequarter of a degree should not be used. Additionally, fuel trucks, tow tractors, or other aircraft containing magnetic metals should not be parked close enough to the compass rose to affect the compass of the aircraft being swung.
AIRCRAFT COMPASS (5) The scale should be readable and its illumination good.
DATE
FOR STEER AUTOPILOT SYSTEM The automatic pilot is a system of automatic con- N 000° 000° trols which holds the aircraft on any selected magnetic heading and returns the aircraft to that head- 030° 033° ing when it is displaced from it. The automatic pilot 060° 060° also keeps the aircraft stabilized around its horizontal and lateral axes.
090° 095° 120° 120° The purpose of an automatic pilot system is primarily to reduce the work, strain, and fatigue of 150° 149° controlling the aircraft during long flights. To do S 180° 175° this the automatic pilot system performs several functions. It allows the pilot to maneuver the air- 210° 205° craft with a minimum of manual operations. While under automatic control the aircraft can be made to 240° 334° climb, turn, and dive with small movements of the W 270° 265° knobs on the autopilot controller.
300° 294° 330° 326° Calibrated by:
FIcURE 12-72. Compass correction card.
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:
Autopilot systems provide for one, two, or three axis control of the aircraft. Some autopilot systems control only the ailerons (one axis), others control ailerons and elevators or rudder (two axis). The three-axis system controls ailerons, elevators, and rudder.
All autopilot systems contain the same basic components: (1) Gyros, to sense what the airplane is doing; (2) servos, to move the control surfaces;
and (3) an amplifier, to increase the strength of the gyro signals enough to operate the servos. A controller is also provided to allow manual control of the aircraft through the autopilot system.
(1) The compass indicator should be checked for correct readings on various cardinal headings and re-compensated if necessary.
(2) Moving parts of the compass should work easily.
(3) The compass bowl should be correctly suspended on an anti-vibration 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 nor have any discoloration.
Principle of Operation The automatic pilot system flies the aircraft by using electrical signals developed in gyro-sensing units. These units are connected to flight instruments which 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 used to control the operation of servo units which convert electrical energy into mechanical motion.
The servo is connected to the control surface and converts the electrical signals into mechanical force which moves the control surface in response to corrective signals or pilot commands. A basic autopilot system is shown in figure 12-73.
Aircraft on course Gyro input signal Amplitier Control surface feed back signal Feedback Servo circuit FIGURE 12-73. Basic autopilot system.
Most modern autopilots can be described in terms of their three major channels: (1) The rudder, (2) aileron, and (3) the elevator channels.
The rudder channel receives two signals that determine when and how much the rudder will move.
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 will develop. 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, which 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 will turn the rudder in the proper direction to return the aircraft to the selected magnetic heading.
As the rudder surface moves, a followup signal is developed which 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 will reach a zero value, and the rud der will be returned to the streamline position by the followup 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 will cause 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 followup 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 streamline, the aircraft will now start moving back toward level flight with the input signal becoming smaller and the followup signal driving the control surfaces back toward the streamline position. When the aircraft has returned to level flight in roll attitude, the input signal will again be zero. At the same time the control surfaces will be streamlined, and the followup signal will be zero.
The elevator channel circuits are similar to those of the aileron channe., with the exception that the 512
Gyro Servoamplifier control unit Controller Elevator servomotor Amplifier Discriminator Pitch Pitcl Clutch transmitter channel Blevator Magnetic amplifier Servomotor Gear 1-1 -Gea Aileron amplifier channel Rate gen Gear Roll Roll Followup transmitter Rudder amp irier channe!
transmitter Rate Aileron servomotor transmitter Servo disconnect Rudder servomotor >>-- FICURE 12-74. Autopilot block diagram.
elevator channel detects changes in pitch attitude of the aircraft. The circuitry of all three channels can be followed by referring to the block diagram in figure 12-74.
The foregoing autopilot system description was used to show the function of a simple autopilot.
Most autopilots are far more sophisticated; however, many of the operating fundamentals are similar. Autopilot systems are capable of handling a variety of navigational inputs for automatic flight control.
BASIC AUTOPILOT COMPONENTS The components of a typical autopilot system are illustrated in figure 12-75. Most systems consist of four basic types of units, plus various switches and auxiliary units. The four types of basic units are:
(1) The sensing elements, (2) command elements,
(3) output elements, and (4) the computing ele- Command Elements The command unit (flight controller) is manually operated to generate signals which cause the aircraft to climb, dive, or perform coordinated turns.
Additional command signals can be sent to the autopilot system by the aircraft's navigational equipment. The automatic pilot is engaged or disengaged electrically or mechanically, depending on system design.
While the automatic pilot system is engaged, the manual operation of the various knobs on the controller (figure 12-76) maneuvers the aircraft. By operating the pitch trim wheel, the aircraft can be made to climb or dive. By operating the turn knob, the aircraft can be banked in either direction. The engage switch is used to engage and disengage the autopilot. In addition, most systems 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.
