Air Crew Propulsion PDF

RESTRICTED 11 CHAPTER-1 INTRODUCTION AND LEADING PARTICULARS Introduction 1. Dornier 228-201/202/202K aircraft is powered by two Garrett air research TPE-331-5/5B-252D/ TPE-331-5B-252D turbo-...

RESTRICTED 11 CHAPTER-1 INTRODUCTION AND LEADING PARTICULARS Introduction 1. Dornier 228-201/202/202K aircraft is powered by two Garrett air research TPE-331-5/5B-252D/ TPE-331-5B-252D turbo-prop engine. These engines are of the latest state of art technology and offer:- (a) Fuel Economy (b) Rapid power response (c) Variable prop RPM (d) Operation simplicity (e) Integral anti-icing (f) Reverse thrust (g) Torque and ITT limiting system (h) Low pollution and low noise (j) Long time between overhaul (k) Minimum maintenance requirement 2. TPE-331 Engines are manufactured by Garrett Turbine Engine Company of USA. The first TPE-331 series engine was produced in 1963 under the leadership of Mr. John Clifford Garrett. These engines are used in several types of aircraft because of its wide range of horse power. 3. TPE 331-5/5B-252D engines are rugged and reliable due to the integral gear box, two stage centrifugal compressors, three stage axial turbines and single annular combustion chamber. As a result of the single shaft cycle power response is instantaneous. Leading Particulars 4. Engine Type : Single shaft TPE with integral gear box. Direction of rotation (DOR) : Clockwise viewed from rear. Model number : TPE 331-5/5B-252 D Type of compressor : Two stages centrifugal. RESTRICTED  RESTRICTED 12 Compression ratio : 10:1 Combustion chamber : Single annular reverse flow Turbine : Three stages axial Dimensions of engine Dry weight : 164 Kg (360 Lbs) Length : 1.10 m (43.3") Width : 0.47 m (18.6") Height : 0.69 m (27.1") Maximum SHP : 715 Flat rated at 100% RPM E SHP : 775.8 HP T SHP : 904 HP Maximum RPM : 100 ± 1% (41,730 RPM) Over speed RPM : 104 ± 1% Cruising RPM : 96-97% Idling RPM : 70 ± 2% Self sustaining RPM : 55% Critical RPM : 18% to 28% Max. ITT limit at 100% RPM : 923° C Max. ITT during starting : 1149° C (for one second) Max residual ITT before starting : 300° C 5. Propeller Data RPM : 1591 at 100% RPM Reduction gear ratio : 26.22:1 Propeller type : Four blade Hartzell HC-B4 (TN) 5 ML Constant speed, feathering & reversible DOR : CCW viewed from rear Prop control : Constant speed single acting Prop tip speed limit : MACH 0.67 RESTRICTED  RESTRICTED 13 Prop diameter : 106" (2.69 m) 6. Engine Limitations Absolute Altitude : 35,000 feet Starting Altitude : Sea level to 20000 feet 7. Ambient Air Temperature - For operation : - 54 oC to 55 oC - For starting : - 40 oC to 55 oC Electrical power for starter : 28 V DC 600 Amps (1000 Amps Momentarily) Type of fuel used : ATF K-50 Fuel Capacity : 2441 Ltrs or 4252 Lbs (DO-228-201) 2895 Ltrs or 5042 Lbs (DO-228-202) Usable fuel : 2386 Ltrs (DO-228-201) 2850 Ltrs (DO-228-202) Fuel consumption : 213 Kgs per hour/ 500 lbs per hour appx. Type of oil used : Type 2 (OX-27) MIL-L 23699C Oil tank capacity : 6.25 Qtz/(5.9 Ltrs) Oil consumption : Max 01Qtz per 12.5 Hrs (01Ltr/13.2 Hrs) 8. Model Identification (a) The engine model number is given on the name plate attached to the gearbox. (b) The engine model number of the DO-228 aircraft is TPE-331-5-252D. (i) TPE - Turbo prop engine (ii) 331 - Engine manufacturer series indicator (iii) -5 - Power class (FAA Certificate) (iv) 252 - Engine configuration (specific location and configuration of Components) (v) D - Dornier (Aircraft Installation) RESTRICTED  RESTRICTED 14 Flat Rating 9. The engines having a greater thermodynamic power capability than required for the designed aircraft performance are often selected. These "over-sized" engines are then certified to the flat rated value in their specific installation. The nominal (flat) rated power is 715 SHP (533 KW) which is available at 33°C (91 °F) at sea level or up to 7300 feet at ISA temperatures. The benefits of using flat rated engines are:- (a) Lower turbine temperature at take off. (b) Improved altitude performance. (c) Longer engine life. Critical Speed Range 10. All turbine engines have a critical speed range, where the resonating frequency of the rotating parts can occur, leading to excessive vibration. This range on the TPE 331 engines is well outside the normal operational speed envelope of 70% to 100% RPM. The critical range of RPM in this engine is from 18% to 28% RPM. Hence the engine operation must be avoided in this speed range to prevent engine damage. During starting the engine should not be allowed to hung or dwell in the critical range. 11. The purpose of this is to caution maintenance and flight personnel that during the engine starting sequence immediate or accumulative engine damage may result if slow (tendency to dwell) engine acceleration is allowed through the engine critical speed range. Each start cycle should be monitored closely to ensure a smooth and continuous acceleration to ground idle RPM. If this is not the case, the start sequence should be terminated and the cause determined. Engine Cycle 12. One cycle is recorded as an engine operating sequence consisting of an engine start, take off, landing and engine shutdown. But when engine is started and number of take offs/landings are carried out before shut down then engine cycles are calculated as follows. 13. Equivalent cycles per start/shut down cycle = 1+ [(damage fraction) X (no. of take offs – 1)] where damage fractions for various turbine wheels are:- (a) 1st stage turbine - 0.5 (b) 2nd stage turbine - 0.6 (c) 3rd stage turbine - 0.2 14. Engine run to carry out maintenance actions, will not be counted as a cycle. RESTRICTED  RESTRICTED 15 Engine Stations 15. All turbine engine manufacturers utilize a station number identification for ease of description of various functions and locations within the airflow path. The adjacent picture shows that:- (a) Station number 1 represents the ambient conditions outside of the engine. (b) Station number 2 is the inlet to the compressor section. (c) Station number 3 is the discharge from the compressor section. (d) Station number 4 is the inlet to the turbine section. (e) Station number 5 is the exhaust discharge downstream of the turbine section. Fig. 01-01 RESTRICTED  RESTRICTED 16 Airflow Stations 16. The following table tabulates some of the major stations as used on the TPE-331 engines:- (a) P1 Ambient pressure T1 Ambient temperature (b) P2 Compressor inlet pressure T2 Compressor inlet air temp. (c) P3 Compressor discharge pressure T3 Compressor discharge temp. (d) P4 Turbine inlet pressure T4 Turbine inlet temp. (e) P4.1 2nd stage turbine inlet pressure T4.1 2nd stage turbine inlet temp. (f) P4.2 3rd stage turbine inlet pressure T4.2 3rd stage turbine temp. (g) P5 Turbine discharge pressure T5 Turbine discharge temp. STN NO 1 STN NO 2 STN NO 3 STN NO 4 STN NO 5 Fig. 01-02 Airflow Stations Bibliography: Airplane Maintenance Manual Vol 1 Chap 12, 20 & Garrett maintenance Manual Chapter 72 RESTRICTED  RESTRICTED 17 CHAPTER-2 HARTZELL PROPELLER (CONTROL SYSTEM) Introduction 1. The Hartzell HC-B4 TN-5 ML (Fig-02-01) is a four bladed constant speed, feathering and reversible propeller. The major components of the propeller are the four blades, hub, and spider assembly, piston assembly with pitch mechanism, blade start locks, spinner and spinner bulk head assembly. 2. Propeller pitch angle is controlled by oil pressure which enters the piston through the beta tube in the centre of the propeller assembly. In flight (prop governor) mode, (Power lever in front of the flight idle) oil pressure to propeller is metered by the propeller governor as a function of RPM. In beta-mode (power lever behind flight idle) oil pressure is metered by the propeller pitch control which is manually controlled by the power lever. Fig. 02-01 Hartzell Propeller Fig. 02-02 Propeller Governor Fig. 02-03 Propeller Pitch Control RESTRICTED  RESTRICTED 18 Propeller System Components Start Lock 3. A start lock assembly is installed aft of each propeller blade on the spinner bulk head to hold the blades at minimum pitch for a successful engine start and to avoid feathering of propeller to avoid aerodynamic load on the starter and power supply. Un-Feathering System 4. An un-feathering system is provided to bring the propeller out of the feather position for an air start or to engage the start locks. The major components of the system are:- (a) Un-feathering pump (b) UNF pump switch (c) Prop system circuit breaker (d) Plumbing system (e) Electrical wiring system Engaging the Start Locks 5. Before starting the engine, the propeller blades must be on the start locks. To engage the start lock when the engine is stationary speed lever should be in the low position and the power lever in a moderate reverse position, the UNF pump switch is set and held in the "ON” position. The energised un-feathering pump draws oil from the tank and delivers it to the propeller piston, through PPC and beta tube. When the blades move past the start lock pins, the pins get engaged and the UNF pump switch can be released. Now the power lever can be set to ground idle or "Flight idle position". Propeller on start lock can be checked by "Red on Red" alignment of spinner and blade root. 6. Engaging the start locks during shutdown is only a preparation of the propeller for next start. As the engine coasts down in speed after shutdown, move the power lever to reverse at 50% RPM or when oil pressure warning is heard. When the RPM reaches 10% power lever can be moved to GI or FI. Disengaging the Start Locks 7. After the engine has been started and accelerated to idling RPM, the propeller may be taken off the locks by merely moving the power lever towards reverse position. Procedure: (a) Speed lever... 