ELECTRICAL SHOPo ANTI SKID BRAKING SYSTEM o LANDING GEAR ASSEMBLY o STALL WARNING o ACTUATOR o THERMOCOUPLE o IGNITION EXCITER o PROXIMITY CARDS o GENERATOR o EXTRACTION FANo

CURRENT TRANSFORMER ASSEMBLY

o EMERGENCY LIGHT POWER SYSTEM

o o o o o

INSTRUMENT SHOPAUTO PILOT COCKPIT VOICE RECORDER RATE OF CLIMB INDICATOR DISPLAY UNIT MEASUREMENT OF PRESSURE

ELECTRICAL SHOP

ANTI-SKID AUTOBRAKE CONTROL UNITIt consists of 4 wheel speed transducers and 4 control valves. The principle function of the control is to process wheel speed information, sense impending wheel skids and produce connection currents. The control unit receives wheel speed information from the 4 transducers and drives the 4 valves with control currents which commands brake pressure radiations and prevent skids.

DESCRIPTIONThe AACU is housed in a container and consists of 4 identical main wheel/BITE cards, 1 power supply card, 1 switching logic card and 1 auto brake card. Indicators are mounted on the rear of the CU connecting to the airplane.

OPERATIONThe AACU senses the speed of the 4 types and 4 transducers connected to the 4 wheels convert the speed into electrical signals. These electrical go to the modules, which compare these signals with some reference signal. If not equal it implies skidding and the valve releasing the brake fluid is closed thus avoiding the braking of that particular wheel. Hence all the 4 tires move synchronously at the same speed.

LANDING GEAR ACCESSORY UNIT ASSEMBLYIt consists of control and safety relays, solid-state circuits and related wiring and connectors mounted in a chassis assembly. The accessory unit assembly includes air and ground sensing indicators and test switches.

OPERATIONThe LGAUA receives signals from proximity sensors on the landing gear. These signals are transmitted to the solid state switching circuits in the accessory unit assembly to control the relays. The relays provide the required control and indication of the landing gear. The air & ground sensing indicators and test switches are used to check for malfunction in the accessory unit assembly and to isolate the safety relays for airplane maintenance purposes. The assembly controls and monitors the following systems: 1. 2. 3. 4. Safety Relays Systems (squat switches). Landing Gear Rearing Systems. Automatic Ground Speed Brake Systems. Take Off Warning Systems.

ACTUATORThe Actuator is an electromechanical unit with 2 ends of travel positions: retracted and extended. The actuator assembly is attached at one end by the adjustable eyeball of a ball screw and at the other by a fixed bearing.

DESCRIPTIONThe actuator assembly consists of: 1. A single phase electric motor equipped with an electromagnetic brake. The motor is sealed to external elements on the casing side and sealed against grease on the motor gear side. 2. A ball screw assembly converting rotational movement of the nut into a translation by the screw. 3. A circuit allowing the balls to circulate reducing friction between the nut and screw. 4. A clutch assembly consisting of 2 friction disks placed between the 2 transmission gears. In the mechanical chain the clutch is located between electrical motor and ball screw. 5. A recopy assembly consisting of a gear potentiometer, 2 cams and 2 microswitches. train, a

6. A transmission gear assembly providing the mechanical link between the electric motor ball screw nut.

OPERATION1. When the electric motor is not supplied, the mechanical system of the electromagnetic is engaged. When the motor is supplied the coil of the electromagnetic brake releases the mechanical system.

2. The rotational movement of the electrical motor axle is transmitted to the ball screw via the gear train. 3. The ball screw assembly transforms the rotational movement of the nut into a translation by the screws. A ball circuit provides mechanical linkage between the nut and screw and reduces friction. Two mechanical stops limit travel of the screw in fully extended and fully retracted positions. 4. The gear on the ball screw controls a recopy system. This system consists of a set of gear providing rotation of the spindle of the potentiometer. The rotation of this spindle gives differing electrical values according to the position of the translating ball screw. 5. Two cams of the potentiometer spindle activates the microswitches, one for the retracted position and one for the extended position. 6. These microswitches switch off the electromagnetic brake coil. The mechanical brake system engages and stops the rotational movement of the electric motor axle.

STALL WARNINGGROUND CRITICAL SYSTEMS1.Drain Mast Heater

AIR MODE

GROUND MODE

Switches the heater from 28V - 115V power source to provide higher heating of the drain mast. Arms the stall warning system.

Switches the heater from 115V - 28V power source to reduce the heating of the drain mast. Deactivates the stall warning system. Deactivates the antiskid touch down protection circuit & allows normal braking application. Arms the APU wheel well fire warning circuit. De-energizes the lever latch solenoid to prevent the landing gear handle from being operated to the up position.

2. Stall Warning

3. Antiskid System Prevents inboard brake application by actuating the antiskid control valves to the full 4. APU Fire dump positions. Detection Hour Deactivates the APU wheel well fire warning circuit. 5. Landing Gear Latch Energizes the lever latch solenoid to enable landing gear retraction without override.

THERMOCOUPLEGENERALA Thermocouple is an instrument used for measuring temperature. Its sensing element delivers a signal, which is a function of the temperature of the medium in which it is placed. When installed on a brake thermocouple provides means of checking the heat generated during brake applications.

CHARACTERISTICSWeight : Probe : Max. Operating temperature : 120 kg Chromel, Alumel 800 deg C

DESCRIPTIONThe major components of the thermocouple are: 1. A housing assembly 2. A connector (BRAKE) A probe attached to the housing is fitted with two leads on which two pins are crimped. These pins are housed in the connector as well as a third pin, which has no assigned function. The connector is secured to the housing by 4 wire locked screws. An O ring placed between the connector and housing provides sealing efficiency. (EXHAUST PIPE) It is a NiCr / NiAl probe assembly designed to give two independent E.M.F. outputs. The thermocouple probe is mounted within the initiator stator guide vane and secured by means of a fixing flange. Four probes are fitted to each engine. Electrical connections to the thermocouple probe are made by means of 4 terminal studs; 2 are of NiCr and 2 are of NiAl. The terminals are mounted to the head can, which is supported by the fixing flange.

The interconnections within the head can are potted in ceramic case. The thermocouple elements are unsheathed and are supported by fine location plates strategically placed along the element lengths.

OPERATION(BRAKE) Under the action of the heat produced braking the probe temperature rises through conduction. Then an electrical current flows through the alumel-chromel circuit. The voltage of this current, which is proportional to temperature, is processed by an indicator. (EXHAUST PIPE) When the functions of the elements are heated an E.M.F. is produced proportional to the junction temperature.

TEST PROCEDURE(A) RESISTENCE TEST: - The Thermocouple shall be allowed

to stabilize at room temperature. The DC resistance of the thermocouple between terminals shall be measured using the wheatstone bridge. 68 deg F or 20 deg C 1.03 to 1.53 ohms(B) INSULATION TEST: - Conduct the test at normal ambient

temperature. Connect the insulation tester between the thermocouple body and its terminals. The insulation resistance shall exceed 100 K ohms using a test potential of 100V DC.

STORAGEHumidity not exceeding 50% Temperature -20deg C to +50deg C

The Thermocouple should be inspected for: 1. 2. 3. 4. 5. Cleanness Security of terminals Serviceability of terminal threads Good conduction of head can, visually check for cracks General condition of all metal parts

IGNITION EXCITERDESCRIPTIONIt is a capacitor discharge system that provides both intermittent duty ignition during starting and a continuous duty ignition after starting. The intermittent duty circuit requires an input of 24V DC and discharges through both outlets firing two igniter plugs. The continuous duty circuit requires an input of 115V AC at 400 Hz and discharges through only one outlet marked firing one igniter plug. Both circuits of the ignition exciter are contained in one compact hermetically sealed housing for long life under severe environmental operating conditions. The exciter has 1 input power connector and 2 output connectors.

OPERATIONINTERMITTENT DUTY CIRCUIT :When the ignition switch is closed, input current flows through filter coils L4 & L5, resistor R5, base to emitter function of power transistor Q1 and resistor R7 through ground back to the power supply. Base to emitter current causes transistor Q1 to be forward biased that allows current flow through the primary of transformer T1, collector to emitter of Q1 and through resistor R7 to ground. Rising primary current induces a voltage across the tertiary or feedback winding of T1 providing a constant current flow through R6 and base to emitter of Q1. When the current level through the primary winding of T1 reaches the saturation value of Q1 the tertiary winding voltage will drop causing a decrease of current in the primary winding. The tertiary polarity will reverse rapidly cutting off current in the primary winding. Voltage across the primary now goes negative causing transistor Q1 to be biased to cutoff. Zener diode CR4

protects Q1 from collector emitter breakdown caused by high amplitude pulse, which may occur during transistor turn off. CONTINUOUS DUTY CIRCUIT : Current from the 400 Hz power source flows through the filter coil L1 and primary of power transformer T2. This induces a high voltage in the secondary winding which charges capacitor C3 & C4. Capacitor C5, in parallel with C3 & C4, charges to the voltage on C3 & C4 added together. When the voltage on C5 reaches the breakdown voltage of G1, it breaks down. A small amount of current flows from C5 through L9, L11 and C11 to ground, through gap, G1 back to capacitor C5. This current induces a high frequency, high voltage in a coil L10 and ionizes the igniter plug gap. The remainder of the charge on capacitor C5 discharges through coil L10, across the igniter plug gap and across gap G1 back to capacitor C5.

