Aircrew Electrical System (PDF)

RESTRICTED 11 CHAPTER-1 INTRODUCTION AND LAYOUT OF COCKPIT Introduction 1. Dornier DO-228 aircraft is provided with both DC & AC power supplies. 24V supply is provided by two nickel cadmium a...

RESTRICTED 11 CHAPTER-1 INTRODUCTION AND LAYOUT OF COCKPIT Introduction 1. Dornier DO-228 aircraft is provided with both DC & AC power supplies. 24V supply is provided by two nickel cadmium alkaline batteries and 28V is generated by two starter/generators. 115V & 26V AC of 400 Hz are provided by two static invertors. 2. Major systems of aircraft are operated by 28V DC supply. Most of the avionic and few instrument systems are operated by AC supply of 26V and 115V. 3. 28V DC external supply can be connected to aircraft for servicing and ground starting. Layout of Cockpit 4. Different panels laid out in cockpit are as under:- (a) Main Instrument Panel (b) Centre Pedestal Panel (c) Overhead Switch Panel (LH)- 5VE (d) Overhead CB Panel- 5VE (e) Overhead Switch Panel (RH)- 6VE (f) AC/DC CB Panel – 13VE 5. The main Instrument Panel houses various instruments for (a) Flight and Engine Monitoring (b) Navigation (c) Landing Gear Operation and Monitoring (d) Central Warning System (e) Environmental Control Systems RESTRICTED  RESTRICTED 12 Fig. 1-1 Main Instrument Panel Fig. 1-2 Central Pedestal Panel Fig. 1-3 Control Panel (4VE) RESTRICTED  RESTRICTED 13 6. The centre pedestal panel houses control and indicating equipment for flight controls landing gear, radio and navigation systems. Switches for landing gear position and warning, nose wheel steering, environmental system, interphone and start and ignition are located on the underside of the panel. These are operated by the engine POWER and SPEED levers. 7. Control panel 4 VE is located in central pedestal panel. It houses various switches and indicators for nose wheel steering, hydraulic, aileron trim and propeller control system. A test button is fitted for checking all warning and indicating lights. Overhead Switch Panel LH 5 VE 8. The overhead panel 5 VE is divided in to the two parts. The upper part contains bus bars and various system circuit breakers of ESS Bus and Bus 1. The lower part contains switches and electrical/Mechanical indicators of engine starting, ignition and fuel system. A lighting panel is fitted, the brightness of which is variable. The various switches fitted on this panel are shown in Figure. Fig. 1-4 Overhead Switch Panel LH 5 VE RESTRICTED  RESTRICTED 14 Overhead CB Panel LH 5 VE Fig. 1-5 Overhead CB Panel LH 5 VE Overhead Panel 6 VE 9. This overhead panel contains all controls and supervision units of the DC/AC power system. In addition the panel contains external lights control switches, instrument lights dimming controls and wind shield wipers control switch. A lighting panel is fitted, the brightness of which can be varied. The switches and meters fitted in this panel are shown. RESTRICTED  RESTRICTED 15 Fig. 1-6 Overhead Panel 6 VE Fig. 1-7 Overhead Panel RESTRICTED  RESTRICTED 16 AC/DC Circuit Breakers Panel (13 VE) 10. It is located on the rear bulk head behind the pilot seat. It houses bus bars AC 115 V, 26 V AC, Bus 1, Bus 2 and Non Ess Bus Fig. 1-8 AC/DC Circuit breakers Panel (13 VE) Bibliography: Airplane Maintenance Manual Vol- V Chapter- 39 ********** RESTRICTED  RESTRICTED 17 CHAPTER- 2 ELECTRICAL POWER SUPPLY SYSTEMS General Description 1. The DO-228 aircraft electrical power supply system consists of following power supply sources:- (a) The primary power supply to aircraft is provided by two 28V DC starter generators feeding parallel Busbars. Each generator is capable of supplying the full load requirement of the aircraft in emergency. (b) Two 24V 27AH nickel cadmium batteries are installed one each on either side of the nose section. Normally both batteries are connected to BATT BUS in parallel. (c) During engine start with internal power both these batteries are automatically connected in series to provide 48V to BATT BUS only. (d) An external DC power receptacle is installed in the front fuselage where an external DC power can be connected to the aircraft. (e)Two static invertors are used to provide single phase 26V and 115V AC at 400 Hz. DC Power System Starter Generator 2. A starter generator is mounted on the accessory gearbox of each engine. The generators have a nominal regulated output of 28V DC, 250 amps each. Following engine start, the generator delivers electrical power when GEN. LH and GEN. RH switches are set to ON. 3. The generators are protected against the over voltage, over excitation and reverse current or differential current by voltage regulators, which maintain the output within 28 ± 0.5V throughout the normal speed/load range of aircraft operation. Technical Data (a) As a Generator (Also known as Generator Limitation) Output Voltage : 28V DC Current : 250 Amps (Continuous) Current : 375 Amps (2 mins) Current : 500 Amps (3 Secs) RESTRICTED  RESTRICTED 18 (b) As a Starter Input Voltage : 28- 30 VDC Current : 860 Amps Current : 1000 Amps (Max) Starter limitations (i) First Cycle : Max. 1 minute ON - Min. 1 minute OFF (ii) Second Cycle: Max. 1 minute ON - Min. 5 minutes OFF (iii) Third Cycle : Max. 1 minute ON - Min. 1 hour OFF Internal Batteries 4. Two 24V, 27AH nickel cadmium batteries are installed in nose section. The batteries supply power to all electrical systems except the Non-ESS BUS, when either battery switch and master switch is set to ON. 5. Both batteries are protected against overheating (71°C) by thermal switch connected to central warning system. When battery overheat warning lights up, the respective battery should be switched OFF immediately. Technical Data (i) Nominal Voltage : 24 V DC (ii) Electrolyte : KOH (iii) Thermostat Assembly : -2 Thermostat Operating at 71°C (iv) Operating Temperature : +50° C to -30° C (v) Saft Battery Type : 2778-2 External Power 6. An external DC power receptacle is installed in the front fuselage where an external DC power unit is connected. Maximum output of the external power unit must not exceed 1000 amps at 28V. The white EXT power light will illuminate as the ground power unit (GPU) supply is made available. The GPU supplies to the electrical system, when both battery switches are set to EXT and master switch set to ON. DC Power Control and Indication Master Switch 7. The MASTER switch is located on the RH overhead S/W panel. It has two positions ON and OFF. In the ON position the switch connects all electrical power sources to the buses. The main function of the S/W is to disconnect the electrical buses from the electrical power source in case of emergency. Note: Both BATT switches must be in desired position (ON or EXT) before the master S/W is set to ON. RESTRICTED  RESTRICTED 19 BATT-1/BATT-2 Switches 8. A BATT-1 &BATT-2 Switches are located on the RH overhead switch panel. The switch has the following positions:- (a) ON: The batteries are connected to the DC electrical buses. External power, if connected, will be isolated. (b) OFF Internal and external power will be disconnected. (c) EXT: External power is supplied to electrical buses. Note: Both BATT switches must be in EXT position to avail external power. TIE Switch 9. A bus TIE S/W is located on the RH overhead panel. The switch has the following positions:- (a) TIE (Centre):. In this guarded position both BUS-1 & BUS-2 are interconnected. Both batteries are charged by GEN-1 and Gen- 2. (b) BATT BUS-1 (Left): BUS-1 & BUS-2 are separated. GEN LH and GEN RH power the BUS-1 and BUS-2 respectively. Both batteries are connected to BUS- 1. (c) BATT BUS-2 (RH): BUS-1 and BUS-2 are separated. GEN LH and GEN RH power the BUS-1 and BUS-2 respectively. Both batteries are connected to BUS-2. Note: (I). With GPU power applied to the airplane, the tie S/W is enabled. (ii). When switching to the left (BATT BUS-1) and the DIM S/W to the position DAY, (while operating on battery power), the warning lights will DIM automatically. GEN LH/GEN RH Switches 10. Two switches labeled GEN LH and GEN RH are located on the RH overhead S/W panel, which is a three position switch with ON, OFF and RESET positions. 11. The ON position connects the generator to electrical system through voltage regulator. During over voltage or under voltage, the generator field circuit is de-energized and the GEN is disconnected from the system. To reconnect the generator, the corresponding, GEN S/W must be held at RESET for 5 seconds (which energizes the generator field circuit) and then engage the S/W to ON position. RESTRICTED  RESTRICTED 20 NON ESS BUS Switch 12. The guarded NON-ESS BUS S/W is located on the RH overhead switch panel. It is used to disconnect NON-ESS BUS to reduce the load from power source, if only limited electrical power is available (GEN or battery malfunction). The S/W is safety guarded in the ON position. 13. The NON-ESS bus is protected by NON ESS BUS CB of ESS BUS on the overhead CB panel. Voltmeter Selector Switch 14. The voltmeter selector S/W is spring loaded and is located on the RH overhead S/W panel. At the centre position, it connects the ESS BUS to indicate its voltage or the output of respective generator, if held to the GEN LH or GEN RH position. DC Voltmeter 15. A DC voltmeter is located in RH overhead S/W panel above the voltmeter selector S/W. (a). The voltmeter is graduated from 0 to 32 volts. (b) The green range is between 26 to 28.5 volts. DC Ammeters 16. A DC ammeter for each generator is located on the RH overhead switch panel to indicate electrical load of the generator. The ammeter is graduated from 0 to 500 Amps. 17. The green range is between 0 to 250 Amps. The yellow range is between 250 to 500 Amps. AC Power System 18. AC electrical system consists of two 1000 VA static inverters which are powered through BUS-1 and BUS-2. Each inverter provides 26V & 115V AC, 400Hz single phase to its own BUS system. If one of the inverters become inoperative, the load will be automatically transferred to the other inverter. The system is protected by:- (a) INV-1/BUS-on the 1 VE Electrical Control Panel. (b) INV-1 CONT/BUS-1 on the Overhead CB Panel. (c) INV-2/BUS-2 on the 1 VE Electrical Control Panel. (d). INV-2 CONT/BUS-2 on the AC/DC CB Panel. RESTRICTED  RESTRICTED 21 1200VA Static Inverter 19. The system consists of a 1200 VA Static inverter which converts 28 V DC into 230 V AC, 50 Hz, and 1200 VA. Protection device CB is fitted in 21 VE panel. Technical