85% RPM (b) Power lever... Reverse (slowly) (c) Observe... Beta light should remain ON (Increase in TQ) (d) Power lever... Back to GI (e) Speed lever... Low RESTRICTED  RESTRICTED 19 Un-Feathering For an Air Start 8. With the speed lever in the low and the power lever in air start position, the UNF pump is energised either by the start stop switch to start position (automatic start) or by the UNF pump switch set to the ON position (manual air start). Oil pressure delivered by the un-feathering pump will bring the propeller out of the feather position. The propeller starts to windmill and cranks the engine. Caution: Un-feathering pump to be operated to minimum of 10% RPM during air start cycle. When the engine has accelerated to its respective RPM, then un-feathering pump is switched off either automatically by the 55% speed switch or manually by releasing the UNF switch to the "OFF" position. Un-Feathering Pump 9. The un-feathering pump is mounted on the left side of the accessory gear box, consists of a DC motor and a pump assembly with an adjustable pressure relief valve. The pump draws oil from the engine oil tank and delivers it under pressure to the propeller piston or to the NTS system if an NTS ground check is performed. 10. Operation of the pump is controlled: (a) By the electric start circuit during an automatic air start. (b) By the UNF pump switch during a manual air start or for putting the propeller on the start locks. (c) By the NTS TEST switch during an NTS ground check. Technical Data (a) Operating limits : 1 min ON 3 min OFF (b) Max. operating pressure : 10.3 bar (150 psi) (c) Flow rate : 40 GPH Prop De-Icing System 11. The system consists of electrically heated mats fitted at the leading edge of each propeller blade. The left and right prop system is supplied with 28V DC from bus 1 and bus 2 respectively. The system is controlled by the two prop toggle switches located on the LH over head switch panel. Correct operation of the system is indicated by the ammeter fitted on the central pedestal panel. Normal current consumption is 24 to 28 amps. 12. When prop de-ice switch is set to on, the supply is provided to the timer. The timer automatically controls the supply to the heating elements by providing two heating cycles each for inner/outer mat circuits, for a period of 45 seconds. RESTRICTED  RESTRICTED 20 13. The system is protected by circuit breakers as follows: (a) DE-ICE PROP LH/BUS 1 located on the over head circuit breaker panel. (b) DE-ICE PROP RH/BUS 2 located on the AC/DC circuit breaker panel. Fig. 02-04 RESTRICTED  RESTRICTED 21 Negative Torque Sensing (NTS) System Introduction 14. The negative torque sensing system is composed of two major components: (a) Torque sensor (b) Feathering valve 15. NTS system on the TPE-331 engine is an automatic drag reduction system. The decision to feather the propeller rests with the pilot. The function of the NTS system is to limit the torque of the engine which can be absorbed from the propeller during wind milling and there by prevent high propeller drag conditions.The NTS system effects the movement of the propeller blades automatically towards feather position (should the engine suddenly lose power in flight) and precisely modulates the propeller blade pitch angle during propeller wind milled engine air start. 16. The torque sensor assembly is internally located in the reduction gear case. The function of torque sensor in the NTS system is to hydraulically actuate the propeller feathering valve when negative torque is applied to the engine. The feathering valve can be manually actuated by the speed lever. The feathering valve is externally mounted on the rear of the reduction gear casing. In response to the torque sensor assembly or by manual actuation the feathering valve blocks the high pressure oil flow to the pitch control and simultaneously relieves high pressure oil in propeller dome, which causes spring loaded propeller blades to move towards feather position. N.T.S. Lockout and Prop-Governor Reset 17. The high oil pressure from the prop governor is supplied to torque sensor and reset valve in the prop pitch control which controls the prop governor reset. When the power lever is positioned at flight idle or forward, the valve will be closed and oil pressure would be normal on propeller governor reset piston and on the NTS system. As the power lever is positioned below FI, the valve in the propeller pitch control is opened and the oil pressure is drained from the propeller governor reset system and the NTS system. Now there is no oil pressure in the system to cause the NTS action. 18. As the power lever is moved below flight idle, the oil pressure on the reset position is allowed to drain into the case. As the oil pressure is removed the heavy Belleville springs recalibrate the propeller governor to about 105%. NTS lockout and PG reset provides an instant response to beta control during the critical phase of landing. An additional minimum drag protection is also available in the event of NTS system malfunction. Move the power lever on the dead engine to full forward position. This positions the PPC follower sleeve to hold the propeller to the highest possible blade angle (resulting in lowest drag) under the conditions of not having NTS protection. This redundant safety feature is called the Beta follow-up system. N.T.S Control and Indication 19. To check the system, a NTS test button is provided on the overhead switch panel which activates the un-feathering pump to provide operating pressure under static RESTRICTED  RESTRICTED 22 condition. Operating status is indicated by the NTS light, which is provided in the overhead panel. When the NTS test button is depressed, the light will illuminate to indicate, that the system is ready. 20. NTS Ground Check (a) There are different NTS checks possible: A static NTS check without engine operation and the NTS check during and after an engine start. (b) The combination of all checks is recommended since the static check verifies the correct opening and closing of reset valve in the propeller pitch control (PPC). (c) To prevent the engine oil tank from being emptied by the un-feathering pump, no excessive time should be spent when carrying out the required steps prior to engine operation. 21. Static NTS Ground Check (Cut-Out Check) (a) When this check is carried out, the correct function of the following components is verified, NTS TEST switch and light, un-feathering pump, NTS check out solenoid valve, and reset valve in the PPC. (b) Set POWER lever to GI. (c) Depress and hold NTS switch, check that the NTS TEST light is extinguished. (d) Move POWER lever to FI, check that the NTS TEST light illuminates which indicates that the reset valve in the PPC has closed. (e) Move POWER lever back to GI, check that the NTS TEST light extinguishes which indicates that the reset valve in the PPC has opened. (f) Release NTS TEST switch. 22. NTS Ground Check during Engine Start (Functional Check) (a) When this check is carried out, the correct function of the entire Negative Torque Sensing System is verified, including the torque sensor and the feathering valve. (b) Carry out the steps as outlined prior to an engine start. (c) Set POWER lever to FI. (d) Depress and holds NTS TEST switch, check that the NTS TEST light illuminates. RESTRICTED  RESTRICTED 23 (e) Initiate engine start, check that the NTS TEST light extinguishes by reaching 5% RPM which indicates that the feathering valve has moved out of its normal rest position. Release NTS TEST switch. (f) When engine reaches 25% RPM, depress NTS TEST switch check that the NTS TEST light re-illuminates by 40% RPM which indicates that the feathering valve is back in its spring-loaded rest position. (g) Release NTS TEST switch, the TEST light extinguishes. (h) When the engine has accelerated to approximately 55% RPM, move POWER lever to ground start position (green line). Engine should accelerate to 70 + 2% RPM. 23. NTS Feather Valve Check (Before Starting Engine) (a) Power Lever - FI (b) NTS Test button - Press and hold (c) NTS Light - Check ON (d) Engine Speed Lever - SHUTOFF and FEATHER (e) NTS Light - Check out (f) Engine Speed Lever - LOW (g) NTS Light - Check ON (h) NTS Test button - Release 24. Supplementary NTS Check Procedure (a) Propeller - Check start lock released (b) Engine SPEED lever - LOW (c) POWER lever - Advance above FI to extinguish Beta lights, check RPM within appx. 93 to 96.5%. RESTRICTED  RESTRICTED 24 Fig. 02-04 Propeller Synchrophaser System 25. The synchrophaser system enables the pilot to maintain a desired propeller to propeller relationship in terms of RPM between the LH and RH propellers by automatically adjusting propeller governor speed setting there by reducing the noise level and vibration in the aircraft. A magnetic pick up mounted on the special bracket behind propeller spinner back plate transmits speed signals to the control box mounted on the central pedestal panel. The control box compares the signal received from both the propeller and sends corrective signal to the bias coil of each propeller governor. If the engines are manually synchronised to +/- 0.5 % RPM engine will be synchronised to zero speed difference level. The synchronised engine RPM may differ up to 1% RPM from the original setting. Turning the knob into the range labelled PHASE SELECT will place the propellers in the desired phase relation with each other to minimise the propeller noise. 