PROXIMITY SENSOR/CARDSGENERALIt is used for identifying position of a thing and is different from magnetic switches. Basically an indicator has 2 arms: - one purely inductive and purely resistive, forms arms of wheatstone bridge. It is used for sensing whether flap/slat/door are in open/closed position. When a metallic body comes in the vicinity of the sensor, the inductance is changed and hence the bridge gets unbalanced. It will operate as switch (using transistor) and lights will be operated when the door is off light gets off (generally).

PROXIMITY SWITCH DC ASSEMBLY DESCRIPTIONThe ELDEC proximity switch assembly is a solid state switch which works together with an ELDEC proximity sensors to detect the presence / absence of metal structure. ELDEC switches operate normally open / normally closed (selected by jumpers at the switch card connector). Switch output can be connected in parallel for a logic OR function. Up to 6 switches may be connected in series for a logic AND function. The sensor appears as an inductance to the switch card. The switch card senses a change of inductance caused by the proximity of a metal target. The switch then completes (or in the case of a normally closed switch, interrupts) the circuit to an external load. The proximity switch is designed so that open or shorted sensor leads inhibit switch actuation regardless of target position. Loss of power ceases the switch output to fail resulting in an open circuit to the load.

OPERATIONIts input is 16 V peak to peak at 1100 Hz.1. VOLTAGE REGULATOR: - It provides stable voltage free of

large transient variation switch circuit. In this block codes are used to provide regulator input protection. One diode protects against reverse zener diode that provides high voltage transient protection. The two other diodes provide base voltages for emitter follower. The voltage regulator provides 18 V output.2. POWER OSCILLATOR: - Sensor oscillation is provided by

the power oscillator. It generates 1100 Hz square swinging between ground and regulated voltage.3. BRIDGE CIRCUIT: - It has two arms, one inductive and

one resistive and are located in the sensor. An unbalanced bridge is reflected by the change in voltage between the red and blue load.4. SYNCHROUS DEMODULATION: - It increase the common

mode reflection of unwanted signals. This voltage is applied to the comparator circuit. Demodulator has a FET, which acts as a switch.7. COMPARATOR: - It is a high gain feedback amplifier and

compares the sensor signal with the threshold voltage.8. SWITCH DRIVER: - If the output of the comparator is not

sufficient it amplifies the signal.9. OUTPUT SWITCH: - It can be operated in normally open /

normally closed condition.

SPECIFICATIONWeight Power Temperature Switching Range -- 2.5 ounce -- 28 V DC, 1.8 W -- 40 to 85 deg C -0.5 A

GENERATORDESCRIPTIONThe generator is a self exciter air cooled, brushless machine designed for use in aerospace application to provide an alternating current power for aircrafts electrical system. It consists of an exciter rotating armature, exciter stationary field, a three phase full wave rectifier bridge and capacitor. It is driven and held with in speed range by a constant speed transmission, which is in turn driven by the aircraft engine.

SPECIFICATIONRating ...40 KVA Voltage .120/208 v(ac) Frequency.380/420 hz Current..111 amp Phase..3 Power factor.0.75 Speed.5700/6300 rpm Rotation (facing end opposite drive)ccw/cw Phase rotation..T1,T2T3/T2,T1/T3 Outline(instalation drawing)....915f317 Size: Max length from mounting flange...12.60 inch Max height11.820 inch Max weight...82.5 pound

OPERATIONThe stationary exciter winding has permanent magnet between the coils. This permanent magnet provide a minimum residual voltage from the generator to the regulator to assure build up under all conditions turning the rotor produces a voltage build up as a result of the residual magnetism or the flux from the permanent magnets. This produced voltage is fed to the voltage regulation. The voltage regulator rectifies the voltage signal and returns a controlled signal to the generator through the pins A and F. this voltage excites the exciter field and produces a three phase alternating current in the rotating exciter armature. This is then fed to the three phase full wave rotating

rectifier where it is rectified to direct current. This current excites the main rotating field, cutting the conductor in the stationary armature produces three phase alternating current output to the main terminals. The exciter, which provides power for the rotating field uses a stationary direct current field and a rotating alternating current armature. The stationary field has two winding in parallel. A thermistor (negative temp coefficient) is placed in series, with one of the winding in order to provide nearly constant exciter field resistance over the temp range encountered by the generator. The alternating current armature is designed to have a high reactance so that its excitation requirements are dependent on the exciter load current rather on the exciter load voltage. The rotating rectifier consists of six silicon power diodes connected in a three phase full wave bridge. The diodes have a reverse voltage rotating equal to 150 percent of the measured transient over voltage under the most severe operating conditions. A capacitor is connected across the direct current output of the full wave bridge to provide transient voltage suppression, this is used to reduce the commutation voltage and to control the other overvoltages caused by system transients which would subject the diodes to excessive peak reverse voltage. The main rotating field consists of eight sets of field coils , connected in service with alternate polarities. The coils are mounted on eight poles which are each on integral position of the main laminated or stacked core rotor, each set of coils consists of two coils, a larger and a smaller which are assembled concentrically on a single pole and connected in a like polarity. Direct current excitation current from the rotating rectifier produces a flux in the eight poles which in turn induces output voltage in the main stator winding. Amorttisseur(damper) winding are provided in longitudinal slots in the pole face. They also function as a wedge to retain the field windings in place. These amortisseur windings are interconnected by circular copper rings on the both sides of the rotor. The copper rings are reinforced by steel bands fitted over them. In isolated generator operation, the amortisseur winding serve to reduce excessively high voltage transients caused by line to line system faults , to damp out tensional vibration

originating in the generator drive and to decrease voltage unbalance during unbalance loads. The output windings or stationary armature of the main generator consists of lap wound coil groups arranged in 120 phase belt winding on a slotted laminator stator. These winding are wave connected providing a three phase alternating current output voltage of 200 volts line to line and120 volts line to neutral at a frequency of 400hz revolutions per minute. Both ends of each of the three phase winding are brought out to a terminal board mounted externally on the generator frame. The leads from the three stator phases are brought directly to upper surface of the terminal board, thus aircraft wiring can be clamped directly against them without current passing through the terminal studs.

EXTRACTION FANGENERALThe extraction fan is used for extracting air from toilet and galley compartments.

DESCRIPTIONThe extraction fan consists of an outer casing(180) containing a receptacle connector(30) and a bonding clamp, an inlet basket assay(130),a fan casing(230) and an outlet diffuser(190). The fan casing (230) accommodates an electric motor assay(360). The inlet basket assay(130) and outer casing are provided with outlet pipes. The outer casing(80) is fitted for mounting of quick release clamps. Motor assay: The motor assay (360) is a three phase squirrel cage type four pole induction motor with neutral point. The motor assay consists of a motor housing (540) and two bearing covers (370,510) in which the ball bearing (460) for the rotor shaft assay (470) are supported. The motor housing (540) comprises of the stator and twelve helical type blades each in two stage. The bearing cover(370) contain the preloaded ball bearing(460) which is held in place by the spring washer(420). The ball bearing(460) are grease packed ball bearings types. The stator in the motor housing(540) includes in addition to the wire winding. These thermo switches (one for each wire winding) connected in series, which switch off a contact current for a separate relay, when the temp of stator reaches +160 0c (+3200c); this relay switches off the power supply for the extraction fan. The three thermo switches are self closing switch types.

Fan wheel: The fan wheel (240) comprises of fourteen helical type blades and it is mounted on the front end of rotor shaft assay (470). The fan wheel (240) is secured by a screw (250) with washer (260).

Identification: For identification purposes, each extraction fan is fitted with an identification plate (10). Amendment status: The amendment status is shown in the amendment plate(20) if applicable.

OPERATIONWhen power supply for the extraction fan is switched on, the fan wheel (240) is driven by the motor assay (360), direction of rotation of the fan wheel and Flow direction of the air are indicated by the identification plate (10) on the inlet basket (80). When during the malfunction of the motor assay (360) the temp of the stator reaches +1600c(+3200F) then three thermo switches included in the stator wire winding are disconnected the power supply for the motor assay at a separate place. The extraction fan is used in the following conditions: Continuous operation mode. Free wheel mode, when the motor assay is inoperative.

CURRENT TRANSFORMER ASSEMBLYDESCRIPTIONIt consists of twelve current transformers plotted into one package. The current transformer package contains the sensing current transformers for overcurrent protection, boost current and power limiting for regulator and for metering. There are three transformers for each function for a total of twelve current transformers in each package.

OPERATIONInput signals are obtained by passing the generator output leads through circular openings in current transformer package. The toroidal core transformer of the package is located such that leads of the generator pass through their center and constitute the one turn primaries of current transformers. The secondary of the three transformers are connected to a terminal board that is also plotted into package and then brought out to connector.

TESTING Check each transformer in package separately. Apply 7.5 A , 400 Hz. , current in primary lead going through transformer opening. The voltage developed across the internal resistors should be 5 - 7 volts. A short or open circuit is indicated if voltage is out of limits.

EMERGENCY EXIT LIGHT POWER SYSTEMDESCRIPTION AND OPERATIONThe emergency exit light power supply is a self - contained unit made up of a package of six nickel - cadmium batteries and an electronic switching circuit with voltage regulator. The unit is used to supply power for emergency exit lighting. The unit is connected to the aircraft 28 volt D.C. power system and is controlled by a switch at the pilots and attendants stations. The emergency for which the unit supplies power to exit lights is the loss of aircrafts 28 volts D.C. power. Under different conditions depending upon presence or absence of power supply and on the pilots switch position; following different phenomenon occur in batteries:

In presence of power supply and when switch is OFF; battery is charging. In presence of power supply and when switch is in ARM position; battery is charging. In absence of power supply and when switch is in ARM position; battery supplies power to the lamps. In presence of power supply and when switch is in ON position; battery supplies power to the lamps.