Data:- Input: 28V DC, Max 54A O/P: 230V AC ± 3% Frequency: 50Hz ± 1% AC Electrical Controls and Indicators INV-1/INV-2 Switches 20. Two inverter switches are located on the RH overhead switch panel and labeled ON/OFF/RESET. They are used to switch the invertors ON & OFF. The RESET function is not used. AC Voltmeter 21. Two AC voltmeters labeled INV-1 and INV-2 are located on the RH overhead switch panel. They are graduated from 0 to130V and are used to measure inverter output. The green range is between 110 to 120 Volts. Electrical Load Distribution System 22. The electrical load distribution system consists of Busbars and protection devices for safe guarding the various consumer systems. Circuit Breaker Panels (Protection Devices) 23. It consists of two panels on which various circuit breakers (CB) are mounted. These CB's act as protection devices for various systems against overload. These panels are:- (a) Overhead CB panel This is divided into two parts. The upper portion has got CB's for various loads of Busbar-1. The lower portion contains CB's for various loads of essential Busbar. (b) AC/DC CB Panel It is located in the LH avionic rack, having six rows of CB's on it. The first two rows pertain to the CB's of 115 and 26 volt AC systems. The next three rows pertain to the CB’s of DC system. The last row pertains to the CB's of NON-ESS loads. Bus bars 24. (a) Battery Busbar The battery Busbar supply is provided by Battery No 1 and No 2 or GPU. Both starter generators are connected to Battery Busbar. During engine starting, using internal batteries this Busbar is supplied by 48V till the completion of starting sequence. RESTRICTED  RESTRICTED 22 (b) Hot Busbars LH/RH These Busbars are directly connected to the battery. Loads connected to these Busbars cannot be switched off. (c) Busbar-1 and Busbar-2 These bus bars are supplied by the generator LH and RH respectively. (d) Busbar-3 and Busbar-4 These Busbars are the auxiliary Busbars of No 1 Busbar. (e) Busbar-5 and Busbar-6 These bus bars are the auxiliary bus bars of No 2 bus bar. (f) Essential Bus bar This bus bar is located in the overhead circuit breaker panel and will be supplied whenever any power supply is switched ON to the aircraft. Location & Source of Supply for Different Busbars D C BUS BAR LOCATION MAIN SOURCE OF SAMPLES 1 PP 1 VE Heart Panel LH Generator 2 PP 1 VE Heart Panel RH Generator 3 PP 5 VE Over Head Panel 1 PP 4 PP 13 VE AC/DC CB Panel 1 PP 5 PP 13 VE AC/DC CB Panel 1 PP 6 PP (NON - ESS) 13 VE AC/DC CB Panel L 2 PP 7 PP (ESS) 5 VE OVER HEAD Panel ALL 8 PP (LH HOT) LH Batt Compartment LH Battery 9 PP (RH HOT) RH BATT Compartment RH Battery BAT BUS 1 VE Heart Panel Both/Any Batt Note: 1. During Internal Start Only Batt. Bus Bar Will Get 48 Volts. 2. 8 PP & 9 PP Can Not Be Connected With External Power. 3. 6 PP Can Not Be Connected With Internal Batteries. Important Note: In case of double generator failure, the batteries can supply some of the loads and sustain for a duration of not less than 30 minutes provided the following circuits are switched off within 10 minutes of the failure occurring:- Standard Options (if installed) COCKP LT Switch OFF COCKP Switch OFF CABIN LT Switch OFF ADF 2 OFF ANTICOLL Switch OFF Weather Radar OFF Picot (RH) Switch OFF HF COM OFF VHF COM 2 OFF OMEGA OFF VHF NAV 2 OFF Radio Altimeter OFF FDS/AP OFF Transponder 2 OFF DME 1 OFF AIRCOND Sel Switch RAM AIR Bibliography: Airplane Maintenance Manual Vol. II Chapter- 24 RESTRICTED  RESTRICTED 23 CHAPTER - 3 CENTRAL WARNING SYSTEM Fig. 3-1 Central Annunciator Panel 1. The annunciator system consists of the central annunciator panel located in the centre of the main instrument panel, an annunciator dimming switch labeled DIM with the position DAY/NIGHT, a press-to-test button labeled LAMP TEST, two red master warning indicators labeled WARNING, two amber master caution indicators labeled CAUTION, and various indicator lights as described in the various systems. 2. The master warning light and the master caution light are located on the upper left and right extremes of the instrument panel. ** The warning light DOORS illuminates with red letters when lit. 3. The caution and warning panel consists of five horizontal rows of seven indicator lights, forming a display of thirty five individual indicators. Five indicators in the upper row (warning lights) and in the second row “DOORS” have red filters. The other twenty nine indicators (caution lights) have amber filters. All indicators illuminate with red or amber letters on a black background. The indicators are bright enough to be read in day light when energized. 4. If a fault occurs which requires immediate corrective action the respective red warning light and the master WARNING light illuminate, actuating an audio warning of 1 Kz which can be heard outside as well as in the pilot head phones also. The master WARNING LIGHT, the warning horn and the 1 KHz tone can be cancelled and the system reset by depressing the master WARNING light. RESTRICTED  RESTRICTED 24 5. If a system is inoperative and/or requires attention but no immediate action, the appropriate amber caution light and the master CAUTION light illuminate. The master CAUTION light can be cancelled and the system reset by depressing the master CAUTION light. 6. If a fault is on or cannot be corrected the corresponding WARNING or CAUTION light in the annunciator panel will remain illuminated. 7. For action to be taken upon WARNING or CAUTION light illumination, refer to Section 3 of pilots operating hand book. 8. A LAMP TEST button is installed on the pedestal panel. When the LAMP TEST button, is pressed all the lamp and annunciator indicators are illuminated with simultaneous audio warning. 9. The DIM switch located on the overhead switch panel (LH) may be used to change the annunciator lighting intensity between DAY (bright) to NIGHT (dim) except for the red master WARNING lights. 10. The system is supplied with 28V DC and protected by:- - WARN PANEL 1/ESS BUS on the overhead circuit breaker panel. - WARN PANEL 2/BUS 1 on the overhead circuit breaker panel. Acoustical Warnings 11. Several different acoustical warnings are triggered in conjunction with either the Master Warning Circuit or the respective systems, allowing to distinguish between different systems in case of a annunciator light malfunction or burned out light bulbs. Frequency Type Can be muted or not Landing Gear 450 Hz Cycling NO (if flaps fully extended) Stall Warning 1000 Hz Continuous No Master Warning 1000 Hz Continuous Yes Fire 1000 Hz Continuous Yes VMO 1000 Hz Continuous Yes Autopilot Disconnect Alert 2800 Hz Continuous Yes Door Warning System 16. The door warning system indicates when the cabin or baggage door is not closed and locked. The door warning system consists of micro switches mounted on the fuselage door frames and connected in parallel. The micro switches are normally open when the doors are closed. When any of the door is not closed and locked, ground is connected to an annunciator DOOR WARN illuminates. RESTRICTED  RESTRICTED 25 17. When the rear baggage compartment door is not closed and locked ground is connected to the baggage compartment lighting system via diode DI48. Diode 147 isolates door warning system from the baggage compartment lighting. Fig. 3-2 Circuit Diagram of Door Warning System & Location Doors Microswitches Bibliography: Airplane Maintenance Manual Vol. III, Chapter-31 ********** RESTRICTED RESTRICTED 26 RESTRICTED  RESTRICTED 27 CHAPTER- 4 AIRCRAFT LIGHTING SYSTEM External Lights 1. The Airplane is equipped with the following exterior lights. (a) Three navigation lights. (b) Two landing lights. (c) Two taxi lights (d) Two anti-collision lights. Navigation lights 2. The navigation lights are conventional, consisting of a red light on the left wing, a green light on the right wing and a clear light on the tail cone below the rudder. The navigation lights are controlled by the NAV switch on the RH overhead switch panel. 28 volts DC electrical power is supplied from Bus1 and protected by a CB marked NAV LTS on the AC/DC CB panel. Landing and Taxi / lights 3. Landing and Taxi light one each are provided in each main landing gear housing. The lights are controlled by a three position toggle switch. It has three positions as follows:- (a) Land - Landing lights are energised (b) Off -. Lights are de-energised (c) Taxi - Taxi lights are energised 4. The system receives 28 volts DC electrical power from:- (a) ‘LAND LTS RH’ of BUS-1 on the AC/DC CB Panel. (b) ‘LAND LTS LH’ of NON-ESS BUS on the AC/DC CB Panel (c) ‘TAXI RH’ of NON ESS BUS on the AC/DC CB Panel. (d) ‘TAXI LH’ of NON ESS BUS on the AC/DC CB Panel. Note: In case of both generators fails only RH LAND LTS can be switched ON. RESTRICTED  RESTRICTED 28 Anti-collision Lights 5. Two red anti-collision lights are installed, one on top and one at the bottom of the centre section of fuselage. The lights are controlled by the two position toggle switch marked ANTICOLL on the RH overhead switch panel. The circuit is powered by 28V DC from bus 2 and protected by a CB marked ANTI-COLL on AC/DC CB panel. Internal Lights 6. The airplane is equipped with the following internal lights:- (a) Instrument Lights. (b) Panel Lights. (c) Cabin Lights. (d) Cabin Signs. (e) Cockpit Dome Light. (f) Toilet Light (g) Baggage Compartment Light. Instrument Lights 7. A row of eight light segments is installed in the glare-shield to illuminate the instruments. The ADI, HSI and the RMI are internally illuminated. 8. The intensity of the four light segments on the pilot's side and the internal instrument lighting are controlled by a rheostat labeled INSTR on the LH overhead switch panel. 9. The intensity of the other four light segments on the co-pilot side and the internal instrument lighting are controlled by the INST rheostat on the RH overhead panel. 