26. The synchrophasing system should be switched off during engine starting, take off, landing, single engine operation and in case of system malfunction. Malfunctioning of the synchrophaser does not affect normal operation of the engines. In case the synchrophaser system is unable to match speed, turn the system to OFF, match speed manually and then turn the system to ON. 27. The system is powered by 28 V DC from the ESS BUS and protected by circuit breakers prop synchro LH and prop synchro RH on the overhead circuit breaker panel. RESTRICTED  RESTRICTED 25 Fig. 02-05 Synchrophasing System Bibliography: Airplane Maintenance Manual Chapter 61 & Garrett maintenance Manual Chapter 72, POH Section 4 RESTRICTED  RESTRICTED 26 Intentionally left blank RESTRICTED  RESTRICTED 27 CHAPTER-3 POWER PLANT CONTROLS Introduction 1. The power plant system is operated by a set of two controls for each engine. They are: (a) SPEED Lever. (b) POWER Lever. 2. These levers are placed at the cockpit quadrant. The respective levers are connected by continuous cable, pulleys and push rods to the various applicable components mounted on the engine. Fig. 03-01 Engine Controls Engine Speed Lever 3. The engine speed lever in the cockpit is connected to the USFG in the FCU, propeller governor, fuel shut off valve and feathering valve of the corresponding engine. It controls the engine speed according to pilot’s selection. 4. The lever has the following positions/range: (a) High The under-speed fuel governor in the fuel control unit is set to 97% RPM for BETA ground operating mode. The propeller governor is set to 100% RPM for propeller governing mode during flight. This position is also known as take off/landing position. (b) Cruise range The propeller governor can be set to as low as 96% RPM for propeller governing mode during flight. (c) Low The under-speed fuel governor in the fuel control unit is set to 70% RPM. The propeller governor is set to minimum stop at 94% RPM engine speed. This position is also known as taxing position. RESTRICTED  RESTRICTED 28 (d) Shut off range In this range the high pressure fuel shutoff valve is closed via a mechanical linkage. To move the lever from LOW to SHUTOFF, the lever must be lifted and pulled back against increasing spring pressure. (e) Feather This extreme position actuates the propeller feather mechanism. To hold the lever in this position, the lever must be pushed in. To release the lever out of the FEATHER position, pull the lever up. Adjustable Adjustable Friction brakes Friction Brakes Fig. 03-02 Engine Controls Engine Power Lever 5. The power lever in the cockpit is connected to the manual fuel valve in the FCU and propeller pitch control of the corresponding engine. The power lever has the following positions:- (a) Max Position to provide highest possible fuel flow (torque) for take-off and in flight. (a) FI (Flight Idle) Position to provide minimum fuel flow and minimum pitch stop for flight. (c) Start Two lines at the power lever segment recommended for respective engine start. (i) Air start marked by a BLUE LINE. (ii) Ground start marked by a GREEN LINE. (d) Reverse Position to reverse propeller blade angle and USFG reset for engine operation to prevent engine bog down. RESTRICTED  RESTRICTED 29 6. Adjustable friction brakes are provided to prevent the power and speed levers from creeping. The engine controls are also connected with aircraft systems by the help of micro switches. The various systems connected with speed lever are air-conditioning and NWS (Nose wheel steering) and the power lever is connected with main landing gear and stall warning. Fig. 03-03 Engine Controls Bibliography: Airplane Maintenance Manual, Chapter 76 RESTRICTED  RESTRICTED 30 Intentionally left blank RESTRICTED  RESTRICTED 31 CHAPTER-4 POWER MANAGEMENT Introduction 1. The power management system consists of the following major components:- (a) Propeller governor (b) Propeller pitch control (c) Fuel control unit PROP PITCH CONTROL PROP GOVERNOR FUEL CONTROL UNIT Fig. 04-01 Power Management Components 2. The inter-related power management controls provide the basic modes of operation: (a) Beta Mode or Ground Mode (b) Propeller governing mode or Flight Mode Beta Mode 3. Beta mode of operation provides for pilot to control propeller blade angle and selection of engine speed. The fuel control unit automatically meters fuel to match engine RPM according to propeller load. Beta mode of operation is normally utilised at reduced aircraft speed for reverse pitch braking on the runway and for ground operation such as taxing. During Beta mode operation of the propeller blade angle is controlled by propeller pitch control as set by power lever and the USFG controls the fuel supply as set by the engine speed lever. 4. A beta light (Fig 04-02) is furnished in cockpit to indicate when illuminated that the propeller has moved into the ground operating range after landing in response to power lever movement. RESTRICTED  RESTRICTED 32 Fig. 04-02 Beta Mode & Light Fig. 04-03 Beta Mode Propeller Governing Mode 5. During propeller governing mode the power lever is positioned between flight idle and take off. Main metering valve controls the amount of fuel metered to the engine for providing the desired power and the engine speed is controlled by propeller governor, so the engine is said to be operating in propeller governing mode. Propeller governing mode of operation is utilised for all airborne operation. Note: The power lever sets the fuel and propeller governor controls the engine RPM as set by the engine speed lever. RESTRICTED  RESTRICTED 33 Fig. 04-04 PG Mode Propeller Governor 6. The propeller governor provides a constant engine speed between 96-100% RPM according to engine speed during the propeller governing mode operation. The gear driven propeller governor assembly is composed of an integral gear type of pump, metering valve, and flyweight governor. Engine oil pressure is boosted and controlled into the propeller hub by the propeller governor via feathering valve, propeller pitch control and beta tube, to balance the feathering forces of the propeller. 7. During beta mode of engine operation the propeller governor assembly is not governing and supplies high pressure oil via the propeller pitch control to the engine propeller control components. Propeller Pitch Controller 8. The propeller pitch control or beta servo holds the cam which positions the propeller blade angle with metered oil through the Beta tube. 9. During propeller governing mode, propeller pitch control serve as oil passage and housing for the beta tube. In beta mode, the propeller pitch controls (PPC) the propeller blade angle. This is accomplished by positioning the beta cam by means of the power lever. Metered oil through the beta tube sets the desired propeller blade angle. Beta Pressure Switch and Beta Light 10. A beta pressure switch is installed in the propeller oil control system. The switch powers a blue beta light, located on the cockpit instrument panel. The beta light is illuminated when the propeller is operating in the beta mode and extinguish when the propeller is operating in the prop governing mode. RESTRICTED  RESTRICTED 34 Fig. 04-05 Bibliography: Garrett Maintenance Manual, Chapter 72 RESTRICTED  RESTRICTED 35 CHAPTER-5 ENGINE OIL SYSTEM Introduction 1. The engine lubrication system is a dry-sump system contains an integral pressure pump to provide jet and spray lubrication of engine bearings and gears. 2. It comprises of the following: (a) Tank (b) Oil pumps (c) An oil-fuel heat exchanger (d) An oil-air heat exchanger (e) Oil filter and by-pass valve with indicator (f) Oil vent valve Oil Tank Assembly 3. An oil tank is attached to the lower side of the reduction gear case to provide an oil reservoir for lubrication of the engine and the propeller governing system. The tank assembly includes a drain plug, a filler opening with filter, an air-oil separator, an overboard vent, a dipstick and a connection for supplying oil to the un-feathering pump. The oil tank is vented to the reduction gear case. The total oil tank capacity is 5.9 Ltrs /6.25 US Quartz. Useable capacity is 5 Ltrs/5.25 US Quartz. A dip stick is provided in the filler cap with markings "ADD" and "FULL". The oil level should be maintained at ‘FULL’. Oil Pump 4. The oil pressure pump is a rotor type pump keyed to the shaft of the fuel control unit drive gear. The pump delivers oil from the oil tank through a pressure regulator and a filter into the reduction gear housing, where the oil is directed to the bearings and gears. The pressure pump displaces approx 12 gals/min. If the pressure drop across the filtering element exceeds 50 to 60 PSI the filter by-pass valve will open allowing oil to bypass filter element to avoid bearing starvation. An external line carries lubricating oil to an oil jet in the turbine section for lubrication of the rear bearing. Pressurised oil is also provided to the propeller governor and the torque sensing system. Oil Scavenge Pump 5. There are three rotor type oil scavenge pumps. Two are located in the gear case and one in the tail cone. Oil returning from bearings, gears and from propeller governor, drains into sump in the bottom of the reduction gear housing. Oil from the sump is scavenged by two scavenge pumps to oil tank via air-oil heat exchanger and the oil-fuel heat exchanger. The capacity of the scavenge pumps in RG casing are approx 9 gals/min each (total 18 gals /min) and rear bearing