RADIO SHOP

WEATHER RADARTRANSMITTERTo generate high energy radar pulses capable of reaching target up to 300 miles from the aircraft with enough energy to develop readable echo signals.

DUPLEXERTo pass the transmitter pulses on to antenna while blocking any RF from going to receiver. As soon as pulse transmission is complete, the duplexer allows the received echo signals to be routed to the receive circuits.

DIRECTIONAL COUPLERIt routes a small amount of transmitted RF to the automatic frequency control (AFC) circuits and VSWR monitor circuits.

AUTOMATIC FREQUENCY CONTROL CIRCUITTo sense the transmitted pulses frequency and to adjust the Local Oscillator frequency as required so that difference will always be 60 MHz (receiver IF frequency).

OVERLOAD CIRCUITTo protect the modulator when modulation monitor senses a magnetron misfiring.

RECEIVING CIRCUITSTo process the received echo signals that are reflected by moisture bearing weather targets seen by weather radars antenna. These signals are routed from the antenna to the receiver via directional coupler, duplexer, RF limiter, and Tunnel diode amplifier.

RF LIMITERIt blocks transmit pulses from the sensitive TDA and crystal detectors in the video receivers mixer stage.

VIDEO AMPLIFIERIt increases the signal level prior to transmission of video data to the indicator.

CONTOUR MODETo highlight the storm cells area of highest precipitation. The display of these areas shifts from brightest to darkest light level when CTR mode is selected.

BENDIX ANT-1T ANTENNA It is split axis type where roll axis is platform stabilized to verticalreference and pitch axis is line of sight stabilized.

TEST CIRCUITSWhen test mode is selected on the control panel, either in flight or on ground,following actions occurs. A) Contour mode is selected. B) Waveguide switch rotate 90 degrees so tha RF is dissipated into a dummy load. C) R/T s video and stabilization test circuits are enabled. In this mode 1. Pitch and roll test signals are introduced into the stabilization circuits for verifying proper antenna performance. 2. Wave-guide switch has to be properly positioned.

GENERAL: FOR BOEING AIRCRAFTSThis section provides a general description and operation of the Bendix RDR 1F Radar received, transmitter, hereafter referred to as R-T unit or by the full nomenclature.

EQUIPMENTThe RDR 1F Radar Rx Tx is apt of Bendix RDR-!f Airborne Weather Radar system. R-T unit is a lightweight, airborne unit consisting of Receiver, Transmitter, and Power supply. In addition, the R-T unit contains the necessary circuitry to perform sensitivity time control (STC), weather penetration contour, self test, performance monitoring and antenna stabilization functions.

LEADING PARTICULARSCharacterstics Power requirements Transmitter: Carrier frequency Peak power output Output pulse amplitude Output pulse width Duty cycle Receiver: Type RF frequency First LO frequency Description 115 volts +/-10%,400Hz +/5% single phase,505va 9375 =/-40MHz. 65 KW(nominal) 13 KV,12 apm nominal 5 microseconds 0.001

Double conversion, superheterodyne 9375+/-40MHz 60 MHz above receiver RF First IF signal frequency. First IF bandwidth 60 MHz Second LO frequency 4 MHz min at 3 Db points Second IF 46.5 MHz Second IF bandwidth 13.5 MHz Dynamic range 0.8 MHz at 3 db points Minimum Discernable signal 26 db minimum level -104 dbm Level(MDS) -102 dbm

OVERALL OPERATIONRDR-1F Rx Tx is a component of the RDR-1F Airborne Weather Radar system. The RDR-1F Airborne Weather Radar system is designed primarily to furnish continuous enroute weather information relative to cloud information, rainfall rate, thunderstorm and areas of turbulence and icing conditions.

The RDR-1F system operates by emitting very short intense pulses of microwave energy, which are reflected within the range of the system. A portion of the radiated energy reflected by and objects having reflective characteristics are returned along the same general path to the aircraft where it is received and converted into a visible indication of range and azimuth bearing. These indications are viewed as luminous spots on the areas on PPI scope. The sensitivity time control (STC) circuit ensures that echo signals are displayed with approx equal intensity from similar targets from near zero range that is consistent with the overall system performance index. STC range settings, set at factory, can be modified as required if antennas smaller than 30-inch sizes are utilized. Penetration compensation circuit in compensation with STC circuit allows storm cells behind intervening cells to be displayed nearly to their true size as though no intervening cell was present. Iso-echo contour circuitry provides a means for determining the relative density of rainfall areas. With contour operating, the pilot can see storm areas in his flight path and can also distinguish corridors of relative calm through the storms. Contour circuitry detects strong returns in high-density rainfall areas and converts them into dark areas on the radar display. Lighted areas or ring surrounds these dark areas, which represent areas of lower rainfall rates. This provides a visual indication of high turbulence areas that should be arounded.

PHYSICAL DESCRIPTION: FOR 320 AIRCRAFTThe receiver transmitter is packaged in a full short ATR case designed to be mounted on a rack that provides cooling air. The front panel of the receiver and transmitter contains maintenance controls and adjustments. Also located on the front panel is the LED fault code read out. The input output wave guide and electrical connections are located on rear 1. 2. Power requirements Operating frequency 115 volt ac, 4 amp 9345 +/- 15 MHz.

3.

4.

Transmitter Characteristics: Power Output (Transmitter module) Pulse width Pulse repetition frequency Receiver Characteristics: I F frequency Overall noise figure

150 W nominal 3,6,10,30 microseconds. 1620,720,180 MHz 53 MHz 5 db

FUNCTIONAL OPERATIONThe basic function of R-T is to generate a high level RF pulse to be radiated from the antenna and to amplify the echoes received from the antenna to a level that can be used by R-T processing circuits prior to application to the indicator or EFIS display. The transmitter at a peak power level of approx 150W generates the RF pulse. Low-level received signals are converted to a n IF frequency, amplified and processed before being routed through R-T to the indicator. The transmitter output pulse (at a frequency 9345 WHZ) is directed to the antenna by the circulator. The antenna flat plate radiator transmits this RF energy in a narrow pencil beam that is perpendicular to the plane of the radiator surface. An area of either 90 degrees (180degrees or normal scan mode) or 45degrees(sector or scan mode), left and right of the forward direction of the aircraft, may be scanned by the antenna in its Search for weather and ground targets. When a target is encountered, RF energy is reflected from target back to the antenna during the time between transmitted pulses. The circulator to RT receiver directs this received RF energy where it is processed for visual display. The RT receiver down converter utilizes a dielectric resonator stabilized oscillator (DRO) as a local oscillator (LO) operating at 9398 WHZ. The mixer combines the LO frequency with the 9345 MHz received signal to produce 53 MHz IF frequency. The down converter output is maintained at exactly 53 MHZ by the AFC circuit. A frequency discriminator develop an error signal that is amplified, filtered, integrated and used to control the frequency of DRO so that it is always 53 MHz from the tx frequency. The TR limiter, at the 1/P to the receiver, protects the rx from the relatively large transmitter pulse that leaks through the

circulator. The53 MHz signal from the down converter is amplified and detected in the IF amplifier. The resultant video is processed in the R-T and sent the indicator. The Bandwidhth of the IF amplifier is changed when the transmitter pulse width is changed to maintain an optimum display resolution. The STC circuit reduces the sensitivity of the receiver for close in targets so that their strong echo returns appear the same as weaker returns from more distant targets. The STC signal is also controlled with the manual gain control. When the manual gain control Increases past its midpoint, it starts reducing the STC signal. The reduced STC signal then increases the rx gain for the short ranges. The blanking pulse prevents the video resulting from the transmitted pulse from appearing on indicator display. The digitized outputs from the IF amplifier are smoothed and stored in range cells by the data processor. Range cells outputs of the data processor are transmitted over independent data buses to a maximum of three indicators for display in response to commands from the data bus microprocessor. The data bus microprocessor also optimizes data processor parameters to match information received on control buses. The data bus/BITE (Built in Self Equipment) microprocessors is used to service the 429 interfaces buses and route received information to the proper subunit. It also controls BITE testing. Azimuth drive signals for controlling the azimuth motor are generated by data bus microprocessors. The stabilization microprocessor performs calculations of the elevation pointing command necessary to keep the antenna pointing at a fixed tilt angle in space independent of pitch and roll motions of the aircraft. The stabilization microprocessor receives pitch and roll signals from the stabilization bus via the data bus microprocessor. Elevation command output goes to the elevation drive servomechanism to control the antenna elevation axis. System configuration programming is provided by means of programming pins to configure the R_T for use in single or dual RT installations with a maximum of 3 indicators. Other programming pins are provided to select analog or digitized stabilization inputs.