10. The lights are powered by 28V DC from Bus1 and protected by CB marked ‘INSTR LTS’ on the AC/DC CB panel. Panel Light 11. All switches, circuit breakers and control-panels, including the avionics are illuminated internally. The centre pedestal panel is additionally illuminated by the dome light on the cockpit ceiling. The intensity of panel lights is controlled by the outer (large) knob of the rheostat labeled ‘PANEL’, situated on the RH overhead switch panel. The lights are powered by 28V DC from Bus1 and protected by a CB marked ‘INSTR LTS’ on the AC/DC panel. RESTRICTED  RESTRICTED 29 Cabin Light 12. The system consists of four lighting units on the ceiling panels, each unit being equipped with an 8.5 watts and an 18.5 watts lamps. Cabin lights are controlled by a three position toggle switch with marked positions FULL, OFF and HALF. 13. The system is supplied with 28V DC from Bus1 and protected by a CB marked ‘CABIN LT’. When the cabin light switch is set to ‘FULL’, both the lamps in each lighting unit are illuminated. With the switch set to ‘HALF’, only the 8.5 watts lamp in each lighting unit is illuminated. Optional 14. Three lights installed in the cabin ceiling are provided for cabin lighting. The three position toggle switch on the RH overhead switch panel labeled CABIN has the positions:- (a) Off. Lights are de-energised. (b) Half. Two lights illuminate the cabin. (c) Full. Three lights illuminate the cabin. 15. Two of the lights are powered by 28V DC and protected by:- (a) CABIN LTS/BUS-1 on the AC/DC circuit breaker panel. The other one is protected by CABIN LTS/NON-ESS BUS on the AC/DC CB panel. Cabin Signs 16. The two toggle switches in the LH overhead switch panel labeled as SEAT BELT and NO SMOKE when operated illuminate the signs "Fasten Belts" and "No smoking" respectively. A 28V DC electrical power is provided from Bus-2 and protected by CB marked ‘SEAT-B & N-SMOKE’ on the AC/DC CB panel. Cockpit Dome Lights and Toilet Light 17. The dome light installed in the cockpit ceiling is provided for cockpit lighting. The light contains a normal bulb and an emergency bulb. It is controlled by the three position toggle switch labeled COCKP which has the positions:- (a) OFF: Light is de-energised. (b) ON: The normal bulb is supplied from Non-ess bus and protected by CB marked ‘DOME LTS’ on the AC/DC CB panel. (c) EMERG: The emergency bulb is supplied from Bus-2 and protected by CB marked ‘CONSOLE PANEL LTS’ on the AC/DC CB panel. RESTRICTED  RESTRICTED 30 Toilet Light 18. The toilet light will always be illuminated with DC power connected to the airplane electrical system. The light is supplied from Non-ESS BUS and protected by CB marked ‘TOILET’ on the AC/DC CB panel. Baggage Compartment Light 19. The rear baggage compartment is illuminated by a single lamp. The light is controlled by a switch installed on the lamp and/or a contact on the baggage compartment door. Map Reading Lights 20. The map reading lights are fitted on to the pilots and co-pilots control wheel. 21. They have red inserts and the light aperture can be varied. The lights are switched ON and OFF with the toggle switch located on the control wheel below the map reading light. 22. Map reading lights are supplied from Non-ESS BUS and protected by CB marked ‘CABIN LT’. Bibliography: Airplane Maintenance Manual Vol IV Chapter-33 ********** RESTRICTED  RESTRICTED 31 CHAPTER -5 FIRE DETECTION AND EXTINGUISHING SYSTEM Fig. 5-1 Fire Detection and Warning System Fire Detection and Warning System 1. The system provides audible and visual warning of fire or abnormal temperature rises in the engine bay. The system consists of three main components namely control units (LH/RH), fire detector elements (LH/RH), fire warning lamps (LH/RH). Control units 2. LH/RH control units are located in the fuselage nose and are the monitoring unit for LH/ RH detection and warning system. The units consists of transistor oscillator, transformer, test circuits with relays and alarm circuit with relay. Fire Detector Element 3. The fire detector elements are flexible elements whose resistances and capacitances vary with ambient temperature. They are fitted on the left and right cowling doors on each engine in such a way that forms a continuous detector loop for their respective control units. Each element consists of a thin stainless steel capillary enclosing a centre electrode which is coaxially located by semiconductor material with a negative temperature coefficient of resistance. RESTRICTED  RESTRICTED 32 4. Each engine compartment is equipped with a fire detection system. The system consists of temperature sensors, a control unit and a FIRE EXT TEST LH &RH warning light test circuit. The system is powered by 28V DC and protected by:- (a) FIRE - WARN LH/ESS BUS on the Overhead CB Panel. (b) FIRE - WARN RH/ESS BUS on the Overhead CB Panel. 5. If excessive heat develops in an engine nacelle, the temperature sensors energize a relay, causing the appropriate RED FIRE warning light to illuminate and triggering a master audio warning. The master warning can be cancelled by depressing the master warning light. If the temperature decreases the FIRE warning light will extinguish and the system is rearmed automatically. When the FIRE EXT TEST LH (RH) light are pressed, the systems are tested and if the test is satisfactory, the FIRE EXT TEST LH (RH) lights and the FIRE warning lights in the FIRE PULL handle will illuminate and a master warning is activated. Self test 6. The system can be tested by pressing the appropriate FIRE EXT TEST switch. This simulate the overheat condition in the detector loop. If the warning light illuminates when the switch is pressed the followings are indicated:- (a) The detector loop is continuous. (b) The insulation resistance of the loop is within limits. (c) The control unit is serviceable. Technical Data 7. (a) Control unit Current Consumption at 30 V - Max 85 mA (Quiescent) - Max 175 mA (Operate) - Max 240 mA(Short Circuits) Relay Contact Rating 2A Switching Point Temp - 195oC (383oF) (b) Fire Detector Element Resistance 0.2 to 0.3 Ohms/foot below 50°C Insulation Resistance - Min 5 Mega Ohm at 250 V Minimum Bending Radius - 25 mm RESTRICTED  RESTRICTED 33 Fire Extinguishing System 8. The fire extinguishing system consists of two independent bottles, a squib assembly, discharge pipes and the operating handle. Each bottle is discharged into its associated engine. The bottles are charged with HALON 1301. 9. The engine fire extinguishing system is actuated by the FIRE PULL handles located on the main instrument panel. Pulling and turning the handle will close the fire cock, detonates the squib assembly and the extinguishing agent is ducted to the engine heat zone and the engine accessory compartment. The system is powered by 28V DC and protected by:- (a) FIRE LH/ LH BATT HOT BUS in the LH battery compartment. (b) FIRE RH/ LH BATT HOT BUS in the LH battery compartment. 10. Operating a fire handle will also supply 28V DC to the close winding of the LP cock to shutoff fuel supply to the engine. Caution: SINCE THE FIRE EXTINGUISHING SYSTEMS OF BOTH THE ENGINES ARE POWERED FROM HOT BUSBAR OF LH BATTARY. ACTIVATION OF THE FIRE HANDLES WILL OPERATE THE RESPECTIVE FIRE BOTTLES, EVEN IF THE ELECTRICAL MASTER SWITCH AND BATTERY SWITCHES ARE IN OFF POSITION IN THE COCKPIT. Fig. 5-2 Fire Extinguisher Bottle, Cartridge, Pressure Guage and Green Indicator RESTRICTED  RESTRICTED 34 11. The serviceability of the firing cartridge can be tested by pressing the FIRE EXT TEST switch mounted below the firing handle. When the switch is pressed, a reduced voltage is supplied to the firing cartridge. If the cartridge is electrically continuous, the green light in the switch will illuminate. 12. The fire detection circuit is also tested when the FIRE EXT TEST switch is pressed. This simulates an overheat condition in the fire detector loop and if the detection circuit is serviceable, the respective RED fire handle will come ON. Technical Data Fire extinguisher bottle Charge Pressure (Nominal) 450 PSI at 70° F Charge Pressure Effective - As Per Temperature Safety Relief Temp Range - 215° F to 226° F (101° C to 108° C) Type Of Agent - Halon 1301 (CBrF3) Weight - Lbs Total Weight - 2.66 Kg (5.86 Lbs) Cartridge Firing Voltage - 18 - 30 V DC Minimum Firing Current - 3.0 Amp Max Test Current - 50 mA Bibliography: Aircraft Maintenance Manual Chap.-26 ********** RESTRICTED  RESTRICTED 35 CHAPTER-6 HYDRAULIC SYSTEM Introduction 1. This system supplies hydraulic power for the following services:- (a) Landing Gear. (b) Wheel Brakes. (c) Nose Wheel Steering. 2. The system is operated by an electrically driven hydraulic pump. Fig. 6-1 Circuit Diagram of Hydraulic Power Pack RESTRICTED  RESTRICTED 36 Electrical Hydraulic Pump 3. The electrically driven pump is of a variable delivery, positive displacement type, driven by a direct current motor. The pump is an axial piston type and contains nine pistons with reciprocating compressors. Technical Data Max Hydraulic Pressure - 206 (Bars), 3000 PSI. Max Supply Rate - 8.25 lb / min Nominal RPM - 7500 RPM Operation Voltage - 28 V DC Max Current Consumption - 130 Amp (Motor P/N 259/1259A) 180 Amp (Motor P/N 1259A-1) (Power pack P/N 68501-0003.4) OEM Manufactured. (Power Pack P/N 115300000) FOR HAL Indigenous Note: Indigenous Hyd Power pack P/No. 115300000 (with pump-motor assy is interchangeable as an assembly with the imported Hyd Power Pack P/N 68501-0003.4(with pump-motor Assy). Pump / Motor assembly of Hyd Power Pack P/No. 115300000 are not to be interchanged with imported Hyd Power Pack P/N 68501- 0003.4. Control and Operation 4. The system is controlled by the "HYDRAULIC S/W" which has the following positions:- (a) NORM. Hydraulic pump operation is controlled by the LDG GEAR lever and landing gear micro switches with GPU or generator power. (b) MAN ON. Hydraulic pump operates on battery, generator or GPU power regardless of LDG GEAR lever and micro S/W position. (c) OFF. Hydraulic Pump is OFF. 5. The system is powered by 28V DC and the control circuits are protected by:- (i) HYDRAULIC NORM/ESS BUS on the overhead CB panel. (ii) HYDRAULIC MAN/ESS BUS on the overhead CB panel. 6. The hydraulic pump motor is powered by BUS-1 and protected by 150 Amps fuse. RESTRICTED  RESTRICTED 37 Operation Normal Operation 7. The system is controlled by the hydraulic switch which has three positions. The positions are Normal, Manual, and Off. Hydraulic pump operation is controlled by landing gear lever and landing gear "UPLOCK" micro switches with the GPU or generator power. When hydraulic pump switch is placed to the normal position with either generator on line or external supply is connected to the aircraft, it energises relay 3 PR which controls the supply to the hydraulic switch under normal position. 