scavenge pump capacity is 3 gals/min. RESTRICTED  RESTRICTED 36 Oil-Fuel Heat Exchanger 6. Oil-fuel heat exchanger is located in the oil tank. The prime function of the oil to fuel heat exchanger is to heat the fuel as required by fuel anti-ice system. Fuel flow through the oil-fuel heat exchanger is controlled by the solenoid valve. The solenoid valve closes by 10% speed switch and opens by 55% speed switch and remains open throughout the engine running condition. Oil-Air Heat Exchanger 7. Hot oil is cooled by oil to air heat exchanger. A temperature control bypass valve is also provided in the oil cooler to bypass the oil in case the oil temperature is too low. It is mounted on the top of the yoke assembly. Oil Filter 8. A replaceable filter element is provided on the aft face of the reduction gear case. If the filter is clogged, the filter bypass valve will open and allow the unfiltered oil to the engine. A tell tale indicator is provided on the RH side of reduction gear case to indicate the clogging of filter. Oil Vent Valve 9. The oil vent valve is mounted on the front face of the reduction gear section. The valve allows gear case air to enter the inlet of pressure pump and the two main scavenge pumps during starts. This unloads the pumps, reducing the load on the engine. This valve is opened by starting circuit and closed by 55% speed switch. RESTRICTED  RESTRICTED 37 Fig. 05-01 Lubrication System Schematic Bibliography: Garrett maintenance Manual Chapter 72 RESTRICTED  RESTRICTED 38 Intentionally left blank RESTRICTED  RESTRICTED 39 CHAPTER-6 ENGINE FUEL SYSTEM Introduction 1. The engine fuel system consists of: (a) Fuel pump assembly. (b) Fuel control unit. (c) Fuel shut off valve. (d) Manifold nozzle assembly. (e) Start/Enrichment fuel system. Fuel Pump Assembly 2. The engine driven fuel pump assembly consists of a centrifugal type boost pump and a gear type high pressure pump with a common drive train. The boost pump boosts inlet pressure and is fed to the high pressure pump. The pressure pump incorporates an anti- icing valve. The pump assembly delivers filtered fuel at the required pressure to the FCU. Fuel filter bypass valve is set to 13 PSI differential pressure. The high pressure discharge is limited by a relief valve set at about 1125±25 PSI. Fuel Control Unit 3. The fuel control unit meters the fuel supplied by boost and high pressure pump to the engine combustion chamber. The FCU consists of three major metering devices. (a) An under-speed fuel governor (b) An over-speed governor (c) Main metering valve/Manual fuel valve (MFV) 4. Fuel flow range is established by two basic limiting devices: (a) A maximum fuel flow schedule is functionally controlled by mechanical means and compressor discharge pressure. This is compensated for by compressor inlet pressure and temperature variance. (b) A minimum fuel flow schedule is functionally controlled by mechanical means and compressor discharge pressure. (c) A manual adjustment for fuel specific gravity is externally located on the fuel control unit to provide compensation for different types of fuel. Under Speed Fuel Governor (USFG) 5. It is a flyweight type of under-speed governor. The under-speed fuel governor is functional at reduced engine speed in beta mode of operation. With power lever at FI or below, the engine speed varies between 70% and 97% engine RPM depending on the speed lever position. RESTRICTED  RESTRICTED 40 6. If the engine speed should decrease below the selected setting, the under-speed fuel governor increases the fuel flow to oppose the speed decrease. The minimum setting (70% engine RPM) provides a minimum speed during ground operation. The maximum speed setting (97% engine RPM) provides engine speed drop control during service operation. Over Speed Governor (OSG) 7. Over-speed limitation is 104 + 1% RPM. The over-speed governor is a safety device to control the engine speed in the event of propeller governor malfunction. Excess engine speed provides over-speed governor flyweight action, which reduces fuel flow to restrict further engine speed increase. Main Metering Valve (Manual Fuel Valve) 8. During beta mode operation the under-speed fuel governor controls the amount of fuel supplied to the engine. As power lever is moved forward towards take-off condition, the under-speed governor is over come, fixing the flyweights in allowing the metering valve to be opened to cause additional fuel to flow. From this point onwards there is no control of fuel by the under-speed governor. During all flight mode operation the fuel flow is controlled by main metering valve. Fuel Shut-Off Valve 9. A fuel shut-off valve is located downstream of the FCU on the left hand side of the engine forward of the fire wall adopter. The fuel shut off valve requires momentary electrical power to set the valve in either the “fuel ON” or "fuel OFF” position. 10. During the start cycle, at approximately 10% engine speed, the10% speed switch automatically energises the valve to the "fuel ON" position. To electrically energise the valve to the "fuel OFF” position the start/stop switch is placed in the stop position to release the mechanical lock but is not closed during a manual start by means of MAN- IGN switch. The valve can be closed manually by positioning the engine speed lever to the shut off position. During an engine shut down by means of the engine speed lever, the fuel shut off valve is manually closed before actuating the propeller feathering valve. When the engine speed lever is in the shut-off position, the fuel shut off valve cannot be opened electrically. The engine speed lever must be returned to the low position before the valve can be opened electrically. Fuel Flow Divider 11. The function of the fuel flow divider is to direct the metered fuel into the primary fuel manifold and nozzle assembly during engine start. As the engine accelerates, the differential pressure across the bellow increases. It causes the fuel flow divider secondary port to open and allowing the metered fuel to flow through both primary and secondary fuel manifolds and nozzle assemblies. 12. Two separate fuel manifold and nozzle assemblies are utilised: (a) Primary manifold assembly. (b) Secondary manifold assembly. RESTRICTED  RESTRICTED 41 Primary Manifold Assembly 13. The primary manifold assembly consists of five fuel nozzles connected through a manifold to the primary side of the fuel flow divider. The primary nozzles are spaced evenly around the outside of the turbine plenum and extended through the plenum into the combustion chamber. During the initial phases of an engine start, the fuel atomisation is provided by the primary nozzles. Secondary Manifold Assembly 14. The secondary manifold assembly consists of the ten fuel nozzles connected through the manifold to the secondary side of the fuel flow divider. The secondary nozzles are spaced around the rear of the turbine plenum and extended through the plenum into the combustion chamber. For all other phases of the engine operation fuel is provided by both the primary and the secondary manifolds and nozzles. Start/Enrichment Fuel System 15. The start/enrich fuel system consists of an electrical solenoid valve and a pressure regulator. The enrich button is located on the over head panel in the COCKPIT, the pressure regulator and solenoid valve are located on the right side of the reduction gear case. The enrich button must be manually actuated to add fuel to the starting sequence, pressure regulated fuel then by passes the fuel control unit and is routed through the fuel shut off valve to the flow divider to aid the initial engine combustion process. During start, release the start/enrich button after ITT registers and re-engage as necessary to assist smooth engine acceleration above 25% RPM. RESTRICTED  RESTRICTED 42 ENGINE FUEL SYS SCHEMATIC Fig. 06-01 Bibliography: Garrett maintenance Manual Chapter 72 RESTRICTED  RESTRICTED 43 CHAPTER-7 ENGINE ANTI-ICE SYSTEM Engine Anti-Ice System 1. It can be seen from this simplified schematic that the air is taken from the compressor discharge. When the pilot has activated the switch in the cockpit and opened the anti-ice valve, the hot air is made available to the inlet shield. The hot air flows between the shield and the inlet housing of the engine. It escapes through the holes back into the nacelle area and is discharged overboard. 2. As that hot air flows past the inlet housing of the engine, it heats the housing and prevents ice from forming. The top part of the inlet housing will be heated by the warm oil that is in the gearbox. The small line provides that warm air to the sensor to keep it from being clogged with ice. 3. This picture also shows air being permitted to flow into the D-duct section of the nacelle mounted to the inlet. Flowing through the duct and discharging back into the nacelle, it transfers heat to the aluminium skin preventing ice from forming. Fig. 07-01 Engine Anti-Icing System P2 Sensor Anti-Ice 4. The P2 T2 sensor is mounted on the right side of the compressor inlet duct. Section protruding into the inlet air passages includes the P2 sensor, which is a ram probe facing into the inlet stream. Hot bleed air from downstream of the anti-ice valve is plumbed to the fitting shown on the left. As the warm air flows to the discharge port, it warms the probe to prevent ice formation. The small amount of anti-ice air that discharges into the compressor inlet has no effect on the temperature of the air going into the engine. RESTRICTED  RESTRICTED 44 Pre-Cool and Flow Limiting 5. This schematic shows a portion of the ducting that provides engine bleed air to the air conditioning system, as well as some auxiliary systems. The hot compressor air is cooled when flowing through the heat exchanger. The flow-limiting venture is part of the shut off valve that will limit the flow taken into the system within the specifications of engine bleed. 