RADAR ANTENNAANT-1T Radar Antenna is designed to perform two primary functions:-

1. To accept and condition r-f energy to form a narrow beam radiation pattern. 2. To receive the returning echo and relay it to the radar receiver incurring only minimum losses. 1. 2. 3. 4. 5. 6. 7. Operating Band Operating frequency Reflector Diameter Focal length Beam width Power capacity Voltage requirements X Band 9375 +/-40 MHz 30 inches 12 inches Pencil 3 db point 2.9 degrees nominal Up to 85 KW 115 volt, 380-420 Hz single phase and reference phase for azimuth, tilt, and roll drive systems. 26 volt, 380-420 Hz single phase and reference phase for antenna stabilization synchros. 28 volt dc to energize ferrite rotator and energize the stabilization relay

The ANT-1T radar antenna is an X band, sector scan weather radar antenna with pitch and roll stabilization about two separate axis. The antenna is platform stabilized in roll upto +/- 43 degrees and line of sigh stabilized in pitch up to +/-25 degrees. Pitch information from the aircraft vertical gyro, after processing by the pitch servo loop is applied to the ANT-1T as the control phase voltage for the tilt drive motor. The tilt drive motor repositions the antenna to counteract pitching of the aircraft. Roll information from the aircraft vertical gyro is applied to the stator of the roll synchro is processed by the roll servo loop and applied to the roll motor generator to reposition the antenna to counteract roll of the aircraft. The function selector on the CON-1R and CON-1S control selects either vertical or horizontal polarization of the transmitted r-f energy as it leaves the antenna field. During antenna mapping operation, the r-f energy is vertically polarized and the antenna parabolic reflector forms a pencil beam of radiated energy. For ground mapping the r-f energy is horizontally polarized and it sees a beam forming a grid located on the upper half of the

reflector. As a result the r-f energy is formed into a beam that spread over a large elevation angle thereby especially illuminating a large strip of terrain at each azimuth position .A ferrite rotator located in the antenna feed, is used to rotate the rf energy. When the function selector is in map mode, 28 V dc is applied to the energizing coil of the ferrite rotator causing the ferrite causes the polarization of the r-f energy to be rotated 90 degrees. In this manner, the transmitter r-f energy is changed from a pencil beam to a mapping beam. Antenna for Boeing 737 The parabolic reflector works on a similar principle to a car headlamp reflector. Energy striking the reflector from a point source situated at the focus will produce a plain wave of uniform phase traveling in a direction parallel to the axis of the parabola. The feed in a weather radar parabolic antenna is usually a dipole with a parasitic element, which, of course, is not a point source. The consequence of a dipole feed is that the beam departs from the ideal and there is considerable spill-over, giving rise to ground target returns from virtually below the aircraft, the socalled height ring. Antenna for A320 The flat plate antenna consists of strips of waveguide vertically mounted side by side with the broad wall facing forward. Staggered off-center vertical slots are cut in each waveguide so as to intercept the wall currents and hence radiate. Several wavelengths from the antenna surface, the energy from each of the slots in summed in space, cancellation or reinforcement taking place depending on the relative phases. In this application the feed to each slot, and the spacing between slots, is arranged so as to give a resultant radiated pattern, which is a narrow beam normal to the plain of the plate. The greater the number of the slots the better the performance; since the spacing between the slots is critical we can only increase the number of slots by increasing the size of the flat plate.

WEATHER RADAR- SCHEMATIC DIAGRAM

SELCALSYSTEMThe SELCAL system alerts the flight crew that a ground station is calling the Airplane. The SELCAL (selective call) system receives SELCAL codes from the airplane communication receivers. When the assigned airplane code is received, the SELCAL system alerts the flight crew with visual and signals.

SYSTEM COMPONENT LOCATIONSThe SELCAL system has two independent channels and consists of the following components: 1- Control System 2- Dual Decoder. The control panel is on the forward electronic equipment rack.

CONTROL PANELThe control panel provides visual alerts and reset capabilities for both SELCAL decoder channels. The alert light (green) comes on whenever that decoder channel has received a radio signal with properly coded audio. The RESET push button is a dual switch, which resets both decoder channels simultaneously and turns off the alert light(s) and HI-LO chime.

DECODERThe SELCAL Decoder senses inputs from communication transceivers and determines if the airplane assigned 4 letter code is being received on either one or both channels. If the coded transmission is the airplane code, the SELCAL decoder activates the aural warning device and illuminates a light. The decoder unit consists of dual decoder chassis mounted in one unit. The unit is contained in a ATR short case and weighs10.8 lbs. (4.9 Kg). Code selector switches are mounted on the front panel of the decoder unit. Channel 1 switches are on the left, facing the panel and Channel 2 switches are on the right. Each switch has letter codes from A through M with the exception of letter I.

OPERATIONSELCAL no.1 and SELCAL no.2 are independently power by switched electronics busses no.1 and busses no.2.The operating description that follows applies to both SELCAL Channels. The audio to the decoder is from the VHF or HF receivers. The SELCAL code is a combination of audio tones. The assigned

airplane 4 letter SELCAL code is set in the decoder. This selects the required tone decoder reeds from the 12 available reeds.

DISTANCE MEASURING EQUIPMENT (DME)

DESCRIPTIONDME is a system combining Ground Based and Airborne Equipment to measure the distance of the aircraft from a ground station. It is used primarily for Position Fixing, Approach to an Airport and Figuring Ground Speeds.

The Airborne DME consists of a Receiver-Transmitter (RT), a Control Unit, a Distance Indicator and an Antenna. The Ground Based DME consists of a Receiver-Transmitter and an Antenna that operates only on a single frequency.

RECEIVER-TRANSMITTER (RT)The transmitter section of the RT unit contains all the necessary circuits to generate, amplify, and transmit the interrogating pulse pairs. The receiver section of the RT unit contains the necessary circuits to receive, amplify, and decode the received pulses.

CONTROL UNITThe control unit provides the necessary control and switching circuits for the airborne DME-RT. The control unit may also provide the frequency selection for a VHF communication or navigation receiver.

DISTANCE INDICATORThe distance indicator displays the aircraft distance in nautical miles from the ground station. The indicator will also display in the form of flag offer dashes, a warning that the system is malfunctioning.

ANTENNA The antenna is a single L - a band type antenna thattransmits and receives with an omni directional radiation pattern.

OPERATION The pilot selects a VOR frequency. The selected frequency automatically selects a DME channel that is paired with that frequency.

The receiver-transmitter of the airborne DME transmits coded interrogating pulse pairs to the ground station. The ground station receives these pulse pairs, delays 50 s, and then transmits a coded reply pulse pairs back to the airborne DME-RT.

The airborne receiver-transmitter receives the reply pulse pairs and verifies that the pulse pairs are valid. The airborne DME-RT computes the slant range (Line-Of-Sight) distance from the ground station as follows: D = {(T 50 s)/12.359} Where; miles from D = Slant range distance in nautical the ground station. T = Time in s b/w transmission of interrogation pulse pair and reception of corresponding reply pulse pair. 50 s = Delay at ground station b/w reception of DME interrogation & transmission of reply. 12.359 = Time in s for RF energy to travel one nautical mile and return. The distance is then sent to the distance indicator where it is displayed for pilots use.

PRERECORDED ANNOUNCEMENT AND BOARDING MUSIC REPRODUCER

(PRAM)DESCRIPTIONThe PRAM is manufactured by THE VIDEO SYSTEMS DIVISION OF MATSUSHITA ELECTRICAL INDUSTRIAL COMPANY LIMITED of OSAKA, JAPAN. The functions of the PRAM are to play prerecorded messages to the passengers and to be a source of prerecorded music programs for the passengers through the aircraft Passengers Address System (PAS) or Passenger Entertainment System (PES). The reproducer requires a remote control unit for message number programming, boarding music ON/OFF and volume control and other Remote/Manual control operations by the attendant. Following are some of the useful accessories provided with PRAM: -

TAPETape is an Audio Cassette. When this term is used as a modifier, for example, the Tape Message or the Tape Music, it means that message or music that is once recorded on an audio cassette tape.

SSSV (SOLID STATE STORED VOICE)When this term is used as a modifier for example, SSSV message, it means that the message is once encoded into digital signal stream and stored in read only memory in the SSSV circuitry.

MESSAGEMessage is a continues speech block, reproduction of which is initiated by manual operation on the remote control unit or one of the external switches connected to the reproducer other than the cabin decompression sensor switch.

EMERGENCY MESSAGEIt is a continues speech block, reproduction of which is initiated by operation of the Cabin Decompression Sensor Switch. This speech block may contain a single message or a number of same messages in different languages. The emergency message is reproduced for a predetermined number of times on one initiation.

ANNOUNCEMENT

It is a collective term referring to messages of all types, used to distinguish them from music. Sometimes, it refers to an ordinary message only to distinguish it from an emergency message.

BOARDING MUSICIt is an Audio (usually Music) that is provided from the Boarding Left and Right Outputs of the Reproducer.

PASSENGER ENTERTAINMENT SYSTEM (PES) MUSICIt is an Audio (usually Music) that is provided from the PES channel 1 to 4 Outputs of the Reproducer.

MUSIC TAPEIt is an Audio Cassette Tape prerecorded in a special Four-Track arrangement, played back in one direction only, contains four channels of audio as boarding or PES music. These four channels of audio are two stereo programs. The tracks 1 & 3 and tracks 2 & 4 are paired with track 1 and track 2 being their respective left channels.