28 V DC is supplied via circuit breaker, contacts 5-6 of relay 3 PR, hydraulic switch in normal position, the three up lock switches and the landing gear lever switch to the coil of relay 4 DA. The relay 4 DA energises and connects hydraulic pump motor to the Bus 1 PP via 150 amps fuse. When landing gear lever is selected to up position the motor will continue to run till all the under carriages are locked up. 28V supply to the pump relay 4DA is interrupted by the landing gear up locks. Thus pumps automatically cuts off to avoid unnecessary running of the pump. Fig. 6-2 Circuit Diagram of Hydraulic Power Pack 8. When the landing gear lever is set to DOWN position the pump will come in to operation and continue to run till it is switched OFF manually by setting the hydraulic switch to OFF position or will be switched OFF automatically when last engine is switched OFF. 9. In 'Manual' position the hydraulic pump operation is not controlled by the landing gear lever and landing gear micro switches. It operates on battery, generator or GPU. RESTRICTED  RESTRICTED 38 Manual Operation 10. If the hydraulic pump does not run when the landing gear lever switch is selected down, the pump relay 4DA can be energised by setting the hydraulic switch to MAN ON position. If pump still does not run the hydraulic switch must be set to OFF enabling the landing gear to be extended using the emergency system. Operation with Other Systems 11. Relay 3 PR is energised from relay 6 PC or 16 PC when a generator is on line or from relay 7 PD when external electrical power is connected to the airplane. The nose wheel steering system is supplied with electrical power via the contacts of relay 6 DA which operates in parallel with the hydraulic pump relay 4 DA. Relay 5 DA is energised by the hydraulic pump supply circuit when the landing gear is retracting and applies a warning signal to the nose wheel steering system illuminating 'by pass' annunciator in the warning light panel. Note: 1. Normally one generator is able to supply to all the electrical buses. To ensure hydraulic pressure during landing, the HYDRAULIC SWITCH should be selected to MAN ON position. 2. During double generator failure use of hydraulic system is only possible with HYDRAULIC SWITCH in MAN ON and the "TIE" switch either in the TIE or BATT.BUS-1. Bibliography: - Airplane Maintenance Manual Chapter- 29 ****** RESTRICTED  RESTRICTED 39 CHAPTER-7 UNDER CARRIAGE SYSTEM 1. Extension and retraction of the landing gear is hydraulically operated and electrically controlled. The extension and retraction system is supplied from ESS- Bus.The nose wheel landing gear retracts forward into the nose section; the main gear retracts sideways into the bottom of the fuselage. The landing gear is controlled by the landing gear lever which is mechanically locked in down position when the aircraft is on the ground. The landing gear system consists of the following main components:- (a) Landing Gear Lever (2 GA) (b) Hydraulic Switch (10 GA) (c) Landing Gear Control Valve (d) LH/RH/NLG Weight Switches (e) NLG Centering Switch (f) LH/RH/NLG Actuators with Down Lock Micro Switches. (g) LH/RH/NLG Up Locks with Up Locks Micro Switches. Landing Gear Lever 2 GA 2. The landing gear lever is located in the LH side of the instrument panel. When the aircraft is on the ground, the landing gear lever is mechanically locked in the DN position by the spring operated plunger. With weight off the landing gear 28 V DC is supplied via weight switches and NLG centering switch to the solenoid. The solenoid energises and removes the mechanical interlock allowing operation of the lever. 3. In the emergency the mechanical interlock can be release by the DN LK REL button. 4. Three green lamps on the top of the lever housing indicates that the gear is locked down. A red lamp on the lever handle flashes indicates that the gear is unlocked. A test button is provided for testing the landing gear warning circuit and a mute horn button is provided to cancel it. RESTRICTED  RESTRICTED 40 Fig. 7-1 Landing Gear Lever Landing Gear Control and Operation 5. The landing gear is controlled by the LDG GEAR lever. This lever is mechanically locked in the DN position when the airplane is on the ground. When the airplane is in the air micro switches in the main landing gear strut, in the nose wheel strut and the NWS centering mechanism close a relay and allow the locking mechanism to be removed electrically. The LDG GEAR lever can now be put in the UP position to retract the landing gear. 6. It is possible to remove the mechanical locking mechanism by pushing the red DN LCK REL slider down, however the landing gear will not retract. 7. Three green lighted indicators labeled LEFT, NOSE, RIGHT, will illuminate when the landing gear is down and locked. 8. A light in the LDG GEAR lever flashes when the landing gear is not locked in the up or down position (intermediate position). 9. A warning horn will give an audio warning when one of the POWER lever is in the FI position and the landing gear is not locked in the down position. The warning can be muted by the MUTE HORN button to the right of the LDG GEAR lever. Subsequent movement of other POWER lever to FI will reactivate the warning circuit. This warning can be muted by the MUTE HORN button as above. However, when flaps are in position 2 (i.e. 20o), irrespective of the position of the power levers, the warning cannot be muted. 10. A TEST button below the MUTE HORN button, when pushed while the airplane is on the ground will activate the warning horn. 11. A MUTE HORN light will illuminate if the MUTE HORN button is pressed. RESTRICTED  RESTRICTED 41 12. 28V DC is provided and the system is protected by:- (a) LANDING GEAR CONTR/ESS BUS on the overhead CB panel. (b) LANDING GEAR VALVE/ESS BUS on the overhead CB panel. (c) LANDING GEAR WARN/ESS BUS on the overhead CB panel. 13. The three green LDG GEAR down lights are protected by the LANDING GEAR WARN CIRCUIT BREAKER. 14. The landing gear lever light is controlled by the INSTR rheostat and the MUTE light by the DIM S/W. Fig. 7-2 Circuit Diagram of Landing Gear System Nose Wheel Steering System 15. The aircraft is equipped with a hydraulically operated and electrically controlled nose wheel steering system, which permits nose wheel deflection of 8° or 45° according to rudder pedal deflection. This system is locked with hydraulic system. 16. The main components of system are feedback and command potentiometers, steering control unit and a steering actuator. The system is controlled by commands from the command potentiometer which is mechanically linked to the rudder pedals. This is a dual track potentiometer. One supplies servo signal and other track supplies the monitor signal to the steering control unit. Feedback potentiometer is mounted on the nose landing gear and is mechanically linked to the steering mechanism. It is also a dual track potentiometer supplying monitor and servo signals. RESTRICTED  RESTRICTED 42 17. The monitor signals from the feedback and command potentiometers are compared in the steering control unit which has three channels i.e. servo channel, monitor channel and control logic circuitry. The steering control unit has the following functions. Functions of Steering Control Unit (a) Reducing steering sensitivity around the nose wheel neutral position. (b) Limiting the steering rate to a maximum of 18° /Sec. (c) Selecting 45° or 8°steering authority as appropriate. (d) Driving the steering actuator to follow command signals from the rudder pedals. (e) Switching out the steering actuator in the event of failure. (f) Centering the nose wheels on takeoff prior to landing gear retraction. (g) Providing 0.5 sec delay for the system to come ON, after the nose wheel is loaded. NOSE WHEEL STEERING SYSTEM 28 V ,DC BUS 1 DTS/ELECT/AC/123 NWS -CB NWS BYPASS MICR-SWITCH SPEED-LEVER BUTTON S/W 45* PUSH 6DA RELAY NWS S/W Servo COMMAND STEERING STEERING Feedback POT-MTR CONTROL UNIT ACTUATOR Monitor POT-MTR CENTRING NLG SWITCH RUDDER WT. PEDALS S/W FEEDBACK SIGNAL Fig. 7-3 Nose Wheel Steering System RESTRICTED  RESTRICTED 43 18. The steering actuator is an electrically controlled hydraulic actuator mounted on the nose landing gear leg. A servo valve in the steering control unit, controls hydraulic fluid flow to the actuator cylinder. A solenoid valve which controls a bypass valve is controlled by monitor channel of the steering control unit. Operation 19. The system is supplied from Bus 1 (4PP) via relay 6 DA. When the rudder pedals are operated (left or right) two steering signals (required position) are applied from command potentiometer to the servo and monitor channel of steering control unit. Two actual position signals are fed from feedback potentiometer to the servo and monitor channel. 20. The error signals are compared in the servo channel amplified converted to regulated DC and fed to the steering actuator. The valve operates and applies hydraulic pressure to actuate the steering. As the nose wheel position changes the error signal reduces, the output voltage to the servo valve reduces proportionally and the rate of change of nose wheel position reduces. When the nose wheel reaches the required position the error signal is reduced to zero causing the servo valve to close. The nose wheel stops at the required position. 21. The position of the nose wheel in relation to the actual position from the rudder pedals is monitored by the monitor channel of the steering control unit. If an anomaly is detected, between the monitor signal from the command and feedback potentiometer the anomaly exceeds the preset threshold, steering control unit cut off signal to the solenoid valve. The solenoid valve is de-energized, port C is open to port R and the bypass valve is held in the bypass position by spring pressure. In this position the valve spool shuts off the system pressure to servo valve, the actuator is then in a hydraulically locked condition. The NWS 'bypass' annunciator illuminates and the nose wheels are free to caster. 22. If bypass valve opens due to a detected anomaly, the system remains in the bypass mode until it is reset by switching off the electrical power to the steering control unit and then switching on back again. 23. Failure to re-install the steering pin or electrical connections on the nose landing gear after towing the aircraft will result in NWS failure without caution/bypass light in the annunciator panel. 24. If the amber NWS bypass light is illuminated on ground, either the hydraulic system has not yet switched on or the NWS has switched to bypass, because the nose wheel is not in line with rudder pedal or another failure has occurred. 