6. Notice that the mounting pad on the plenum chamber for the ducting system includes an orifice plate with a gasket on either side. The hole in the orifice plate is precisely calibrated to limit the bleed from the engine within the specified tolerances, in the event that a massive leak should occur in the aircraft ducting this would prevent over bleeding the engine 7. When the inlet De-ice switch is activated in the cockpit, the anti-ice valve opens allowing the compressor discharge air to the inlet shield. The hot air flows between the shield and the inlet housing of the engine. It escapes through the holes back into the nacelle area and is discharged overboard. As the hot air flows past the inlet housing of the engine it heats the housing and prevents ice formation. The top part of the inlet housing will be heated by the warm oil that is in the gear box. The small line provides the warm air to the P2 sensor to keep it from being clogged with ice. It is a self testing system, when the switch is activated, there is a slight increase in ITT can be observed. Caution: For ground operations do not operate engine anti-ice if ambient temperature above 5° C for more than 10 seconds. 8. During the start cycle, the priority of the enrichment system is obvious. It is more important that available fuel is sent to the enrichment system for good acceleration than to be concerned with anti-icing the fuel system filter. This priority is simply accomplished by the solenoids operating in opposite directions. The lockout solenoid is a normally open solenoid. The enrichment solenoid is a normally closed solenoid. During the start phase, between 10% and 55% RPM, the lockout solenoid is energized and closes to prevent fuel flow through the oil-to-fuel heat exchanger and thereby fuel pressure loss. 9. Beyond 55% RPM, the solenoid is de-energized and the anti-ice lockout valve re- opens allowing fuel to be available through the fuel heater up to the anti-ice valve. When the temperature-sensing element of the anti-ice valve senses that the temperature of the fuel in the area of the filter is getting to a point less than 40⁰ F, the anti-ice valve would open. The warm fuel from the fuel heater would mix with the inlet fuel and would keep the filter from icing. P2 T2 Sensor 4. The P2 probe identified on the sensing element is a total pressure pickup. The air pressure sensed at this probe will be sent through plumbing connected to the fuel control unit. This pressure is available to a bellows that will position the three dimensional cam to make the adjustment necessary to the metered fuel. There is a supply of warm air from the compressor section made available to make sure that P2 probe does not ice up. 5. The T2 system is made up of a sealed tube assembly filled with an alcohol solution. Increasing air temperature flowing over the coil of the T2 sensor will cause the fluid to RESTRICTED  RESTRICTED 45 expand. The expansion of the bellows will adjust linkage assembly within the fuel control to make the necessary correction to the metered fuel. Fig. 07-01 Bibliography: Garrett maintenance Manual Chapter 72 RESTRICTED  RESTRICTED 46 Intentionally left blank RESTRICTED  RESTRICTED 47 CHAPTER-8 ENGINE LIMIT SYSTEM Torque and ITT Limiting System 1. The torque and ITT limiting system is factory preset to 100% torque and 923⁰ C, ITT. Whenever the ITT or torque exceeds the limit, a signal is sent to bypass valve which opens and bypasses the excess fuel to the inlet of the fuel pump. The system will be tested at partial power, by selection of a spring loaded test switch to either ITT or TORQUE position or decrease in fuel flow and ITT can be observed. The system may be switched off by the "ENG LIMIT" switch. During starting phase the torque and ITT limiting system is isolated automatically from 0 to 55% of engine RPM. Torque Temperature Limit System 2. With the limiter system turned on, the pilot may advance the power lever to make maximum torque for takeoff. The maximum torque will then be limited by the automatic system and he can direct more attention to the handling of the aircraft. The torque/temperature limit controller receives its input signals from the torque/temperature indicating systems. The tachometer generator is sending an engine speed signal to the limiter controller through speed switch and there is also a 28 volt source of DC power. The one output of the controller goes to the torque limiter assembly, which is a fuel bypass valve. When either the torque or temperature limit is reached, this bypass valve is driven to an open position sufficient to bypass enough fuel back to the fuel pump inlet to keep the engine from exceeding those limits. Fig. 08-01 Torque/ITT Limiting System RESTRICTED  RESTRICTED 48 Limiter Bypass Valve 3. Above artwork illustrates the major portions of the "Limiter Fuel Bypass Valve". The metered pressure on the way to the atomizers is connected to the right side of this valve. The discharge, on the bottom of the picture, returns to the pump inlet. The metering valve, shown in the centre, is a calibrate orifice closed by a torque motor operated flapper valve. When the valve receives a signal at the electrical connector from the torque/temperature limit controller, it will move the flapper valve away from the orifice sufficiently to allow the right amount of fuel to bypass back to the inlet side of the pump. It appears that this could cause a problem to the engine if this valve were to malfunction in the wide-open position. However, this valve has a failsafe feature.The metering valve is a calibrated orifice that will limit the maximum flow being bypassed to approximately 60 Lbs per hour when fully opened.There is a screen in the inlet line which needs periodic inspection at every 200 hours inspection. Bibliography: Garrett Maintenance Manual, Chapter 72 and Chapter 77 RESTRICTED  RESTRICTED 49 CHAPTER-9 PNEUMATIC SYSTEM Fuel Purge System 1. The fuel that remains in the flow divider and in the fuel manifold circuit during the shut down cycle is purged into combustion chamber by an air charge and burned. The purpose of the purge system is to minimise the ITT in the subsequent start and to prevent the carbon deposits on the nozzle tips. Components 2. (a) Filter (b) Check valve (c) Accumulator (d) Solenoid valve 3. The P3 air bled from the engine to provide pneumatic power for the other systems is taken from different ports within the plenum chamber. The fuel control P3 signal is taken from a bulkhead fitting generally located near the top of the plenum chamber. The fuel manifold purge system usually takes air from the same port that supplies the inlet anti-ice valve. 4. When we discussed flow divider operation, we identified the spool type drain valves that allowed fuel to drain from the manifolds out a drain port on the flow divider. In recent years, the Environmental Protection Agency has issued regulations prohibiting the discharging of fuel into atmosphere. The "FUEL" manifold purge system pictured here has been designed to utilize air pressure to push that fuel into the engine and burn it as the engine is being shut down. 5. The source of air pressure is plumbed from the engine through a filter and a one- way check valve into an air storage tank. Downstream of the tank is an electrically operated shutoff valve that can discharge that air down to a check valve attached to the drain port of the flow divider. During normal operation, the electric shutoff valve will be closed and air will pressurize the air storage tank. 6. The check valve on the flow divider drain port will be seated preventing the possibility of any internal fuel leakage from getting back into the pneumatic purge system. When the stop switch is completing the circuit to the fuel shutoff valve to energize it closed, it also energizes the control valve in the purge system to the open position. Since the pressure in the storage tank is higher than the pressure from the engine at reduced speed, the differential pressure closes the check valve downstream of the filter. 7. Thus the only way out for the air stored in the tank, is to discharge through the open control valve, unseat the check valve on the flow divider drain port and push the flow divider drain port and push the fuel that remains in the manifolds into the combustion section to be burned before the flame goes out. This will result in a slight increase in rpm, just after energizing the stop switch closed. That slight increase in rpm is an indication that the purge system is operating normally. RESTRICTED  RESTRICTED 50 8. It is important for the engine to run at least 95% rpm prior to shutdown so that the charge in the air storage tank will be high enough to satisfactorily purge the system. It is also advisable to hold the stop switch circuit to the control valve for at least five seconds, or until the engine rpm has dropped below 50%. This provides the time necessary to discharge the air storage supply through the flow divider and completely purge the manifolds. Fig. 09-01 Fuel Manifold Purge System Bibliography: Garrett Maintenance Manual, Chapter 72 RESTRICTED  RESTRICTED 51 CHAPTER-10 ENGINE STARTING SYSTEM General 1. The engine can be started on the ground or in the air either automatically or manually. For ground starting, the starter-generator drives the engine between 0% and 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. 