ANNOUNCEMENT TAPEIt is an Audio Cassette Tape prerecorded in arrangement, played back in one direction 256 different Messages sequentially from Track 3 and/or 4 may contain a number of messages. a special Four-Track only, contains up to track 1 to track 4. identical emergency

OPERATIONTwo identical Announcement Tapes are loaded on Announcement Tape Decks A and B. A pair of Music Tapes are loaded on Music Tape Decks A and B. Internal settings are necessary for the music tapes and also for the emergency message. Power to the reproducer is turned on. A built-in test equipment (BITE) test sequence takes place for about one second, during which DC power line voltages, announcement tape deck track selection and a part of the Input Interface are checked for normal operation. When the announcement tapes are new, an initialize mode is activated either from the RCU or by setting the PRAMs front panel switch to TAPE INITIAL and then back to NORMAL.

SCHEMATIC DIAGRAM

EMERGENCY LOCATOR TRANSMITTER (ELT)

DESCRIPTIONARTEX ELT 110-406 is an automatic activated Emergency Locator Transmitter (ELT). It is a device to detect the aircraft after it has been crashed. It can be manually activated via the MANU-OFAUTO switch on the unit, or via the optional remote AUTO/MANU switch on the front of the aircraft. It gets activated automatically with the longitudinal thrust of 5g for 55 milliseconds. It is an Orange plastic box of (216*82*60) mm dimension, fixed on a mounting tray and locked by a metallic strap with Quick Operating latch.

OPERATIONThe ELT is equipped with an impact g switch that will automatically activate the transmitter when a g forces of at least 5g is applied to the longitudinal axis of the aircraft, from nose to tail for 55 milliseconds. Due to this it transmits the standard swept tone on 121.5MHz and 243.0MHz. The 406.025MHz transmitter turns on every 50 seconds for 440 milliseconds (standard short message) or 520 milliseconds (optional long message). During this time an encoded message is sent to the satellite. The information contained in this message is shown below: Serial number of the Transmitter. Country code. I.D. code. Position coordinates (optional). The 406MHz Transmitter will operate for 24 Hours and then shuts down automatically. The 121.5/243.0MHz Transmitter will continue to operate until the unit has exhausted the battery power, which typically is at least 72 Hours. One of the advantages of the 406MHz transmitter is that it will produce a much more accurate position, typically 1 to 2 Km as compared to 15 to 20 Km for 121.5/243.0MHz Transmitters. It also transmits a digital message which allows the search and rescue authorities to contact the owner/operator of the aircraft through a database.

Information contained in the database that may be useful in the event of crash is shown below: Type of the aircraft. Address of the owner. Telephone Number of the owner. Aircraft registration number. Alternate emergency contact. Once the ELT is activated and the 406MHz signal is detected from the satellite and a position is calculated, the 121.5/243.0MHz transmissions are used to home in on the crash site.

LOW RANGE RADIO ALTIMETER (LRRA)GENERALIt measures the distance from the airplane to the terrain and provides a read out of altitude in the cockpit. This device is only operational from 0 to 2500 feet and is of low range. This device is only used during takeoff and landing. Its frequency of operation is 4300MHz

OPERATIONIt measures altitude by transmitting a signal to the ground and comparing this signal to the reflected from the terrain. Reflected signal is received by the receive antenna and routed to the receiver/transmitter. The R/T compares the reflected and transmitted signal and converts the difference to a signal to drive the indicator and R/T also supplies to other airplane system. Transmit Antenna LRRA LRRAindicator

Receive Antenna

Receiver Transmitter

Altitude to other systems

The R/T antenna transmits a varying frequency signal to the transmit antenna and receives the signals. Signals are then mixed together to develop a difference frequency .the difference frequency is inversely proportional to the time it took the transmit signal to travel to the terrain and back to the plane .the frequency difference is converted to a DC voltage which is the output from the R/T. the R/T provides to separate DC o/p signals the frequency of operation is 4300MHz modulated at a 160mz rate.

SYSTEM THEORYThe system transmits an FM signal of linearly low varying frequency which when delayed by round trip of the ground and then mixed with a portion of the transmitted signal gives a

difference frequency proportional to the aircraft height above the ground+

The transmitted frequency (ft) is 4300 -A MHz at some point in time t1 the emitted frequency f1is transmitted to ground and the reflected signal received at LRRA mixer at time (t2) During time (t2 t1) the transmitted frequency has increased to a new frequency f2 when f1 and f2 are mixed at time (t2) the output of the mixer is the difference frequency ( f) which is proportional to ( t) and also altitude (H). By means of proper calibration, this difference in frequency is converted to a DC voltage and sent to interfacing systems

Transmitter

altitude output

Mixer

Freq to DC counter

f

Difference frequency Transmit Antenna Receive Antenna

R/T powered by 115v AC The transmitter applies a variable frequency output to the transit antenna and through coupling through the mixer. The reflected signal is reflected from terrain back through the receiver antenna through the mixer. The mixer output is a frequency difference, which is sent to two identical frequency counters.

The Frequency counters convert the difference frequency to a DC voltage that is proportional to airplanes altitude. The Radio altimeter contains a C band transmitter which is frequency modulated 70 MHz on each side of 4.3 GHz at a rate 155 Hz i.e. the transmitter output frequency sweeps from 4.23 to 4.87 GHz and back 155 times a second. The transmitter output is conducted to the transmitting antenna with a small sample coupled to a mixer. The transmitted signal is reflected from the ground, picked by the receiving antenna, filtered (to eliminate possible interference from the ground radars and other interference sources) and also applied to mixer as an injection frequency. The time required for the signal to travel from the transmitting antenna to the ground and back to the receiving antennal, is directly proportional to the distance from the aircraft to the ground. The mixer output is the difference between the transmitting frequency being coupled directly to the mixer and the frequency of the reflected transmitted signal picked up by the receiving antenna. Therefore, the mixer output is directly proportional to the time required for the signal to travel to the ground (altitude). The difference frequency will be the same for any particular altitude whether the transmitted frequency is sweeping up towards 4.37 GHz or down towards 4.23 GHz. The amount of time required for the transmitter frequency to reverse its direction of sweep is immaterial. The difference frequency output from the mixer is counted and converted to a DC analogue voltage, which actuates the point on the Radio Altimeter indicator to indicate altitude above the ground at which the aircraft is flying.

TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS)

DESCRIPTIONTCAS-2 is an airborne traffic alert and collision avoidance system that operates without support of ATC ground station. The system detects the presence of the nearby intruder aircraft equipped with the transponders that reply to the modes interrogation. TCAS-2 displays the nearby transponder equipped aircraft on a Traffic Advisory Display and during threat situation provides Traffic Advisory Alerts and vertical maneuvering resolution advisories alert to the pilot in avoiding MID AIR COLLISION. Detection and Tracking of intruder aircraft is performed via transmission and reception on top and bottom mounted TCAS-2 directional antenna.

OPERATIONBy using data (Range, Relative Bearing, Altitude, Vertical Speed and Closure Rate) the TCAS-2 predicts the speed and flight path of tracked aircraft to determine if there is any chance of collision. If the intruder aircraft enters the safe boundary of the particular TCAS-2 equipped aircraft then the TCAS-2 will issue a Traffic Advisory (TA), to alert the crewmembers that closing aircraft is in vicinity. This advisory gives the pilot about 30-45 sec as a reaction time to react before some mishap happens. If the intruder aircraft continues to approach the particular TCAS-2 equipped aircraft then the TCAS-2 will issue a Resolution Advisory (RA) to maintain safe vertical separation between two aircraft. This advisory gives the pilot about 10-25 sec as a reaction time. Two TCAS-2 equipped aircraft will coordinate their advisories using a mode S transponder data link. The crew then promptly but safely and smoothly follows the advisory. Since the maneuvers are coordinated therefore the crew should never maneuver in opposite direction of the advisory.

AIR TRAFFIC CONTROL (ATC) SYSTEMDESCRIPTION

The ATC system contains Airborne Components controlled by Ground Facilities to identify the Flight Number and Altitude of the Airplane. In addition, the Ground Facilities monitor the Airplanes Location and Direction of Travel. The Power Requirement Is 115V AC. Transmit Frequency is 1090+3MHz and Receive Frequency is 1030+0.2MHz.

The Airborne Equipment consists of a Transponder (ReceiverTransmitter), a Control Unit, an Antenna, and a Digitizer. The Ground Based equipment consists of a Primary Radar System and a Secondary Surveillance Radar (SSR) System. The Primary Radar System consists of an Antenna, a ReceiverTransmitter (RT) and an Indicator. The Primary Radar System works like other Radar System. A narrow RF type beam, transmitted through a rotating Antenna, is reflected by any targets in its path and returned to the Antenna. By calculating the elapsed time between transmission and reception of the RF beam, the distance to the target is determined. This information is displayed on the Two Dimensional Radar Screens. The SSR system consists of an antenna, a ReceiverTransmitter (RT), and the Necessary Interface and Control Equipment for the ground station. The SSR system is used to separate aircraft by altitude. The SSR system uses an Antenna that is mounted directly to the Primary Radar Antenna and pointed in the same direction or synchronized with the same rotation as the Primary Radar. This system interrogates the aircraft about its identity and altitude by transmitting two sets of pulses.

TRANSPONDER The Airborne Transponder is an important part of the Air Traffic Control System. The safety of passengers, aircraft and the crew members depends on the ability of the air traffic controllers to accurately identify and locate aircraft within controlled airspace. A Transponder is the Airborne Receiver-Transmitter (RT) portion of the ATC Radar Beacon System that sends an

identifying coded signal, in response to a transmitted interrogation from a ground-based radar station, in order to locate and identify the aircraft. Air Traffic Controllers use the coded identification replies of transponders to differentiate between the aircrafts displayed on their radar screens. This aids the controller in maintaining aircraft separation, collision avoidance, and distinguishing types of aircraft.