25. Sufficient airplane controllability during take-off/landing and all ground maneuvering with NWS bypass light illuminates can always be assured by means of differential power brakes and rudder. 26. The rate of change of the signal from the rudder pedals is electrically controlled by a ramp function in the control unit circuit of the control unit so that the rate of change of the nose wheel position limited to maximum of 18° per second. RESTRICTED  RESTRICTED 44 27. A bypass signal is also produced if a failure occurs in the required/actual position, potentiometer, the servo valve or the servo channel electronics. A signal applied via valve switch contacts to the warning light panel to illuminate NWS by pass annunciator. During retraction relay 5 DA is energized if the bypass valve is in the bypass mode the NWS by pass annunciator is connected with GND and illuminated. The NWS by pass light is prevented from illuminating during extension via micro switch 7 GA, 12 GA, 14 GA. 28. When the aircraft is airborne nose gear weight switch 5 GA opens and the control logic causes the nose wheel to return to the neutral position. The rudder pedals now have no effect on the nose wheel position. The maximum deflection of the nose wheel is limited to +8° during takeoff and landing and 0o when air borne. The required limits are controlled by control logic in the control unit switched by steering switches and actuated by the speed levers and steering button 4 GH. (a) The 45° steering authority is applicable when following condition are fulfilled:- (i) Both speed lever set to low. (ii) Nose gear weight switch closed (weight ON) (iii) Nose gear extended and locked down. (iv) Steering press button 11 GH depressed. (b) The 8° steering limits is applied during landing and takeoff when the following conditions apply:- (i)) LH/RH speed levers set to High/Low. (ii) Nose wheel extended and locked down.(Weight Switch closed). (iii) Nose wheel steering switch 'ON'. (c) The nose wheel is centered to 0° under the following conditions:- (i) With weight off the nose gear. 29. When 45° steering limit is applied 28V DC is supplied to the master warning system and the NWS 45° annunciator illuminates. To prevent damage to the nose wheel during retraction the nose wheel must be centered. If the nose landing gear not centered the nose landing centering switch is opened and the bypass valve is de-energized. The landing gear cannot be retracted. 30. When landing with weight ON there is 0.5 seconds delay before the 8 ° limit is applied. This is controlled by the control logic of the control unit. Bibliography: Airplane Maintenance Manual Chapter- 32 RESTRICTED  RESTRICTED 45 AIRCRAFT ACCIDENTS/INCIDENT CASE STUDY 1. Nose Wheel Steering (NWS) Bypass Light 'On' During Short-Finals. (a) KD-714 dated 20 Nov 19, 19 Wing (59 Sqn) (b) Brief On 20 Nov 19 at 2015 hrs, Nose Wheel Steering (NWS) Bypass light came 'On' in 00-228 ac KO-714 during short finals at 10 Wg. After landing, NWS and brakes were not effective and aircraft gently veered to the left. Aircraft was stopped on the left lane of runway using reverse thrust and parking brakes. Aircraft was switched-off on runway after suspecting hydraulic failure. (a) Reason Failure of Hydraulic Power Pack. (b) CAS Remarks N/A (c) Action Taken Aircraft was recovered back to base after changing the Hydraulic Power Pack. (d) Lessons Learnt / Remedial Action A study on frequent failure of Hydraulic Power Pack is in progress at HAL (TAD). 2. NWS By Pass Caution Light Came ON. (a) KD-714 dated 19 Aug 20, 19 Wg (59 Sqn). (b) Brief On 19 Aug 20, aircraft KD-714 aircraft was on routine cat I sortie (day circuit and landing) at Guwahati. The sortie was uneventful till NWS By pass caution light came ON while hydraulic pressure was found to be normal (206 bars). A normal landing was carried out. On landing roll while attempting to control direction through differential braking, the brakes were found to be non-responsive. Reverse power was used to slow the aircraft down and the aircraft was stopped with the help of parking brakes. Aircraft was switched off on the runway. Crash crew was activated and aircraft was cleared from the runway. Oil leak from the defective Check valve had led to draining out oil from the reservoir. (c) Reason Faulty check valve. (d) Action Taken Defective Check valve replaced with serviceable check valve and aircraft cleared for operation. (e) Lessons Learnt/ Remedial Action Failure of check valve has happened for first time, maintenance crew are required to be more vigilant 10 during servicing/DI and due attention may be given to any leakage/seepage of hydraulic oil in Hydraulic system. RESTRICTED  RESTRICTED 46 AIRCRAFT ACCIDENTS/INCIDENT CASE STUDY 3. NWS Bypass Caution Light Came ON. (a) HM-688 dated 26 Aug 20, 19 Wg (59 Sqn). (b) Brief On 26 Aug 20, aircraft HM-688 aircraft was on RTR sortie from Tezpur to Guwahati. The sortie was uneventful till approaching for landing at Guwahati. After lowering landing gear, NWS By pass caution light came ON and hydraulic pressure was found to be zero. A normal landing was carried out. Brakes were found to be non-responsive, hence directional control was maintained with the help of differential power and the aircraft was brought to a halt on the runway with the help of parking brakes. Crash crew was activated and immediately aircraft was cleared from the runway. (c) Reason Faulty Hydraulic Power Pack. (d) Action Taken Defective Hydraulic Power Pack has been replaced with serviceable component as per Chapter 29-10-19 of AMM and aircraft cleared for operation. (e) Lessons Learnt/ Remedial Action HAL (TAD) has carried out the root cause analysis for frequent failure of HPP brushes and initiated measures to enhance the reliability. Accordingly, SI on motor of HPP was issued by OEM (M/s Thales) vide dte-12-196. The SI is implemented on motor Sl no 1482 at 59 Sqn, which is under performance trials for 6 MO/100 FH. In addition, an additional brush bedding facility is also developed at 5 BRD. 4. NWS Bypass Caution Light Came ON. (a) KD-720 dated 27 Aug 20, 3 Wg (41 Sqn). (b) Brief On 27 Aug 20, aircraft KD-720 aircraft was started up at 41 Sqn for DP- Jammu commitment and the start-up of aircraft was normal. The hydraulic system was functioning satisfactorily and all other aircraft parameters were observed normal. During taxi out, at a distance of approximately 50 meters from start-up point, NWS By pass caution light came on and hydraulic pressure was checked and observed gradually reducing to zero. A precautionary shut down was carried out by crew immediately after stopping the aircraft with help of parking brakes. The aircraft was towed backed to the unit dispersal. (c) Reason Faulty Hydraulic Power Pack. (d) Action Taken Defective Hydraulic Power Pack has been replaced with serviceable component as per Chapter 29-10-19 of AMM. Aircraft cleared for operation. (e) Lessons Learnt/ Remedial Action HAL (TAD) has carried out the root cause analysis for frequent failure of HPP brushes and initiated measures to enhance the reliability. Accordingly, SI on motor of HPP was issued by 11 OEM RESTRICTED  RESTRICTED 47 (M/s Thales) vide dte-12-196. The SI is implemented on motor Sl no. 1482 at 59 sqn, which is under performance trials for 6 MO/100 FH. In addition, an additional brush bedding facility is also developed at 5 BRD. ********** RESTRICTED RESTRICTED 48 RESTRICTED  RESTRICTED 49 CHAPTER-8 FLIGHT CONTROLS SYSTEM Flap System 1. The flaps are driven by an electric motor through screw jacks and bell cranks. The flap motors is equipped with an electro-magnetic brake. The safety clutch is incorporated to protect the system against overloading. A torque limiter is incorporated to protect the shaft drive and the screw jacks against over-stressing. 2. The flap motor is powered by 28V DC and protected by the CB labeled, ‘FLAPS CTRL’/ESS Bus on the overhead CB panel. Flap position is indicated on the flap position indicator. The flap position indicator is powered by 28V DC and is protected by the CB labeled, ‘FLAPS IND’ /ESS BUS on the overhead CB panel. Flaps Selector Switch 3. Flaps extension and retraction is accomplished by the flaps selector switch located on the central pedestal panel. It permits the selection of three fixed positions UP, 1 & 2. The DN position is mechanically locked. To raise the flaps from 1 to UP, the flaps lever must be pulled out and placed in the UP position. The FLAPS lever position marks are internally illuminated. They can be dimmed by the PANEL rheostat. Flaps Position Indicator 4. The FLAP position indicator is located on the central pedestal panel and is labeled UP, 1, 2 and DN. Aileron Trim Control System 5. An electromechanical trim actuated by a three-position toggle S/W with a spring loaded centre position which is located on the centre pedestal panel is attached to the artificial feel springs. To trim the airplane the neutral position of the artificial feel springs is changed to re-position the ailerons. The trim position indicator located in the central pedestal panel indicates the position of trim (ailerons). 28V DC electrical power for the trim motor is provided by the CB labeled,‘AIL TRIM’ /ESS.BUS on the overhead CB panel. Stabilizer Trim Control System 6. The aircraft is equipped with an electromechanical trim actuator which positions the horizontal stabilizer. To avoid overloading of the electrical trim motor the stabilizer trim actuator is equipped with a mechanical clutch. The clutch protects the trim motor against overload. Stabilizer deflection is 4.8o Nose UP and 1.6o Nose DN. 7. The three position stabilizer trim control S/W is located on the outboard horn of each control wheel. Each switch has a Nose UP and Nose DN position and is spring loaded to the centre OFF position. Each S/W consists of two self-contained switches RESTRICTED  RESTRICTED 50 combined into a single unit and must be depressed simultaneously to activate the horizontal stabilizer trim. 8. The trim motor will stop if the pilot and co-pilot actuate their respective trim switches in opposite direction. 