2. 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 un-feathering pump. At 10% RPM, fuel and ignition are introduced automatically or manually in the same sequence as described above. During engine starting, fuel can be added manually to assist in starting of a cold engine by pressing the fuel ENRICH button. Abort All Starts 3. Abort the engine start immediately if any of the following conditions are observed:- (a) Propeller fails to rotate. (b) Start sequence does not begin. (c) Battery temperature warning light comes ON (71⁰C) (d) RPM does not reach 10% in 10 seconds. (e) ITT not rising within 10 seconds after 10% RPM. (f) No ITT rise until 20% RPM. (g) ITT approaches start limit rapidly to 1149⁰ C. (h) No oil pressure indicated by ground idle. (j) RPM stops increasing prior to reaching the normal setting. (k) Any unusual noise or vibration. (l) Engine instruments indicate abnormal conditions. RESTRICTED  RESTRICTED 52 Fig. 10-01 Start-Control Panel Start Selector Switch 4. The START selector switch located in the middle of the starting panel (Fig 10-01), 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. 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 un-feathering 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. RESTRICTED  RESTRICTED 53 START/STOP Switch 5. 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 un-feathering 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 6. Manual ignition is provided for the following reasons: (a) To enable manual AIR or GROUND starts. (b) To allow battery starts in the manual mode, when battery voltage is only 24 V. (c) To enable an engine start with the 10% and/or 55% speed switch defective. (d) To prevent flame-out under icing/rain conditions (e) To enable an engine start when residual ITT is more than 300 0 C. Manual Ignition Switch 7. 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 during ice and rain condition. (b) OFF: This position allows tenderization 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. Engine Starting Modes, Key Steps Of Operation 8. Note: Before starting engines, it is necessary to carry out a pre-flight inspection of the airplane, using the Pilot's Operating Handbook. RESTRICTED  RESTRICTED 54 9. Automatic Ground Start (a) Set START selector switch to GROUND. (b) Set respective START/STOP switch to START and holds momentarily. The start cycle is initiated and the following circuits are energised. (c) To the coil of the start relay and its hold-in circuit. (d) To the starter which cranks the engine. (e) To the oil vent valve which opens. (f) To the power supply relay of the torque/temperature limiting system, which opens and prevents operation of the system? (i) When engine RPM has reached 10%, the "10% contacts" in the speed switch assembly is made, energising the following circuits: (aa) To the ignition unit which fires the igniter plugs. (ab) To the two ignitions indicating lamps which illuminate. (ac) To the "open" circuit of the engine-side fuel shutoff valve which opens and allows fuel flow to the engine fuel nozzles. (ad) To the anti-ice lockout valve which closes and prevents fuel flow through the oil-fuel heat-exchanger. (ae) To the fuel enrichment switch which, if pressed, opens the fuel enrichment valve and allows additional fuel flow to the engine. (ii) When engine RPM reaches 55%, the "55% contacts" in the speed switch assembly open which interrupts power supply to: (aa) The starter circuits (switches over to "generator" mode). (ab) The ignition unit, stopping the firing of the igniter plugs. (ac) The ignition lamps which extinguish. (ad) The oil vent valve which closes. (ae) The "open" circuit coil of the high pressure fuel shutoff valve. However, the valve will remain open due to its internal Belleville spring washer mechanism. (af) The anti-ice lockout valve which re-opens. (ag) The power supply relay of the torque/temperature limiting system which closes and re-powers the system. RESTRICTED  RESTRICTED 55 Note: The engine limit switches must be ON for engine starting. 11. Manual Ground Start (a) Set START selector switch to VENT. (b) Set respective START/STOP switch to START and hold. The circuits energised are: (i) 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. (ii) To the starter which cranks the engine. (iii) To the oil vent valve which opens. (iv) 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. (g) When engine RPM has reached 10% to 15%, set the manual ignition switch (MAN IGN) to the START position. This will energise the components as outlined in the AUTOMATIC GROUND START procedure. (h) When engine RPM reaches 55%, release the START/STOP switch to its neutral position and set the manual ignition switch (MAN IGN) to the OFF position. (j) Releasing the START/STOP switch to its neutral position will De-energize: (i) The starter. (ii) The oil vent valve which closes. (iii) The power supply relay of the torque/temperature limiting system which closes and re-powers the system. (k) Setting the manual ignition switch to the OFF position will de-energize: (i) The ignition unit, stopping the firing of the igniter plugs. (ii) The ignition lamps which extinguish. (iii) The "open" circuit of the engine-side fuel shutoff valve. However, the valve will remain open due to its internal Belleville spring washer mechanism. (iv) The anti-ice lockout valve which re-opens. (v) The fuel enrichment switch. RESTRICTED  RESTRICTED 56 12. Automatic Air Start (a) Set START selector switch to AIR. (b) Set respective START/STOP switch to the START position and hold momentarily. The following circuits are energised:- (i) To the coil of the start relay and its hold-in circuit. (ii) To the oil vent valve which opens. (iii) To the power supply relay of the torque/temperature limiting system which opens and prevents operation of the system. (iv) To the un-feathering pump. Operation of the un-feathering pump brings the propeller out of the feathered position. Decreasing the propeller pitch angle will cause the propeller to windmill and crank the engine. Note: The negative torque sensing system will control the propeller pitch angle according to the increasing RPM to minimize drag. (c) 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" procedure. (d) When engine RPM has reached 55%, the "55% contacts" in the speed switch assembly open which interrupts power supply to:- (i) The un-feathering pump (ii) The system and components as outlined in the “AUTOMATIC GROUND START” except to the starter which had not been energized. 13. Manual Air Start Note: The START/STOP switch is not to be actuated to initiate a manual air start. (a) Set START selector switch to AIR. (b) Unlock the un-feathering pump switch, set and hold the switch in the ON position. Operation of the un-feathering 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. RESTRICTED  RESTRICTED 57 (c) When engine RPM has reached 10% initiate the following: (i) Release the un-feathering pump switch to the OFF position and close the guard lid. (ii) Set the manual ignition switch (MAN-IGN) to the START position. This will power the components as outlined in the automatic ground start. (d) 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). (e) When engine RPM reaches 55%, set the manual ignition switch to the OFF position. This will de-energize: (i) The ignition unit, stopping the firing of the igniter plugs. (ii) The ignition lamps which extinguish. (iii) The "open" circuit of the engine-side fuel shutoff valve. However, the valve will remain open due to its internal belleville spring mechanism. (iv) The anti-ice lockout valve which re-opens. (v) The fuel enrichment switch. Bibliography: Garrett Maintenance Manual Chapter 72, POH Section 02 RESTRICTED  RESTRICTED 58 Intentionally left blank RESTRICTED  RESTRICTED 59 CHAPTER-11 EMERGENCY HANDLING PROCEDURES Operational Guidelines 1. Adherence to the following aircraft handling practices like, careful engine operations within limitation can enhance performance, improve engine life and reduce cost of maintenance:- (a) Do not exceed flight manual (POH) prescribed limits. (b) Rotational freedom and unusual sound from engine to be checked by hand cranking of propeller (DOR) prior to every start. (c) If a start attempt was aborted due to no combustion or excessive ITT, a clearing or engine ventilation run is recommended prior to another start attempt. (d) Carefully observe ITT and RPM rate of rise and limitation during engine starts. (e) If beta light goes out during start lock removal, hesitate momentarily until it re- illuminates. (f) Time in excess of limit RPM, torque and temperature can reduce the life of bearings and hot section components. (g) Hand rotation of the engine (DOR) limits peak post shut down engine temperature and will enhance fuel nozzle life. (h) Complete post flight inspection which includes looking for smooth engine rotation, oil by-pass indicator normal, rear turbine, tail pipe and propeller condition. (j) Install inlet and exhaust covers of engine only after 15 to 30 min cooling period. 