RECEIVER-TRANSMITTERThe Receiver portion of the Transponder contains the necessary circuitry to Receive, Demodulate, Amplify, and Decode the Interrogation Signal. The Transmitter portion of the Transponder contains the necessary circuitry to Encode, Modulate, Amplify, and Transmit the coded reply signal. The Transponder also contains the circuitry required for checking the validity of the received interrogation signal and monitoring the integrity of the Transponder.

CONTROL UNITThe control unit contains the circuits and controls necessary to select the identifying codes. It also contains the controls necessary for selecting an altitude source, initiating self test condition, and selecting the transponder reply mode.

DIGITIZERThe digitizer is a simple converter that converts an Analog Signal, representing Barometric Altitude, to a Digital Format. The digitized barometric altitude can then be encoded and transmitted as a part of the reply signal.

ANTENNAThe Antenna is a Monopoly Blade Type Antenna. The antenna is usually mounted in an area of the aircraft that will not be shielded from the interrogation. This prevents the aircrafts identification from disappearing from the controllers radar screen.

OPERATION The Pilot selects an identification code from the Air Traffic Controller.

The SSR system transmits a coded interrogation signal as the Primary Radar System detects the aircraft. The interrogation signal is received, detected and decoded by the Airborne Transponder. The Transponder then encodes and transmits a set of reply signals. The reply signal is then received, decoded and displayed at the ATC ground station screen.

HIGH FREQUENCY (HF) SYSTEM

DESCRIPTIONThe HF communication system provides long range communication between: The Aircraft and Ground Stations. The Aircraft and other Aircraft. The system operates in the 2 to 30 MHz frequency range in Amplitude Modulated or SSB mode to transmit and receive information that can be in the form of a transmitted voice or a coded digital signal. The HF system uses the skip distance phenomena to achieve long distance transmission. Skip distance transmission is most effective in the 2 to 30 MHz ranges and varies with frequency and time of day. The HF communication provides a reliable way to transmit and receive Flight Information, Landing Instruction and Voice Communication. There are two HF communication systems HF-1 and HF-2 installed in the aircraft. Each HF communication system is composed of one receiver-transmitter, an antenna coupler, lightning arrester, an antenna, a remote control unit, a microphone, a speaker or handset and necessary relays. The HF1&2 communication systems use 115V, 400Hz, 3-phase primary power from 2.0000 to 29.9999 MHz or 2.8000 to 23.9999 MHz on channels spaced at 1KHz or 100Hz.

RECIEVER-TRANSMITTERThe RT provides transmit signal and power to the antenna, and processes the received signal from the antenna. The HF RT operates in the 2.0000 to 29.9999 MHz range providing USB and LSB data communication. The RT contains the following circuit, a frequency synthesizer, a receiver, an exciter, and a power supply. The transceiver may also contain circuit to monitor the VSWR of the antenna and transmission line. The receiver portion consists of the necessary circuit to receive, demodulate, amplify, and filter the received signal. The exciter contains the circuit to excite, modulate, amplify and transmit the voice or coded data communication.

ANTENNA COUPLERThe HF probe operates in 2.0000 to 29.9999 MHz region of the radio spectrum. An antenna coupler is required in HF system to maintain the quality of the received signal over a wide range of frequency by matching the antenna to the transmitter.

HF COMMUNICATION SYSTEM

VERY HIGH FREQUENCY (VHF) SYSTEM

DESCRIPTIONThe VHF communication system provides short-range (Line Of Sight) voice communication between: The Aircraft and Ground Stations. The Aircraft and other Aircraft. VHF communication system transmits data in the form of codes. The frequency range allocated for commercial aviation VHF communication is 118.00 to 136.975 MHz in 25KHz increments. The VHF communication system receives RF energy via the antenna, processes the RF and sends resulting audio to the Audio Integrating System. The VHF communication system also sends an audio signal to the SELCAL system to alert the flight crew of an incoming call. During transmission, the VHF communication system and RF energy transmitted via the antenna process the microphone audio from the Audio Integrating System. There are two VHF communication systems installed and wiring provisions for a third, and each system contains the following components: 1. CONTROL PANEL: The control panels are located on the AFT electronic panel. 2. TRANSCEIVERS: The transceivers are located on the E2-1 shelf in the electronic equipment compartment. 3. ANTENNA: The number-1 antenna is located on the top of fuselage and the number-2 antenna on the bottom. The VHF communication system requires a primary input voltage of 27.5V DC at 1 ampere in receive mode and 27.5V DC at 6 ampere (7.5 ampere maximum) in transmit mode.

OPERATION The pilot selects the communication frequency of the Airport or Air Traffic Control that is responsible for the airspace in which the aircraft is flying. The pilot actuated the transmitter section of the RT by pressing the microphone key. The pilot's voice is transformed into an electrical signal, amplitude modulated and mixed to produced a signal that is radiated from the transmit antenna.

The receiver section of the ground station RT receives the radiated signal, filter and mixes it and then detects the modulated signal. A speaker or headset then transforms the modulated signal into an audio signal. The Airport or Air Traffic Controller replies using the same frequency.

VHF COMMUNICATION SYSTEM

VERY HIGH FREQUENCY OMNI-DIRECTIONAL RANGING SYSTEM (VOR)

DESCRIPTIONThe VOR-700 is a solid-state Microprocessor controlled Very High Frequency Omni-Directional Range/Marker Beacon Receiver. VOR Receiver is airborne equipment and is in the aircraft and its corresponding indicator is in the cockpit and its counterpart transmitters are along the runway.

VOR PRINCIPLESINE WAVE COMPARISON : Sine Waves are used to illustrate the radiation of the VOR audio signals and to compare the phase differences of two audio signals (30Hz). Both the 30Hz Reference signal and Variable signal radiated from ground station, combines in the space into a Rotating Cardioid. At receiving end, the Rotating Cardioid is demodulated by a demodulator on board in the aircraft to give 30Hz signal of variable phase produced by the

Rotating Antenna on the ground and also an Omni-Directional signal carrying a 30Hz signal of a fixed phase. The receiver detects the phase shift between the signals and gives the relative reading on the indicator as to where the aircraft is flying relative to the VOR station.

DIFFERENT MODES OF VOR1. AUTOMATIC VOR: In this mode the pilot has to select or tune to a particular frequency and corresponding Bearing Information is displayed on the Radio Magnetic Indicator (RMI). 2. MANUAL VOR: In this, the display is on Course Deviation Indicator (CDI). The CDI has a needle, which is divided, into three parts Upper end, Lower end and Course Deviation Bar. The Indicator also has an Omni Bearing Selector Switch.

OPERATIONLet us take, that we are flying at 0o north and now we have to go 60o northeast. There is Lubber Line, which always points out to the position where the nose of the aircraft is. Now in order that the aircraft should fly to 60o northeast the Lubber Line should coincide with it. Now the pilot has to align its aircraft accordingly. The pilot will select the 60o on the selector and the Upper end and the Lower end align themselves with 60o but the middle bar will not, it will remain either above or below the point. Now the pilots main task is to align the three parts of the Line. It must be noted that the straight line is divided into number of equal parts and has spacing equal to the 5o. Therefore the line above or below the center point gives the angle to which the plane must be turned around. As the pilot turns the aircraft, the central bar starts aligning itself with other two and when all are aligned it means the plane is heading due course (60o northeast).

AUTOMATIC DIRECTION FINDER

(ADF)

DESCRIPTIONThe Automatic Direction Finder (ADF) is the oldest and most widely used radio navigation system. The automatic direction finder (ADF) is an airborne system used to determine the relative bearing from the aircraft to the ground-based transmitter (with respect to the aircraft centerline). ADF is the oldest of the radio navigation systems and one of the most widely used throughout the world, because of the availability of numerous ground stations to tune the ADF. The concept of the ADF navigation is based on the ability of the airborne system to provide the bearing indication with respect to the aircraft centerline, based upon the direction of arrival of radio wave from a selected station. If the indicator compass card is adjusted so the aircraft's present heading is set below the Lubber Line (centerline), then the indicator pointer against the compass card provides a direct magnetic bearing to the station. ADF has been in use longer than most radio navigational aids, and its use has become quite common. ADF was first mandatory aboard commercial air carriers in 1937. The simplicity of the system and its independence of other systems are two reasons for it's continued use. ADF is used by itself or in conjunction with a VOR system. Although VOR systems may be more accurate, ADF has an advantage in that there are more ADF ground stations (low frequency beacons and standard broadcasting) than there are VOR stations. This means that finding an ADF station close to the direct line from the city A to city B is much more likely than finding a VOR station in the same path. ADF is used for navigation, position fixing, and position holding.

THE ADF SYSTEMThe airborne portion of the ADF system consists of a Receiver, Control Unit, Indicator, Fixed Loop Antennas, and a Sense Antenna. The ground facility consists of a Transmitter and an Antenna. A typical ground facility used for ADF would be an AM radio station or a Non-Directional Beacon (NDB).

RECEIVER

The ADF receiver contains the necessary circuits for the reception and processing of radio signals (in the 190 to 1750 KHz range) to provide relative bearing information to an indicator. The receiver also contains the circuits required to confirm the validity of the received signals and the reliability of the receiver itself.

CONTROL The ADF control unit provides the control and switching circuitsto select the ADF receiver operating mode and frequency.