9. The trim position indicator is located on the pedestal panel. The green range is marked for normal take-off. Note: (a) Incorrect pitch trim settings may result in high stick forces. The system is designed fail passive. A single electrical failure in the system may render the system inoperative. But no single electrical failure in any part of the system will cause runaway of the actuator. (b) 28V DC electrical power for the trim motor is provided through the CBs:- (i) ‘PITCH TRIM UP’/ESS BUS on the overhead CB panel. (ii) ‘PITCH TRIM DN’/ESS BUS on the overhead CB panel. Bibliography: - Airplane Maintenance Manual Chapter- 27 ********** RESTRICTED  RESTRICTED 51 CHAPTER-9 STARTING AND IGNITION SYSTEM Introduction 1. The engine can be started on the ground or in the air either automatically or manually. 2. For ground starting, the starter-generator drives the engine from 0% to 55% RPM. Fuel and ignition are introduced automatically at 10% RPM through an electronic speed switch or manually through the MAN IGN switch. At 55% RPM, electric power to the ignition unit and the starter is switched off either automatically by an electronic speed switch or manually by the MAN IGN switch. The engine-side fuel shutoff valve remains open due to its internal spring Belle-Ville spring. 3. For air starting, propeller wind milling forces are used instead of the starter. The propeller is driven out of the feather position by oil pressure from the unfeathering pump. At 10% RPM, fuel and ignition are introduced automatically or manually in the same sequence as described above. 4. During engine starting, fuel can be added manually to assist in starting of a cold engine by pressing the fuel ENRICH button. Starting System Components 5. The starting system comprises the following components:- - START CONTROL Circuit Breakers 1JA and 11JA - IGNITION Circuit Breakers 1JB and 11JB - START Selector Switch 2JA - START/STOP Switches 3JA and 13JA - Various Relays in the System - Battery Switching Relay 6JA - MAN IGN Switches 2JB and 12JB - IGN lamps 5JB and 15JB - Speed Switch Assemblies 1EN and 2EN START Selector Switch 6. The START selector switch located in the middle of the starting panel, is used to select the mode of start for both engines, It is a three position toggle switch. The three positions control the following:- (a) GROUND When set to this position, the GROUND mode of start may be initiated by pushing the START/STOP switch briefly to the START position. The starter winding of the starter generator is energized and cranking of the engine begins. (b) VENT If the START selector switch is set to the VENT position, the START/STOP switch must be held in the START position for cranking the engine. RESTRICTED  RESTRICTED 52 This position allows cranking without the ignition and fuel shut-off valve circuits being activated. (c) AIR When set to this position, the AIR mode of start may be initiated by pushing the START/STOP switch briefly to the START position. The unfeathering pump is energized, and the engine is cranked by wind milling action of the propeller. Ignition and fuel shut-off valve circuits will be automatically energized at 10% RPM or should be made available manually by setting the MAN IGN switch to START. Start/Stop Switch 7. The START/STOP switches are located to the left and right of the START selector switch. It is a three-position toggle switch, with a guarded centre (OFF) position. The other two positions are used for the following:- (a) START If pushed briefly to this position, power is applied to the starter winding of the starter generator, provided the START selector switch is in the GROUND OR VENT position. If the switch is pushed briefly to the START position with the START selector switch set to the AIR position, the unfeathering pump will be energized. (b) STOP The STOP position is used to shut down the engine. The CLOSE circuit of the fuel shut-off valve and the purge solenoid are energized simultaneously. The switch should be held in this position for about 5 seconds for efficient purging action. Manual Ignition- 8. Manual ignition is provided for the following reasons:- - To enable manual AIR or GROUND starts. - To allow battery starts in MAN mode, when battery voltage is only 24V. - To enable an engine start with the speed switch defective. - To prevent a flame-out under icing/rain conditions - To enable an engine start when residual ITT is more than 300°C. Manual Ignition Switch 9. The three-position toggle switch has the following positions/functions:- (a) CONT In this position the ignition circuit is kept continuously energized. The IGN light remains continuously lit up. (b) OFF This position allows energisation of the ignition circuit during an automatic AIR/GROUND START. (c) START This position energizes the ignition circuit and the OPEN solenoid of the fuel shut-off valve. RESTRICTED  RESTRICTED 53 Engine Starting Modes, Key Steps of Operation Note: Before starting engines, it is necessary to carry out a pre-flight inspection of the airplane, in accordance with the Pilot's Operating Handbook. Automatic Ground Start 10. - Set START selector switch to GROUND. - Set respective START/STOP switch to START and hold momentarily. Power is supplied:. To the coil of the start relay and its hold-in circuit.. The START/STOP switch can be released.. To the starter which cranks the engine.. To the oil vent valve which opens.. To the power supply relay of the torque/temperature limiting system, this opens and prevents operation of the system. - When engine RPM has reached 10%, the "10% contacts" in the speed switch assembly close and power is supplied:. To the ignition unit which fires the igniter plugs.. To the two ignitions indicating lamps which illuminate.. To the "open" circuit of the engine-side fuel shutoff valve which opens and allows fuel flow to the engine fuel nozzles.. To the anti-ice lockout valve which closes and prevents fuel flow through the oil-fuel heat-exchanger.. To the fuel enrichment switch which, if pressed, opens the fuel enrichment valve and allows additional fuel flow to the engine to achieve sufficient acceleration. - When engine RPM reaches 55%, the "55% contacts" in the speed switch assembly open which interrupts power supply to:. The starter, now it becomes as generator.. The ignition unit, that stops the firing of the igniter plugs.. The ignition lamps which extinguishes.. The oil vent valve which closes. RESTRICTED  RESTRICTED 54. The "open" circuit of the high pressure fuel shutoff valve. However, the valve will remain open due to its internal belle-Ville spring.. The anti-ice lockout valve which re-opens.. The power supply relay of the torque/temperature limiting system which closes and re-powers the system. Note: The ENGINE LIMIT switches are ON for engine starting. Manual Ground Start - Set START selector switch to VENT. - Set respective START/STOP to START and hold. Do not release the Switch. Power is supplied:. To the coil of the start relay. In the manual starting mode, the relay is not provided with a hold-in circuit, so the START/STOP switch must be held in the START position.. To the starter which cranks the engine.. To the oil vent valve which opens.. To the power supply relay of the torque/temperature limiting system, this opens and prevents operation of the system. Note: When engine RPM reaches 10%, the closing of the 10% speed switch contacts will not energize the power supply relay of the fuel and ignition circuits. The control circuit to the relay coil will remain open since the START/STOP switch is held in the START position. - When engine RPM has reached 10% to 15%, set the manual ignition switch (MAN IGN) to the START position. This will power the components as outlined in the AUTOMATIC GROUND START. Ventilation- 11. Vent start or DRY start can be carried out either with internal batteries or GPU. Ventilation will continue as long as the start switch is held in the 'start' position but should be allowed to continue only for 20 seconds maximum or 17% RPM whichever is earlier. RESTRICTED  RESTRICTED 55 Automatic Air Start - Set START selector switch to AIR. - Set respective START/STOP switch to the START position and hold momentarily. Power is supplied to: * To the unfettering pump. Operation of the unfeathering pump brings the propeller out of the feather position. Decreasing propeller pitch angle will cause the propeller to windmill and crank the engine. * To the power supply relay of the torque/temperature limiting system which opens and prevents operation of the system.  To the oil vent valve which opens Note: The negative torque sensing system will control the propeller pitch angle according to the increasing RPM to minimize drag. - When engine RPM has reached 10% the "10% contacts" in the speed switch assembly close and power is supplied to the systems and components as outlined in the "AUTOMATIC GROUND START. - When engine RPM has reached 55%, the "55% contacts" in the speed switch assembly open which interrupts power supply to:-.The unfeathering pump.The system and components as outlined in the “AUTOMATIC GROUND START". Manual Air Start The START/STOP switch is not to be actuated to initiate a manual air start. - Set START selector switch to AIR. - Unlock the unfeathering pump switch, set and hold the switch in the ON position. Operation of the unfeathering pump brings the propeller out of the feather position. Decreasing propeller pitch angle will cause the propeller to windmill and crank the engine. - The negative torque sensing system will control the propeller pitch angle according to the increasing RPM to minimize drag. When engine RPM has reached 10%: - Release the unfeathering pump switch to the OFF position and close the Guard it. RESTRICTED  RESTRICTED 56 - Set the manual ignition switch (MAN IGN) to the START position between 10 to15% RPM. This will power the components as outlined in the AUTOMATIC GROUND START. - When engine RPM reaches 10%, the closing of the 10% speed switch contacts will not energize the power supply relay of the fuel and ignition circuits. The control circuit to the relay coil will remain open since the start relay is not energized during a manual air start (START/STOP switch is not actuated). - When engine RPM reaches 55%, set the manual ignition switch to the OFF position. This will de-energize.. The ignition unit, stopping the firing of the igniter plugs.. The ignition lamps which extinguish.. The "open" circuit of the high pressure fuel shutoff valve. However, the valve will remain open due to its internal belle-Ville spring mechanism.. The anti-ice lockout valve which re-opens.. The fuel enrichment switch. Bibliography: - Airplane Maintenance Manual Vol. V Chapter- 74 ********** RESTRICTED  RESTRICTED 57 CHAPTER-10 AIR-CONDITIONING SYSTEM Fig. No. 10-1 Pneumatic Panel Introduction 1. To operate Air conditioning system, a pneumatic control panel is installed on the co-pilot's side of main instrument panel. Air-conditioning is possible when either of bleed air switches are in the ON position. The three position air-conditioning selector switch must be in NORMAL/HIGH position for normal operation. In the position RAM AIR electrical power to the air-conditioning master switch relay will be interrupted and the RAM AIR shut-off valve will Closed. The TEMP MODE S/W is either in the AUTO or MAN POSITION. In the AUTO position temp is maintained as selected by the heat control rheostat by means of a temperature sensor and temperature controller. In the MAN position setting the three position toggle switch either to HOT OR COLD can control temperature. The S/W is spring loaded to the Centre (OFF position). 