One type of automatic pilot system has an engaging control that manually engages the clutch mechanism of the servomotor to the cable drum. A means of electrically disengaging the clutch is provided through a disconnect switch located on the control wheel (s).
Sensing Elements The directional gyro, turn-and-bank gyro, attitude gyro, and altitude control are the sensing elements. These units sense the movements of the aircraft, and automatically generate signals to keep these movements under control.
Computer or Amplifier The computing element consists of an amplifier or computer. The amplifier receives signals, determines what action the signals are calling for, and ampli- 00 De fies the signals received from the sensing elements.
It passes these signals to the rudder, aileron, or elevator servos to drive the control surfaces to the position called for.
Output Elements The output elements of an autopilot system are the servos which actuate the control surfaces. The 513
Sensing elements Command elements Output elements Pressure source Directional gyro indicator он OтF Flight Aileron controller servo actuator Turn-and-bank indicator gyro Rudder servo actuator Computer Attitude indicator Elevator Navigation signals servo actuator Altitude control Heading selector Trim actuator FIGURE 12-75. Typical autopilot system components.
majority of the servos in use today are either electric motors or electro/pneumatic servos.
An sircraft may have from one to three servos to operate the primary flight controls. One servo operates the ailerons, a second operates the rudder, and a third operates the clovators. Each servo drives ita 514
OFI FIGURE 12-76. Typical autopilot controller.
associated control surface to follow the directions of the particular automatic pilot channel to which the servo is connected.
Two types of electric motor-operated servos are in general use. In one, a motor is connected to the servo output shaft through reduction gears. The motor starts, stops, and reverses direction in reponse to the commands of the gyros or controller.
The other type of 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.
The electro/pneumatic servos 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 electro/ magnetic valve assembly and an output linkage assembly.
FLIGHT DIRECTOR SYSTEMS A Hight director system is an instrument system consisting of electronic components that will compute and indicate the aircraft attitude required to attain and maintain a preselected flight condition.
"Command" indicators on the instrument indicate 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 operator of many of the mental calculations required for instrument fights, such as interception angles, wind drift correction, and rates of climb and descent.
A flight director system has several components;
the principal ones are the gyroscope, computer, and the cockpit presentaton.The gyro detects deviations from a preselected aircraft attitude. Any force exerted against the gyroscope is electrically transmitted to the computer, which in turn, sends a computed signal to the flight indicator, telling the operator what must be done with the controls. When using a flight director system, the operator is, in a sense, acting as a servo, following orders given by the command indicators.
The computers used in the various types of flight director systems are basically the same; however, the numbers and types of functions available will vary between systems because of the mission of a particular aircraft, the limited aircraft space available for installation, and the excessive cost of functions not absolutely required.
The instrument panel presentations and operating methods vary considerably between different sys.
tems. Command indications may be presented by several different symbols, such as bar-type command indicators with different types of movements, a phantom aircraft symbol, or two-element crossbar indicators.
Many flight director systems are equipped with an "altitude-hold" function, which permits selection of a desired altitude; the flight director computes the pitch attitude necessary to maintain this particular altitude.
A flight director greatly simplifies problems of aerial navigation. Selection of the VOR function electronically links the computer to the omnirange receiver. After selection of the desired omnicourse, the flight director will direct the bank attitude necessary to intercept and maintain this course.
Flight director systems are designed to offer the greatest assistance during the instrument approach phase of flight. ILS localizer and glide slope signals are transmitted through the receivers to the computer, and are presented as command indications.
With the altitude-hold function, level flight can be maintained during the maneuvering and procedure turn phase of the approach. Once inbound on the localizer, the command signals of the flight director are maintained in a centered or zero condition.
Compensation for wind drift is automatic. Interception of the glide slope will cause a downward indication of the command pitch indicator. Any de- 515
viation from the proper glide slope path will cause a fly-up or fly-down indication on the flight director pitch command symbol. When altitude-hold is being used, it automatically disengages when the glide slope has been intercepted.
A flight director system not only shows the present situation, but also predicts the future consequences of this situation. For example, a momentary change in attitude is detected by the computer, and command symbol movement is made to correct this condition possibly before an altitude error can result. Thus, greater precision is achieved with less mental effort on the part of the aircraft operator.
AUTOPILOT SYSTEM MAINTENANCE The information in this section does not apply to any particular autopilot system, but gives general information which 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.
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 mechanism 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 compo- a to the nent, or whenever a malfunction in the autopilot system 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 system using the engaging control (switch, handle).
After the system is engaged, perform the operational checks specified for the particular aircraft. In general, the checks consist of:
(1) Rotate the turn knob to the left; the left rudder pedal should move forward, and the control column wheel should move to and the control column wheel should move 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. The control column should move slightly aft. Return the turn knob to the center position; the flight controls should return to the level-fight position.
(3) Rotate the pitch-trim knob forward; the control column should move forward.