INDICATORSThere are several types of indicators that can be used with the system. All indicators used with the ADF system indicate the bearing of the ground station. That is, the needle of the indicator always points to the station that the receiver is tuned on. An ADF indicator will have the needle rotating against a fixed Azimuth card. This type of indicator was also called as radio compass indicator.

ANTENNASThe ADF receiver requires two types of antenna. An OmniDirectional sense antenna is required to help tune the receiver and loop antenna is required to provide the bearing. Depending on the ADF system used, there are different antenna types (older systems require older antenna types). These include loop antennas that are mechanically rotated, electrically rotated, or mounted in a fixed position to the aircraft.

OPERATIONThe 51Y-7 is the principal component of the Automatic Direction Finding System. The system may be operated in either ANT Mode or ADF Mode.

ANT MODE: When 51Y-7 is in the ANT Mode, the Loop Antenna circuits and the input to the Bearing Servo System are disabled. The signal received by the Sense Antenna is applied to the 51Y-7 and processed to produce an audio output to the aircraft audio system.

ADF MODE: In the ADF Mode, a signal received by the Fixed Loop Antenna is applied to the 51Y-7 through a Quadrantal Error Corrector. The signal received by the

Sense Antenna is also applied to the 51Y-7. The Loop Sense signals are combined and processed in the 51Y-7 to produce bearing information and audio output to the aircraft audio system.

INSTRUMENT LANDING SYSTEM (ILS)DESCRIPTIONILS guides the aircraft as to where to land on the runway. The total system comprises of three parts, Localizer, Glidescope and Marker Beacons, each with a Transmitter on the ground and Receiver and Signal Processor in the aircraft. ILS uses Microprocessor circuits to process these signals and provides Digital Data to the Pilot and Automatic Flight Control System through use of ancillary equipment. The LOCALIZER provides Lateral Steering for both the Front-Course and the Back-Course approaches. The GLIDESCOPE provides Vertical Steering for the Front-Course only. The MARKER BEACONS give the Distance Checks. International Civil Aviation Organization (ICAO) has defined three categories of visibility, the Third of which is subdivided. All are defined in terms of Runway Visual Range (RVR) and, except category HI, Decision Height (DH). Due to some problems, category III has not been installed in India. The various categories are defined in Table 1. Table 1: ICAO VISIBILITY CATEGORIES. CATEGORY VISUAL (DH)I II

DECISION HEIGHT 60m (200 ft) 30m (100 ft) -------------------------------

RUNWAY RANGE (RVR) 800m (2600 ft) 400m (1300 ft) 200m (650 ft) 30m (100 ft) ZERO

IIIA IIIB IIIC

OPERATIONThree separate Ground Station Transmitters are required to radiate the ILS guidance signals required for Instrument Landing System. LOCALIZER TRANSMITTER: The Localizer Transmitter radiates frequencies in the range from 108.1MHz to 111.9MHz in odd decimals to provide HORIZONTAL DATA. The Localizer Transmitter Antenna radiates two intersecting Lobes. One Lobe is to the; Right of Runway and the other is to the Left. The Lobe to the Right of Runway is modulated with 150Hz Frequency and to the left is modulated with 90Hz. When the modulated signals are equal, this means that the aircraft is flying towards the middle of runway. Otherwise the airborne equipment detects the 90Hz and 50Hz tone and hence cause the deviation indicator to show a Fly-Left or Fly-Right command.

GLIDESCOPE TRANSMITTER: Glidescope Transmitter the generates frequencies in the range from 329.15MHz to 335KHz spacing to provide vertical guidance data. Glidescope path is an angle of descent for any instrument landing. The glidescope transmitter antenna radiates two intersecting lobes, one lobe is above glidescope path and the other is below. The upper lobe is modulated with 90Hz frequency and the lower lobe is modulated with 150Hz. When the modulated signals are equal, this means that the aircraft is flying at the proper angle of descent. Otherwise the airborne equipment detects the 90Hz and 150Hz tone and hence provide the resultant Glidescope deviation signal to the indicator and also to the Automatic Flight Control System. The Pilot ten aligns its aircraft to the indication on indicator in order to ensure safe landing.

MARKER BEACON SYSTEMThe purpose of this system is to indicate to the flight crew that the airplane is passing through a particular geographical location or points along an Instrument landing path. It provides both type of indication: Audio Indication Visual Indication Three indicator lights provide visual identification.

Oral identification is provided by three audible tones to the interphone system. Receiver frequency is 75MHz.

THE MARKER BEACON SYSTEM LOCATION In India we use three Markers. These markers are located some distance away from the runway. Three are located before the runway i.e The Marker Beacons are: 1. Outer Marker. 2. Middle marker. 3. Inner marker. The location of these markers from the runway (O. M) Outer Marker: Approx. 3.9 NM from the Glideslope Its Radiation is Modulated at 400Hz. The Blue Lamp will glow in the Cockpit giving two dashes/sec. (M. M.) Middle Marker: Approx. 0.5NM from the Glideslope. Its Radiation is Modulated at 1300Hz. The Amber Lamp will illuminate in the Cockpit with one dot dash in every two third second. (I. M) Inner Marker: Approx. 0.1NM from the Glideslope, Its Radiation is Modulated at 3000Hz. White Lamp will illuminate in the Cockpit with six dots per second. The Marker Beacon System is consists of the following components. 1) Two light assemblers 2) Sensitivity Selector Switch 3) Receiver 4) Antenna: Located on the bottom of airplane.

INSTRUMENT SHOP

AUTO PILOTPITCH CONTROL CHANNELThe pitch control channel consists of ATR rack consisting of four plugs in modules, a rack gyro or derived rate assembly. A rugged self test meter, rotatory self test switch, gyro test toggle switch, handle, two hold down hooks and test point are mounted at the front of the rack. The pitch rack is the base of the pitch control channel. The pitch rack supplies the necessary support and the interconnection of the PC channel components and the sub assemblies. The front panel of the rack consists of the self test switch and meter, gyro test switch, hold down hooks and test points. The rear of the rack consists of dual electrical connector for the interconnections and the pitch control channel with the airplane wiring. Pitch calibrator is mounted at the rear of the pitch rack. The pitch calibrator is a plug in the module containing fixed resister

to the tailor gain with in the pitch control channel for the specific airplane configuration.

OPERATIONSP.C.C provides the pitch axis control axis control by amplifying, shaping and computing and compiling error and command signal to drive the elevator control surface through the hydraulic actuator. Synchronization mode: there are following functions of the pitch control channel. 1) pitch altitude synchronization 2) valve amplitude synchronization 3) elevator position synchronization The pitch control channel consists of following number of parts. 1) Pitch servo amplifier The pitch servo amplifier provides a quadrature rejection and power amplification of the pitch axis error and stabilization signal for commanding elevator movement. This unit also provides quadrature rejection. Signal spacing and steady state signal wipe out for the rate gyro signal and implement automatic stabilizer trim commands. The P.S.A is a tape sealed assembly that is rigidly attached to the pitch rack by the 4 captive. a 50 pin connecter in the bottom of the module mate with the corresponding sockets mounted in the pitch rack. The pair of the alignment pin on the bottom contains the component cards that are held in place by the pairs of Teflon rollers in the module frame and the card lid. The component card hinged at one corner to allow partial removal for maintenance it firmly place of the notch in the card lid, which engages a corresponding notch in the edge of the component card. Rate gyro signals are applied to amplitude de-modulator where they are amplified, demodulated and filtered. 2) Pitch Computer The P.C provides pitch control error signal to the servo amplifier. The unit consists of an electromechanical computer, which consists of the motor.

Amplifier driving the dual ratio gear reduction servo amplifier. The P.C is a tape sealed assembly that rigidly attached to the pitch rack by the 4 captive screws in the module bases. a 50 pin connecter in the bottom of the module mate with corresponding sockets in the pin rack. The pair of the alignment pins on the bottom contains the component cards that are held in the place by the pairs of the Teflon pairs of Teflon rollers in the module frame and card lid. The vertical gyro senses the pith amplitude. It supplies the signal to the servo assembly control transformer. If the transformer is not synchronized with the vertical gyro, a signal is generated in the control amplifier and sent to the pitch control amplifier. The vertical gyro signal is applied to the control transformer resulting CT output, which is routed to pitch servo amplifier to cause necessary in elevator position. Movement of the both the servo assembly and the elevator continue until the mechanical unit of the servo assembly is reached. When the CWS force is applied to control wheel is released or until the mechanical unit of servo assembly is clamped at the new reference position. Output from the control transformer continues, causing the airplane altitude changes until the vertical gyro output results in control transformer null output. Vertical path coupler glides slope receiver and radio altimeter signal to the AFCS. The module receives the signal from the glide slope receiver and processes this signal to command the signal to command the airplane to descend on the glide slope radio beam. The P.S.A is the tape sealed assembly that is rigidly attached to the pitch rack by 4 captive screws in the module base. A 50 pin connector in bottom of the module mate with the corresponding sockets mounted in the pitch rack. The pair of the alignment pins on the bottom contains the component cards that are held in the place by the pairs of the Teflon rollers in the module frame and the card lid. Vertical path coupler is to implement airspeed hold mode. The vertical beam sensor senses glide slope signal level. When the glide slope signal level decreases to a predetermined level, Vertical sensor changes state to provide an output to start 10sec timer. The glide slope signal does not pass through gain programmer until 10-sec timer activated easy on the switch. 3) Control Wheel Steering Coupler

The C.W.S.C is tape sealed assembly that is rigidly attached to the pitch rack be 4 captive screws in the module bases. Operation: The pair of the alignment pins on the bottom contains the component cards that are held in place by the pairs of the Teflon rollers in the module frame and the card lid. The C.W.S.C provide the quadrature rejection and level detection of the pitch axis control wheel force sensor signal to the pitch control channel, which responds to elevator movement commands. A/c glide slopes capture bias signal to be applied to summing point to command a predetermined rate of descent. An altitude rate signal from C.A.D.C is also applied to the summing point so that the input from summing point is filtered and amplified by the amplifier demodulator and amplifier movement commands ten seconds after the altitude signal begun passing through the switch, switch changes state, removing the altitude rate signal from the pitch computer.