2. Power Plant Instrument Markings INSTRUMENT Red Yellow Green Arc Yellow Red Orange Line Arc Normal Arc Caution Line Triangle Caution Open. or take off Max Min Range Limit. Tachometer (% 70 to 101% - 96 to 101 - 104% (a) RPM) 96 ITTo C - - 300 to 923 885 to 923(c) 923 1149 (b) Oil Temp (o C) -40 -40 to 55 to 110 110 to127 127 - 55 Oil Pressure (PSI) 40 to 120 40 70 to 120 - - 70 Torque (% TQ) - - 0 to 101.4 - 101.4 - RESTRICTED  RESTRICTED 60 3. Starting Limitations (a) Air-Starting Limitations (i) Maximum altitude for air starts is 20,000 ft (only applicable if certified maximum operating altitude is above 15,000 ft). (ii) Minimum airspeed for air starts is 92 KIAS (iii) Minimum oil temperature for air starts if 4o C (39o F) (b) Starter Operation Limitations (i) First cycle - max 60 sec ON – minimum 60 sec OFF (ii) Second cycle - max 60 sec ON - minimum 5 min OFF (iii) Third cycle - max 60 sec ON - minimum 1 hr OFF 4. Power Plant Limitations The nominal (flat) rated power is 715 SHP (533 KW) which is available at 33°C (91⁰F) at sea level or up to 7300 feet at ISA temperatures. Power Plant Limitations Condition Torque Maximum RPM % Oil Pr. PSI Oil Temp Time % 0 oC (d) ITT C (a) Maximum 101.4 923 101max (b) 70 to 120 55 to 110 Continuous Power (c) 110 to 127 5 minutes - - 96 to 100 70 to 120 55 to110 Continuous Flight Idle 110 to 127 5 minutes - - 67 min (g) 40 to 120 -40 to 127 Continuous Ground idle (h) Maximum - - Min 93 70 to 120 55 to 110 Continuous reverse 110 to 127 5 minutes (landing) Maximum 50 - min 94.5 70 to 120 55 to 110 Continuous reverse 110 to 127 5 minutes (static) - 1149 - - 1 second Starting -40 to 127 Continuous - - 00-05 - - Continuous Wind 05-10 30 minutes milling(5) 10-18 05 minutes 18-28(f) - 28-100 1 minute - - - - 55 to 110 Continuous Ground 110 to 127 (below Operation 100% TQ) RESTRICTED  RESTRICTED 61 5. Legends (a) 8850C is maximum ITT for RPM below 99% except during starting. (b) Avoid operation between 18 and 28% RPM except for transients occurring during engine start and shutdown. Observe following RPM limitations:- (i) 96% RPM is minimum during flight. (ii) 101% RPM is maximum continuous for normal operation. (iii) 101.0-101.4% RPM - 5 minutes. (iv) 101.5-105.5% RPM - 30 seconds. (v) 105.6-105.9% RPM - 5 seconds. (vi) 106.0% RPM - never exceed. (c) 101.4 % torque is maximum for normal operation. If this torque limit has been exceeded refer to Aircraft Maintenance Manual. (d) Minimum oil temperature for: (i) Starting and ground operation is -40C (-400F). When temperature is below, preheat oil prior to starting. (ii) Take off and flight is 550 C (1300 F). Operate engine on ground and in flight on a regime at which temperature limits can be maintained. (iii) Air-start is 40C (390 F). (e) Reverse rotation not permitted. (f) Do not allow wind-milling in this speed range longer than time required to transition through this speed range. (g) Avoid sustained engine RPM at or below 67% during ground operation (Propeller limits). (h) Transients below 40 PSI are permissible at Ground Idle. 6. Engine Shut Down Note: Allow engine to operate for a minimum of three minutes with the Power levers in GI position before shut down, to ensure proper ITT stabilization. Caution: When the starter switch is set to stop, a slight rise in RPM and ITT followed by an immediate drop in fuel flow should occur. If the immediate drop does not occur, immediately position the engine speed lever to shut off to prevent the possibility of an engine compartment fire. RESTRICTED  RESTRICTED 62 7. Fuel Limitation Difference in fuel quantity shall not be more than 700 lbs between the LH &RH wing fuel system. Use 10% of aileron trim into the light wing for every 100 lbs of fuel unbalance. 8. Oil Limitations (a) Minimum contents per tank before takeoff - 4.7 Litres (b) Maximum capacity per tank - 5.9 Litres (c) Maximum usable oil per tank - 5.0 Litres (d) Maximum oil Consumption per engine - 01 Litre per 13.2 hrs 9. Propeller Limitation (a) Propeller de-icing must not be switched on as long as the respective propeller is stationary to prevent overheating of the heating mats. (b) All propeller blades of an individual airplane must be of the same models. (c) For known icing condition operation only with suffix- “B” are allowed. 10. Engine Air Inlet Anti-Icing Maximum ground operating time is 10 seconds if OAT is above 5 degree centigrade. 11. Engine Fire on Ground Following actions must be taken in the event of engine fire on ground: (a) FIRE COCK switch - Close Note: It may take up to 30 seconds until the engine flames out. Engine operation will continue until all fuel in the lines between fire wall and engine is burned. (b) BLEEDAIR switch - OFF After the engine has flamed out: (c) POWER lever -REVERSE to engage start lock (at 60 to 30% RPM). (d) START selector switch - VENT (e) Starter switch - START and hold to 17 % or 20 sec Note: Maintenance action is required before next flight. RESTRICTED  RESTRICTED 63 If fire warning persists: (f) Fire handle -Pull and turn (g) MASTER switch -OFF (h) ENGINE SPEED levers – SHUTOFF (j) Airplane –Evacuate 12. Cabin Fire on Ground (a) Engines -Normal shutdown (b) MASTER switch – OFF (c) Portable fire extinguisher(s) - Fight fire (d) Airplane –Evacuate 13. Engine Fire in Flight (a) ENGINE SPEED lever - SHUTOFF and FEATHER (b) FIRE COCK switch –Close (c) BLEEDAIR switch –OFF (d) AIRCOND sel. switch - RAM AIR In case of engine fire (e) Fire handle -Pull and turn 14. Cabin Fire in Flight (a) Portable fire extinguisher(s) - Fight fire (b) Land - ASAP Bibliography: Garrett Maintenance Manual Chapter 72, POH Section 02, 03 RESTRICTED  RESTRICTED 64 Intentionally left blank RESTRICTED  RESTRICTED 65 CHAPTER-12 INTRODUCTION TO GLASS COCKPIT Introduction 1. The EFIS or Glass cockpit system is a complete flight and navigation instrumentation system that intuitively provides information to a pilot/co-pilot via computer generated screen displays. 2. Visual Difference between Conventional Cockpit and Glass Cockpit Fig. 12-01 Conventional Cockpit View Fig. 12-02 Glass Cockpit View RESTRICTED  RESTRICTED 66 LRUs of EFIS (Electronic Flight Instrument System) DESCRIPTION QPA LOCATION Integrated Display Unit (IDU- 04 MIP 680P) Slip Indicator 02 PFD USB Cover Blank 02 MFD Analog Interface Unit (AIU) 01 RH Avionic Rack Data Acquisition Unit (DAU) 02 14/15 LH & RH WRM 04 All IDUs AEM 04 All IDUs Remote Bug Panel (RBP) 02 MIP Serial A/D converter (SAND) 02 RH Avionic Rack EFIS XFILL Switch 01 CPP TAWS Inhibit Switch 01 CPP TAWS G/S cancel Switch 01 CPP Cooling Fan 02 Front Baggage Compt. Integrated GPS 02 PFD Pilot and Co-Pilot Pressure Transducer Unit 01 At Frame No. (Hydraulic System) 11 RESTRICTED  RESTRICTED 67 Integrated Display Unit 3. The two outboard IDUs are configured as Primary Flight Display (PFD pilot/co- pilot) (CPU #1) and the two inboard displays as Multi Function Display (MFD pilot/co- pilot) (CPU #2) by system limit settings (screen position setting) in IDUs during installation. PFD pilot and co-pilot (CPU #1) has a primary flight information (PFI) page on the top area and a pilot-selectable multi-function page on the bottom area. The MFD pilot and co-pilot (CPU #2) in normal mode has pilot-selectable multifunction pages on both top and bottom areas. 4. The display pages available on IDUs based on configuration are Primary Flight Information, Map Page, HSI Page, NAV LOG Page, Traffic, Wx RDR, AUDIO/RADIO Page, Full/Half Engine Page and CAS Panel. Hardware and software are identical for all IDUs, and functionality is determined by system limit settings in IDUs during installation. The IDUs are independently connected to all external sensors and independently perform all integrated functions. 5. Personality module inside J1 connector contains the CPU number (#1 PFD and #2 MFD) and system designation (Pilot or co-pilot).Aircraft configurations are initially read from flash drive storage to provide IDU switch a default configuration setup in the event of personality module failure. Pilot IDU #1 reads aircraft configuration from its personality module. IDU #1(PFD Pilot) transmits this configuration to the other IDUs. The other IDUs save the transmitted configurations to flash drive storage. RESTRICTED  RESTRICTED 68 Fig. 12-03 Location of Components RESTRICTED  RESTRICTED 69 Fig. 12-04 Indication and Recording System – Equipment Location RESTRICTED  RESTRICTED 70 Fig. 12-05 Glass Cockpit CBs on 13VE CB Panel Fig. 12-06 Glass Cockpit CBs on 5VE CB Panel RESTRICTED  RESTRICTED 71 Fig. 12-07 IDU-680P Input Identification  Bezel key L1-L8 Left side  Bezel key R1-R8 Right side  Encoder 1  Encoder 2  Encoder 3  Encoder 4  Turn and slip indicator (TSI) RESTRICTED  RESTRICTED 72 Fig. 12-08 PFD Normal Mode Pages RESTRICTED  RESTRICTED 73 Fig. 12-09 MFD Normal Mode Pages Bibliography: Airplane Maintenance Manual Glass Cockpit System (H-9501) RESTRICTED  RESTRICTED 74 Intentionally Left Blank RESTRICTED  RESTRICTED 75 CHAPTER-13 ENGINE INDICATING AND CREW ALERTING SYSTEM (EICAS) AND DATA ACQUISITION UNIT (DAU) 1. Engine Indicating and Crew Alerting System (EICAS) is a dual redundant system that can be manually selected by the pilots through the Electronic Flight information System (EFIS) menu in case of a failure event. 