(4) Rotate the pitch-trim knob aft; the control If the aircraft has a pitch-trim system installed, it should function to add downtrim as the control column moves forward, and add uptrim as the column moves aft. Many pitch-trim systems have an automatic and a manual mode of operation. The above action will occur 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 re-engage the system and check the emergency disconnect release circuit. The autopilot should disengae each time the release button is actuated.
When performing maintenance and operational checks on a specific autopilot system, always follow the procedure recommended by the aircraft or equipment manufacturer.
Annunciator System Instruments are installed for two purposes, one to display current conditions, the other to notify of unsatisfactory conditions. Colored scales are used; usually green for satisfactory; yellow for caution or borderline conditions; red, for unsatisfactory conditions. As aircraft have become more complex with many systems to be monitored, the need for a centralized warning system became apparent. The necessity to coordinate engine and flight controls emphasized this need. What evolved is an annunciator or master warning system (figure 12-77).
Certain system failures are immediately indicated on an annunciator panel on the main instru ment panel. A master caution light and a light indicating the faulting system flash on. The master light may be reset to "Off, but the indicating light will remain "On" until the fault is corrected or the equipment concerned is shut down. By resetting.
the master caution light is ready to warn of a sub- 516
SYSTEM ATA NUMBER INDICATION uses bimetallic strips set at predetermined temperatures. The second group responds to pressure Aircraft Fuel Fuel Pressure Low changes and uses a flexible chamber that moves Engine Fuel Fuel Pressure Low when pressurized. The third group consists of Electrical Inverter Out mechanically operated switches and/or contacts on Generator Generator Generator Overheated Generator Out a relay.
Starting Starter Engaged An annunciator system may include any or all Oil Pressure Low of the following indications or others as applicable.
Brake Pressure Low Not Locked Down Aural Warning System Anti-Skid Out Cabin Pressure High Aircraft with retractable landing. gear use an una Cabin Pressure Low aural warning system to alert the crew to an unsafe Speed Brake Extended condition. A bell will sound if the throttle is re- Not Set for Takeoff Thrust Reversal locked condition (figure 12-78). tracted and the landing gear is not in a down and Aural warning systems range in complexity Doors ked from the simple one just described to that system necessary for safe operation of the most complex Doors ked ater transport aircraft.
Out A typical transport aircraft has an aural warnpen ing system which will alert the pilot with audio signals to: An abnormal takeoff condition, landing Low condition, pressurization condition, mach-speed con- Firewarning dition, an engine or wheel well fire, calls from the crew call system, and calls from the secal system.
Shown in figure 12-78 are some of the problems FIGURE 12-77. Warning in annunciator system. which trigger warning signals in the aural warning system. For example: a continuous horn sounding sequent fault even before correction of the initial during landing would indicate the landing gear is fault. A press to test light is available for testing not down and locked when flaps are less than full the circuits in this system. up and the throttle is retarded. The corrective vices divided into groups, according to their method One late model business jet has the sensing de- throttle. action would be to raise the flaps and advance the of operation. The fast group responds to heat and (See Agure 12-78 on next page) 517
STAGE OF WARNING WARNING CAUSE OF WARNING SIGNAL CORRECTIVE
OPERATION SYSTEM SIGNAL ACTIVATION ACTION Landing Landing gear Continuous Landing gear is not down and locked Raise flaps ATA 3200 born when flaps are less than full up Advance throttle and throttle is retarded to idle.
In flight Mach warning Clacker Equivalent airspeed or mach Decrease speed ATA 3400 number exceeds limits. of aircraft Takoff Flight contro Intermittent Throttles are advanced and any Correct the aircraft ATA 2700 hom of following conditions exist. to proper takeoff ATA 4900 Aux power 1. Speed brakes are not down 2. Flaps are not in takeoff range conditions.
3. Auxiliary power exhaust door is open
4. Stabilizer is not in the takeoff setting.
Inflight Pressurization Intermittent horn If cabin pressure becomes equal Correct the ATA 2100 to atmospheric pressure at the specitic condition.
altitude (altitude at time of occurrence).
Any stage Fire warning Continuous bel Any overheat condition or 1. Lower the heat ATA 2600 fire in any engine or in the area wherenacelle, or main wheel or nose wheel in the F/W was well, APU engine or any compartment activated.
having firewarning system installed. Also whenever the firewarning system 2. Signal may be si- lenced by pushing is tested. the F/W bell cutout switch or the APU cutout switch Any stage Communications High chime Any time captain's call button is Release button or if ATA 2300 pressed at external power pane. button remains locked forward or rearward cabin attend in, pull button out.
ant's panel Any stage Communications Tow tone hi-low Whenever a signal has been Press reset button on secal system® chime or received by an HF or VHF com- secal system control ATA 2300 single low chime. munication system and decoded by panel.
the secal® decoder.
*NOTE: Secal system is the Selective Calling System: Each aircraft is assigned a particular four tone audio combination for identification purposes. A ground station will key the signal whenever contact with that particular aircraft is desired. The signal will be decoded by the airborne secal decoder and the crew allerted by the secal warning system.
FIGURE 12-78. Aural warning system.