ROLL CONTROL CHANNELPHYSICAL DESCRPTIONThe roll control channel contains the electronic and electromechanical assemblies and sub assemblies, which provides the signal processing, shaping and computing functions required for driving automatic flight control system. 2) This unit consists of half A.T.R rack containing four plugs in modules, a rate gyro or a derived rate assembly, and self test assembly. 3) Mounted on the front panel of the rack are: a rugged zed self test meter, rotator self test switch, gyro test toggle switch, handle, test points and hold down hooks. 4) Roll back assembly provides the mechanical support and interconnections for the roll control. 5) Roll calibrator is plug in module of the roll control channel, containing fixed resisters to adapt the gains of the roll control channel for a specific airplane configuration. The roll calibrator is mounted at the rear of the roll rack.1)

6) Roll rack assembly provides the mechanical support and interconnections for roll control. 7) Roll calibrator is a plug in module of the roll control channel containing fixed resisters to adopt the gains of the roll control channel for the specific airplane configuration. The roll calibrator is mounted at the rear of the roll rack.

FUNCTIONAL DESCRIPTIONThe roll control channel provides roll axis control by amplifying, shaping, computing and coupling error and command signals to drive an electro-hydraulic transfer valves in the airplane aileron power units. The roll control channel consists of the following four modules:-

1)

ROLL SERVO AMPLIFIER

PHYSICAL DESCRIPTION: The roll servo amplifier is a tape sealed assembly that is rigidly attached to the roll rack by four captive screws in the module base. A 50 pin connector in the bottom of the module mates with the corresponding socket mounted in roll rack. The module contains components cards that are contained the place by the pairs of Teflon rollers in the module frame and card lid. FUNCTIONAL OPERATION: The roll servo amplifier shapes and amplifies aileron command and control wheel steering signals. It detects the C.W.S signal level and indicates, by the logic output, when control wheel is out of the detent. The rate gyro signals are applied to the amplifier demodulator where they are amplified, demodulated and filtered.

2)

ROLL COMPUTER

PHYSICAL DESCRIPTION: The roll computer is a tape sealed assembly that is rigidly attached to the roll rack by four captive screws in the module base. A 50 pin connector in the bottom of the module mates with the corresponding socket mounted in roll rack. The module contains components cards that are contained the place by the pairs of Teflon rollers in the module frame and card lid.

FUNCTIONAL DESCRIPTION: The roll computer generates the aileron commands signals proportional to difference between the desired roll attitude and actual roll attitude. It shapes, limits , amplifies and integrates roll axis error and command signals. The roll attitude as senses by the vertical gyro is applied to the stator of the control transformer rotor position, when compared with vertical gyro input, results in the attitude error signal. This signal ultimately actuates the aileron in the direction to eliminate the original error, i.e. heading error. The resolver generates the feedback signal proportional to its displacement from the zero or the wings level position. This resolver feedback cancels the error signal when proper amount of control transformer rotation is accomplished. The control transformer then stops driving. As bank angle changes, the vertical gyro input changes accordingly; in effect following the control transformer. When vertical gyro input matches the control transformer position, the attitude error is zero and a bank angle is established.

3)

LATERAL PATH COUPLER

PHYSICAL DESCRIPTION: The lateral path coupler is a tape sealed assembly that is rigidly attached to the roll rack by four captive screws in the module base. A 50 pin connector in the bottom of the module mates with the corresponding socket mounted in roll rack. The module contains components cards that are contained the place by the pairs of Teflon rollers in the module frame and card lid. FUNCTIONAL DESCRIPTION: The lateral path coupler shapes, amplifies and limits the beam deviation and error signals. It also senses signal level and its rate of change. The lateral path coupler has no effect on the roll channel operation during modes other then the localizer and V.O.R. the timer functions as a part of the roll logic circuitry. When it receives the logic signal for heading hold it applies the one second ground to logic circuitry. This action prevents entry into heading hold mode for the one second. During V.O.R intercept radio beam deviation signals from the V.O.R receivers are applied to lateral beam sensor. The L.B.S

changes state when the amplitude of the modulated radio beam deviation drops below the preset level. For V.O.R mode, the level is established at 30 mill volts by the action by the energized switch. When the L.B.S changes the state, a signal is fed to the latching circuit in logic circuitry, which terminates intercept and initiates the capture.4)

HEADING SYNCHRONIZER:

PHYSICAL DESCRIPTION: The heading synchronizer is that is rigidly attached to the roll rack by four captive screws in the module base. A 50 pin connector in the bottom of the module mates with the corresponding socket mounted in roll rack. The module contains components cards that are contained the place by the pairs of Teflon rollers in the module frame and card lid. FUNCTIONAL OPERATION: The heading synchronizer generates error signals proportional to the difference between actual heading and desired heading. It also shapes and amplifies heading select signals and provides beam integration. The airplane heading as sensed by the directional gyro is applied to the stator of the control transformer. It control transformer rotor is not aligned with the electromagnetic field created by the stator; a control transformer output is generated. This heading error signal is applied through summing point to the motor amplifier. Generator feedback is also applied to the summing point. This feedback signal is proportional to the rate of the motor- tachometer generator. The control transformer rotor is driven in the direction to eliminate the original error. When control transformer rotor is aligned with stator, heading error signal is eliminated and heading synchronizer has effectively stored the airplane heading.

YAW DAMPER COUPLER (YAW CHANNEL)DESCRIPTION1) Yaw damper coupler provides the sensing, interlocking, controlling and computing functions required for yaw damping. 2) It consists of rack, rate gyro, yaw computer plug- in module and yaw calibrator plug- in module.

3) Yaw rack is the base of all the units and provides necessary support and interconnections for the remaining modules. The front of the rack contains a self test switch; self test meter, gyro test switch, test point and rack fastening handles. The rear of the rack contains a connector for the interconnections with airplane wiring. The yaw rack houses position sensor, rate gyro excitation supplies, 30 V dc supply, module excitation and power sources and interlock circuitry. 4) The yaw rate gyro is miniature rate gyroscope. It is attached to the yaw rack with four captive screws. It is protected by the dust cover to yaw rack by the two captive screws. 5) The yaw computer is sealed assembly rigidly attached to the yaw rack by the four captive screws at its base. Connections between the yaw computer and yaw rack are made through a 50 pin connector in the module base. It contains three cards, which are hinged at one corner and held in the place by the pairs of Teflon rollers n module frame and card lid. 6) Calibrator module is rigidly attached to the yaw rack by two captive screws at its base. Connections between the two are made through a connector in the module base. It contains the one resister card, which is held in the place by the four captive screws.

FLIGHT DATA RECORDERIt is capable of storing about 350 parameter when aircraft is flying. It is of two types: SSFDR Solid state flight data recorder. Solid state box. Present in Boeing DFDR Digital flight data recorder Endless magnet tape Advanced from present in A-320

DFDRIt is an endless loop magnetic tape device, which works either alone or as a portion of the flight data acquisition recorder system. When installed in an aircraft the DFDR recording operating information consecutively on 6 tracks for maximum of 25 hours. All data is read out 3 seconds after recorder and is fed to BITE continuously monitors such things as the incoming data and power faulty track changes and internal operation of DFDR. Displays in the aircraft flight data entry panel indicated the operating status of DFDR as determined by the BITE. DESCRIPTION Crash and fire protection capable: The tape transport capsule is triple protected by an inner AI easing an isothermal protection shield, a stainless steel crash casing cover and the external stainless steel dust cover. 2) Drive unit assembly: consists of 2 phase, 115V, and 400 Hz induction motor drawing approximately 5 watts. 3) Reel and tape assembly: it consists of precision bridge assembly containing the record and reproduced. The pitch roll assembly, which controls the rate of movement of the tape and tape path of the channel assembly, which guides the tape into the tape hub and reel. 4) Tape configuration: the tape is and endless loop of inch, it is approximately 0.0012 inch thick and is reeled as inter periphery of the bundle, after having passed over write and read head. A long window on the tape allows a light signal from the LED to pass from the optical track change sensor to switch the recording from one track to the next. 5) Magnetic heads: DFDR uses 2 six track heads, the write head, which puts the data into the tape and read head, which is used for self monitoring as well as data reproduction by a read data unit (RDU). Since the actual recording on the tape is non return to zero mode the write head automatically erase and any 25 hours old data as it records new incoming eliminating the need for separate need for erasing.1)

OPERATIONIt consists of various units, which work simultaneously so as to bring out the desired operation. The various units are:-

1) Synchro/ mix upto six analog signals can be inputted to a single synchro card with one or two boards being used depending upon the number of inputs. These signals are resolved into the sine and cosine components. 2) Synchro/digital conversion, The resulting sine and cosine analog signal output from the mux is then routed t