2. The system includes an over temperature S/W which monitors the temperature in the supply ducts to the compartments. In the event that the duct temperature exceeds the nominal value (65°), the S/W supply the signals and illuminates the amber CABIN TEMPERATURE caution light. 3. If this occurs in AUTO mode, change the ‘TEMP MODE Switch’ to MAN position and place the ‘MAN CTRL S/W’ to the COLD position for 10-15 seconds. The ‘CABIN TEMP’ light must extinguish. 4. The over pressure switch monitors pressure regulation of the pressure reducing and shut off valve. Whenever the supply air line pressure reaches 68 –76 PSI, ‘BLEED PRESS’ caution light comes on. RESTRICTED  RESTRICTED 58 5. Distribution of air in the cockpit is controlled by the ‘CKP AIR’ slider. Depending upon slider position more or less air will flow into the cockpit via the floor outlets. Demisting is controlled by the DEMISTING slider between full open (UP) or Closed (DN).The temperature control valve (TCV) controls the cabin temperature by regulating the flow of hot air to the mixing chamber. The operation of TCV is controlled by temperature controller which has two modes operation namely AUTOMATIC and MANUAL. 6. If NLG is extended and one or both speed levers are set to more than 85% RPM then the switch AIR COND ON / OFF switches automatically comes to OFF position and shuts down the air conditioning system to avoid an excessive engine power loss during take off and landing. Cockpit Fans Fig. No. 10-2 Cockpit Fans Switches There are two cooling fans installed in the cockpit, one in pilot’s side and the other in the co-pilot’s side and are meant to provide the mean of Positive airflow to pilot and co-pilot during the ground operations and in flight. They are 28V DC fans powered from the NON- ESS BUS on the pilot’s AC/DC CB panel and protected by the circuit breaker labeled “Cockpit Fans” (3 amps). The cooling fans are controlled by their corresponding switches “FAN PILOTS” installed in the panel-RH in Front avionics rack near the Elapse time Indicator for air-conditioning system and hydraulic power pack and “ FAN CO-PILOT “installed in the CB panel LH in front avionics rack. Note: Switch OFF the cockpit fans during R/T communications, if any noise of airflow is heard in the headset or adjust the fans to direct the airflow away from the microphones. Bibliography: - Airplane Maintenance Manual Chapter- 21 ********** RESTRICTED  RESTRICTED 59 CHAPTER-11 PROPELLER SYSTEMS Introduction 1. Engine power is transmitted to the propeller via the reduction gear section and the prop shaft. Propeller blade pitch is controlled by the engine power lever and prop (engine) RPM by the speed lever. During flight the propeller is controlled by the power lever in the range between flight idle and maximum and by the speed lever in the range between low and high. Propeller pitch angle is controlled by oil pressure which enters the piston through the beta tube in the centre of the propeller assembly. Increasing oil pressure forces the piston forward. This movement is transmitted by the pitch change links to the blades. 2. The propeller control system is designed to operate in either of the modes namely, propeller governing mode, or flight mode and Beta mode or ground operating mode. In flight mode oil pressure to the propeller is metered by propeller governor as function of RPM. In Beta mode oil pressure to the propeller is metered by propeller pitch controller which is mechanically controlled by power lever. 3. If it is required to feather the propeller during flight the engine speed lever is first moved to the shut off position and then to the FEATHER position. This actuates the feathering valve and the blades are then moved to feather position by a combination of spring counter weight and aero dynamic force. In feather position, the blade face straight into oncoming air stream to reduce drag by preventing wind milling of engine. Propeller Unfeathering System Description 4. Unfeathering system is provided to bring the propeller out of the feather position for an air start or to engage the start locks prior to initiating an engine start. The main components of the system are:- (a) Unfeathering Pump (b) Unfeathering Pump Switch (c) Prop System Circuit Breaker The prop blades are unfeathered by an electrically driven oil pump. Operation 5. If the engine speed lever is moved away from the FEATHER position to low and the power lever in the AIR START position (Blue line on quadrant) and AIR VENT GND switch is set to AIR position the unfeathering pump energized either by the START/STOP switch when moved to the START position (Automatic air start) or by the UNF pump switch set to the 'ON' position (Manual air start). Oil pressure delivered by the unfeathering pump will bring the propeller out of the feather position. The propeller starts to wind mill and cranks the engine. RESTRICTED  RESTRICTED 60 Beta Pressure Switch 6. When the power lever is positioned between reverse thrust and flight idle oil pressure to the propeller is metered by propeller pitch controller and at this stage selected prop blades allows for maneuvering aircraft during ground operation. 7. The Beta pressure switch is installed in the prop hydraulic control system. This switch is used to power an indicator light referred as a Beta light located on the instrument panel. This light is illuminated when the propeller is operating in Beta mode and is out when propeller is operating in PG (prop governor) mode because during Beta mode the oil pressure is high and the switch contacts closes at 360 ± 10 PSI to provide indication in the cockpit. The blue beta light will illuminate whenever the corresponding propeller is at the minimum blade angle. Negative Torque Sensing System 8. The function of negative torque sensing system is to limit the negative torque caused by the wind milling propeller and driving the dead engine, thus preventing the large drag force on the airplane. If the engine suddenly loose power in flight the NTS system automatically changes the propeller blade angle towards higher pitch angle to minimise the negative torque. 9. The main components of the system are:- (a) Torque Sensor (b) Propeller Feathering Valve (c) Speed Pick Up 10. The system can be tested by operating the NTS test switch. 28 V DC is supplied to the unfeathering pump and the NTS check out solenoid valve. The pump runs and the solenoid valve opens allowing oil pressure to build up and act on the negative torque pressure switch. The switch contact closes and illuminates the NTS test lamp and indicates the system is ready. If light do not illuminate the system is inoperative if the power lever is in the normal flight range or indicates a cut-out-function when the power lever below FI. NTS lock out and propeller governor reset. NTS lock out dumps the NTS supply pressure. At the same time propeller governor resets 105% RPM so that minimum pressure is supplied to the propeller control, to ensure proper propeller response during rejected take-off and high landing speed. Propeller Synchrophasing System 11. The propeller synchrophaser system automatically adjusts propeller governor speed setting to match the speed of the two engines. A magnetic pick up mounted on a special bracket behind the propeller spinner transmits speed signal to the control box. These impulses are compared in the control box which sends corrective signal to the coils in the governor, thus synchronizing the engines, both the system operates on 28V DC and is protected by, ‘PROP-SYNCH’ CB /NON ESS BUS on the AC/DC CB panel. RESTRICTED  RESTRICTED 61 Control and Operation 12. If the engines are manually synchronized within 0.5% RPM engine speed and the SYNCHROPHASER control knob is turned ON, the engines will be synchronized to a zero speed difference. The synchronized engine RPM may differ up to 1% RPM from the original setting. Turning the knob into the range labeled PHASE-SELECT will place the propellers in the desired phase relation with each other to minimize propeller noise. 13. The system must be OFF during take off, landing and single engine operation. Any subsequent power changes can be made by moving the ENGINE SPEED levers closely together. 14. Malfunctioning of the synchrophaser cannot prevent normal operation of the engines. In case the synchrophaser system is unable to match speeds, turn the system to OFF, synchronize manually, then turn the system to ON. Propeller Deicing System 15. The system consists of an electrically heated de-icer mat fitted to the leading edge of each propeller blade. Each mat contains separate inner and outer heating circuits which are alternately supplied with 28 V DC via a timer, brush block assemblies and slip rings. The timer cycle is 68 seconds. Operation 16. When the DE-ICER PROP switch is set to ON the relay energizes and supplies 28 V DC via a shunt to terminal 'B' of timer. The timer now connects 28 V DC to the inner de-icing circuit of all four propeller blade de-icing mats for a period of 34 seconds. On completion of this first half cycle the timer switches the supply from the inner to the outer circuit for a further 34 seconds to complete the cycle. The cycle is repeated as long as the DE-ICER mats current consumption is indicated on the ammeter fitted on the center pedestal panel in cockpit. The system has to be switched ON before entering icing zones. When the system is switched ON, the heating cycle will start where the previous cycles has left OFF. Indication is on the two ammeters on the centre pedestal panel. 