2. Data Acquisition Unit (DAU) is a microprocessor-based signal conditioning system. The system will accept aircraft sensor data, perform signal conditioning and validation and then present this information to compatible systems via ARINC 429 (Aeronautical Radio Incorporation) compatible serial data bus. 3. The DAU configuration is designed around dual channel architecture. Two independent processing channels (DAU Channel #1 and DAU Channel #2) interface with engine and aircraft sensors. Both the channels are capable of processing and conditioning the sensor data and output the sensor data. Operation 4. System The two Data Acquisition Units (DAUs) #1 and 2 collects sensor data (e.g. torque, ITT, fuel flow, RPM, Oil pressure and temperature) and discrete warning (e.g. IGN and Beta light, oil pressure, fuel pressure etc.) from LH and RH engines respectively. DAUs also collect other aircraft data from various sensors like hydraulic pressure, fuel quantity, stabilizer trim, aileron trim, rudder trim, flaps, prop deice and their respective discrete warnings. The DAU perform signal conditioning and validation on analog sensor data and then present this information to EFIS via ARINC 429 serial data bus. 5. Channel #1 of both the DAUs 1 & 2 forms EICAS 1 display and Channel #2 of both the DAUs 1 & 2 forms EICAS 2 display for EFIS. At EFIS start-up, the system will default the Pilot side IDUs to display EICAS 1 data and the co-pilot side IDUs to display EICAS 2 data. 6. The DAU contains two independent power supplies Power-A and Power-B (one for each channel) that accept 10 to 32VDC. Power-A input (Channel #1) to both the DAUs is provided by DAU1-A and DAU2-A circuit barkers respectively on the overhead CB panel (5VE). DAU1-B and DAU2-B circuit barkers on the CB panel 13VE provide protection for Power-B input (Channel#2) of DAU#1 (LH) and DAU#2 (RH) respectively. 7. DAU at EFIS start-up, the system will default the Pilot side IDUs to display EICAS 1 (Channel #1) data and the co-pilot side IDUs to display EICAS 2 (Channel #2) data of both the DAUs. EICAS channel can be manually selected by the pilots through the Electronic Flight information System (EFIS) menu in case of a failure event. RESTRICTED  RESTRICTED 76 8. In the event of an EICAS channel failure, one of the following yellow caution alerts: “EICAS1 FAIL” or “EICAS2 FAIL”, CAS message displayed on EICAS display page with a caution tone in headset and triggering master CAUTION light. Perform the following procedure in the event of an EICAS failure: (a) Press MENU (R1): (b) Press SOURCE (L2): (c) Rotate BTM to scroll down the Source list and highlight EICAS without checkmark: or (d) Push the encoder to select the new EICAS source. (e) Rotate the encoder to highlight DONE and push the encoder to exit or press EXIT (R1). Note: When both Pilot side and co-pilot side IDUs are operating from the same EICAS source, the cyan “SAME EICAS” advisory CAS message is displayed on the CAS message on EICAS display page along with an advisory chime in headset. RESTRICTED  RESTRICTED 77 EICAS Miscompare Threshold 9. For each of the following EICAS parameters, if the difference between left and right data sources(Channel #1 and Channel#2 of DAU) exceeds the miscompare threshold for more than the delay time, an asterisk (*) will appear alongside the digital display of the EICAS parameter, as listed in the following table: Note: Miscompare situation is an awareness to the flight crew that the two data sources for parameter/s being displayed on EICAS page are not in agreement within defined threshold limits (as defined in table above). Fig. 13-01 Sample ITT Miscompare on EICAS Display Page Bibliography: Airplane Maintenance Manual Glass Cockpit System (H-9501) RESTRICTED  RESTRICTED 78 Intentionally Left Blank RESTRICTED  RESTRICTED 79 CHAPTER-14 DATA ACCUSITION UNIT’S OPERATION WITH OTHER SYSTEMS 1. Pressure Refueling Both DAU #1 and #2 receives 28 V DC discrete from refueling panel (installed at RH wing) for respective LH and RH wing tanks. If Master Refueling switch is in DOWN (REFUEL) position and PRE-CHEK Switches on refueling panel is ON, “REFUELING” message displayed on fuel quantity bars and only total fuel on LH and RH side tanks displayed. Press and hold fuel counter reset switch (7KX) to view fuel quantity of individual tank during refueling process. Fig. 14-01 Fuel Quantity Display 2. Fuel Dump DAU #1 (LH) receives 28 V DC discrete from both LH and RH fuel dump valve. If active, there is a yellow “L FUEL DUMP” and “R FUEL DUMP” CAS message on the EICAS display page with a caution tone in headset and triggering master CAUTION light. 3. Fuel Indicating Once in four second DAUs receive the fuel quantity of Outer/feeder/Total values from Fuel amplifier. Outer/feeder/Total tank selection signal (+13 V DC) to Fuel amplifier is provided via SAND unit (part of EFIS system). The values for LH and RH fuel quantity are transmitted to DAU #1 and #2 for respective sides to display fuel quantity on EICAS display page. The fuel quantity is shown on the EICAS display page graphically and additionally in numeric values. The total of LH and RH side and the total in complete are indicated as numeric value only. Fuel displayed on inner tank is calculated from total minus outer minus feeder tank. 4. Low level Warning Each fuel quantity minimum switch provides a GND discrete signal to DAU #1 and #2 for respective side, which activates the yellow “L FUEL QTY” and “R FUEL QTY”CAS message on EICAS display page with a caution tone in headset and triggering master CAUTION light. 5. Hydraulic Power Indicating At frame #11 a pressure sensor is installed for measuring the pressure in the hydraulic system. The DAU #1 (LH) provides an excitation of 28 V DC to pressure sensor and sensor transmits 0.5 -5.5 V DC output signal to the DAU #1 (LH). Hydraulic pressure indication from 0-300 BAR displayed graphically on the EICAS display page. From 180-230 BAR hydraulic pressure indication scale is green outside this range scale is white coloured. RESTRICTED  RESTRICTED 80 6. Ice and Rain Protection Propellers At frame 15 (Equipment rack) there is one current shunt (50AMP/50mV) for each propeller installed measuring the current consumption of the propeller de-icing system. The output current signal is transmitted from LH and RH propeller shunt to DAU 1 & 2 respectively. The current consumption of the LH and RH propeller de-icing system displayed numerically on the EICAS display in Ampere. Current numerical indication remains green between 26 to 28 Amp, outside this range indication become white. 7. Steering DAU #1 (LH) receives a GND discrete signal from the Nose Wheel Steering (NWS) CAM Switch (LH MLG Actuator) and a 28 V DC discrete signal from the NWS Steering control unit. Depending on the signals either the yellow “NWS BYPASS” or “NWS 45°” CAS message annunciated on EICAS display page with a caution tone in headset and triggering master CAUTION light. 8. Air Intakes Each inlet anti ice selector provides a 28 V DC discrete signal to DAU #1 and 2 for respective LH and RH engine. Under icing conditions, the DE-ICE INLET switch must be switched on, so that the ANTI-ICE VALVE opens allowing compressor hot air to be fed to the system. If a valve malfunctions, a micro-switch in the ANTI-ICE VALVE Activates the yellow “L INLET DE-ICE” and “R INLET DE-ICE” CAS message on EICAS display page with a caution tone in headset and triggering master CAUTION light. 9. Propeller Beta Indication Each beta pressure switch provides a GND discrete signal to DAU # 1 and 2 respectively for LH and RH engine, which activates the blue BETA light indication on EICAS display page. 10. Engine Fuel System Each fuel pressure switch fitted between the LP cocks and engine connection hoses provides a GND discrete signal to DAU #1 and #2 respectively for LH and RH side. If switch is closed, it activates the yellow “L FUEL PRESS” and “R FUEL PRESS” CAS message on EICAS display page with a caution tone in headset and triggering master CAUTION light. 11. Each differential min pressure switch provides a GND discrete signal to DAU #1 and #2 respectively for LH and RH engine, which activates the yellow “L FUEL FILT” and “R FUEL FILT” CAS message on EICAS display page with a caution tone in headset and triggering master CAUTION light. 12. Fuel Flow Fuel flow transmitter, installed on each engine provides both the flow rate and temperature value to the signal conditioner. Signal conditioner converts these signals into 0-5 V DC aux flow rate output to DAU # 1 and 2 for respective LH and RH engine to display Fuel Flow on EICAS display page. Fuel flow indication is available with a pointer and the respective value as digital number in LBS/HR. 13. Signal conditioner also provides combined total fuel consumed pulse signal to DAU # 1 (LH) for displaying cyan coloured numerical value of total fuel consumed in LBS on EICAS display page. Fuel counter of all the IDUs can be rest using fuel cntr reset switch installed on centre pedestal. Each IDU has individual counter for total fuel consumed, in case of IDU powered down/restarted, fuel counter of that IDU will get reset to Zero. RESTRICTED  RESTRICTED 81 14. Ignition and Starting Both DAU #1 and #2 receives 28 V DC discrete from respective LH and RH engine, when ignition is active. If ignition is active, there is a yellow “IGN” indication on the EICAS display page. 15. RPM Indication A tachogenerator, installed on each engine converts engine RPM into electrical pulse (whose frequency is proportional to RPM) and provide to DAU # 1 and 2 for respective LH and RH engine to display % RPM on EICAS display page labeled RPM %. RPM indication is available with a pointer and the respective value as digital number scaled from scale from 0 to 110%. 16. Scales are graduated in increments of 10 with subdivisions of 5, and with numerals at 0, 20, 40, 60, 80, and 100. In IDU ground mode, from 0- 96% RPM scale is yellow, while in air mode from 0 to 68% is white and 68.1 to 96% is yellow. From 96 to 101% the scale is green and from 101 to 106% the scale is red coloured. As indication enters in yellow or red range, it starts flashing. Press master WARNING or CAUTION light, to stop flashing of indication. Note: RPM