17. Connection to the rotating propeller is accomplished by the brush block assembly and the propeller slip rings to the inner/outer de-icing circuits. The normal current consumption of the each propeller circuit should be 24 to 28 amps (Max 34 Seconds) for long mats and 22 to 26 Amps for small mats (Max 45 Seconds). Caution Propeller de-icing must not be switched on as long as propeller is stationary to prevent over heating of mats. Bibliography: - Airplane Maintenance Manual Chapter- 61 ********** RESTRICTED RESTRICTED 62 RESTRICTED  RESTRICTED 63 CHAPTER-12 AIRCRAFT FUEL SYSTEM Introduction 1. The complete fuel system contains a separate tank group and supply system for each engine. The integral wing tank consists of separate feeder, inner, outer and the optional auxiliary tank in each wing. The wing centre section above fuselage does not contain any fuel. 2. Fuel is transferred from the outer and inner tanks to the feeder tank by booster pump driven jet pumps. Two jet pumps are installed in the outer and two in the inner tank. If the optional auxiliary wing tanks are fitted, an additional jet pump in each auxiliary tank pumps fuel via the existing transfer line directly into the respective feeder tank. Motive flow for this jet pump is taken from the existing motive flow line. The booster pump also supplies fuel from the feeder tanks to the engines. Indication of fuel quantity, low fuel level and low pressure is provided in the cockpit. 3. The separate tank groups can be interconnected by means of the cross feed system. Technical Data Tank capacities without optional auxiliary wing tanks Total usable fuel LH wing : 1193 LTR (2078 lbs) Total usable fuel RH wing : 1193 LTR (2078 lbs) Total usable fuel both wings : 2386 LTR (4156 lbs) Total fuel both wings : 2441 LTR (4251 lbs) Tank capacities with optional auxiliary wing tanks Total usable fuel LH wing : 1425 LTR (2482 lbs) Total usable fuel RH wing : 1425 LTR (2482 lbs) Total usable fuel both wings : 2850 LTR (4964 lbs) Total fuel both wings : 2895 LTR (5042 lbs) RESTRICTED  RESTRICTED 64 Fuel Booster Pumps Description 4. The booster pump feed fuel from feeder tank to engine driven HP pump. It also supplies motive flow for jet pumps. 5. Two fuel booster pumps (immersion type DC fuel pumps) are fitted in the appropriate feeder tank. Pump 1 (aft) is the main pump and pump 2 (forward) is an auxiliary pump. The booster pump consists of two main assemblies, pump assembly and motor assembly. The pump assembly includes the body base casting, outlet port, impeller cavity, inlet diffuser and inlet screen. The pump is mounted in the aircraft via lugs on the body base casting. 6. The motor assembly includes the pump upper body section, armature, commutator, carbon brushes, impeller and fuel lubricated carbon bush bearing. The motor is a permanent magnet type. The three check valves in each collector tank are identical. The fourth valve, located in the fuel supply line to the jet orifice pumps is of different type. (i) The two check valves just downstream of booster pumps prevent reverse flow from an operational pump to a non-operational pump. (ii) The check valve in the motive fuel supply line prevents air being drawn into the engine fuel supply line if both booster pumps are u/s. (iii) Check valve downstream of the flow divider prevents fuel flow to the jet orifice pumps and thereby filling the tanks, if the cross feed valve is open and fuel is supplied from the opposite tank group. 7. A bypass in the lower body allows fuel flow to the engine if the electric motors of the both booster pumps are failed and the fuel is drawn to the engine by the engine driven fuel pump. However this mode of operation is applicable at altitude less than 10000 ft to ensure sufficient fuel flow. Fuel flow in to the non-operating pump is prevented by the check valves. The pumps are fuel lubricated. Technical Data Operation Voltage : 28 V DC Current Consumption : 8 A (Pump immersed) Fuel Flow : 1000 Lb / hour at 20 PSI minimum. Pressure : 1.8 bars (26.1 at Zero flow) 8. Two pump switches on the LH overhead switch panel energise the booster pumps of the fuel system via the following bus bars and circuit breakers:- (a) LP PUMP LH 1 CB of BUS-1 on the Overhead CB Panel. (b) LP PUMP LH 2 CB of BUS-2 on the AC/DC CB Panel. (c) LP PUMP RH 1 CB of BUS-2 on the AC/DC CB Panel. (d) LP PUMP RH 2 CB of BUS-1 on the Overhead CB Panel. RESTRICTED  RESTRICTED 65 LP Cock Description 9. The engine fuel supply system feeds the fuel from feeder tank to engine. This system consists of following components:- - Booster Pump - LP Cock 10. A fuel LP cock is installed in each engine nacelle on the forward wall of the inner tank. It is controlled by a 28 V DC motor which is operated by a guarded switch labeled FIRE COCK in the LH overhead switch panel in the cockpit. 11. Fuel from the booster pump circuit passes under pressure to the LP cock via a check valve and a pressure relief valve. When the LP cock is open fuel flows to the engine HP pump via a quick disconnect coupling. A pressure switch is fitted in the fuel line from the LP cock to the engine. The LP cock is a DC motor driven shut off valve with integral limit switches. The motor has separate windings for opening and closing the valve and each winding is supplied by its own circuit breaker. 12. During normal operation the LP cock is open allowing fuel to flow to the engine. The LP cock can be switched off during maintenance check. In emergency the cock is closed (shutting down the engine) automatically by a signal from the appropriate fire handle if operated. Operation 13. LP cocks are operated by switches labeled ‘FIRE COCK ‘. With the switch set to ON, the LP cock is opened and the motor is stopped by the integral limit switches. These limit switches also supply an 'OPEN' signal to the LP cock indicator in the cockpit. A mechanical ON/OFF indicator is fitted on the body of the cock for inspection purposes. 14. With the switch set to OFF, the LP cock is closed and the motor is stopped by the internal limit switches and supply a close signal to the indicator in cockpit. 15. The cock is closed automatically by the signal from the fire handle if operated. Technical Data (a) Operating Voltage : 28 V DC (b) Opening / Closing time : 1 Second max (c) FIRE COCK LH OPEN CB of BUS-1 on the Overhead CB Panel. (d) FIRE COCK RH OPEN CB of BUS-2 on the AC/DC CB Panel. (e) COCK LH CB of HOT BUS-1 in the LH Battery Compartment. (f) COCK RH CB of HOT BUS-1 in the LH Battery Compartment. RESTRICTED  RESTRICTED 66 Crossfeed System 16. The fuel supply lines to each engine are interconnected by a cross feed system consisting of a cross feed valve and piping. This enables both engines to be fed from one wing or one engine to be fed from both wings. Fuel cannot be pumped from one wing to the other. The cross feed valve is a DC motor driven shut off valve with integral limit system. 17. The cross feed valve is fitted in the cross feed piping in the area of the LH feeder. The valve has two position fully opened and fully closed. In normal operation, cross feed valve is closed and each engine is supplied with fuel from its own wing tanks. If for any reason, the pilot wishes to supply both engines from one wing or one engine from both wings that is in case of an engine failure he can open the cross feed valve using the cross feed switch. When the valve is opened its motor is stopped by its integral limit switches which also supply an open signal to the cross feed valve indicator in the cock pit. 18. To supply both engines from LH fuel system, switch LH booster pumps ON and RH booster pumps OFF, open the cross feed valve and check for ON indication. To supply one engine from both fuel systems, set all booster pumps to ON and cross feed open. Fig. No- 12-1 Cross Feed System RESTRICTED  RESTRICTED 67 Indication and Warning System 19. The indicating system gives the pilot, visual indication of fuel quantity with warnings for low level and low pressure. The low level warnings provides a separate indication for each side when fuel quantity in the feeder tank is 195 ± 10 lb. The low pressure warning provides separate indication for each side when the engine fuel feed pressure drops below 8.7 PSI. The pressure switch opens when the fuel feed pressure reaches to 14.5 PSI. The pressure switches are fitted between the LP cocks and engine connected hoses in LP fuel feed to each engine. In closed condition, the switches supply a ground connection to 2 lamps (Fuel pressure LH & RH) on central annunciator panel. Pressure Refueling System 20. A single point pressure refueling system can be fitted to the airplane as an optional system. The refueling connector is located in the right wing inner leading edge under an access door and the airplane can be pressure refueled or suction defueled via this connector. Fuel may also be transferred from one wing to the other (on the ground only) via the pressure refueling system. Description 21. The pressure refueling line leads from the connector to the refueling valves in the two inner wing tanks, and control lines connect each refueling valve to the fuel control pilot valves mounted in the top of the left and right inner tanks. A manual shutoff valve is mounted near the refueling connector and provides the connection to the cross feed line via which the feeder tanks can be defueled. A safety valve is installed in the bottom of each outer tank to prevent excessive pressure building up in the tank should the refueling valve fail during pressure refueling. The pressure refueling procedure is controlled from the refueling panel mounted below the refueling connector. 22. Maximum refueling pressure is 3.5 bars (50 PSI). System Components a. Refueling Panel b. Pressure Refueling Connector c. Refueling Valves d. Fuel Control Pilot Valves e. Safety Valves f. Manual Shut-Off Valve g. Refuel Valves (Electrically Operated) Switches of Pressure Refueling System 23. Following switches of pressure refueling system are mounted on Refueling panel. MASTER Switch In the ON position, 28V DC is supplied to the fuel quantity control units and solenoid valves of the fuel control pilot valves (via LH / RH REFU-SEL switch when selected). The drain port of each pilot valve opens and the system is ready for pressure refueling. When refueling is completed, the master switch must be set to OFF. RESTRICTED  RESTRICTED