A380RR_71..80_B12X1

Airbus A380

RR RB211 Trent 900

ATA 71−80 Power Plant

EASA Part-66 B1/B2_(excluding−Level-1−Contents) A380_71−80_B12x1

Revision: 1OCT2010 Author: WzT For Training Purposes Only E LTT 2007

Training Manual For training purposes and internal use only. E Copyright by Lufthansa Technical Training (LTT). LTT is the owner of all rights to training documents and training software. Any use outside the training measures, especially reproduction and/or copying of training documents and software − also extracts there of − in any format all (photocopying, using electronic systems or with the aid of other methods) is prohibited. Passing on training material and training software to third parties for the purpose of reproduction and/or copying is prohibited without the express written consent of LTT. Copyright endorsements, trademarks or brands may not be removed. A tape or video recording of training courses or similar services is only permissible with the written consent of LTT. In other respects, legal requirements, especially under copyright and criminal law, apply. Lufthansa Technical Training Dept HAM US Lufthansa Base Hamburg Weg beim Jäger 193 22335 Hamburg Germany Tel: +49 (0)40 5070 2520 Fax: +49 (0)40 5070 4746 E-Mail: [email protected] www.Lufthansa-Technical-Training.com

Revision Identification: S The date given in the column ”Revision” on the face of this cover is binding for the complete Training Manual.

S Dates and author’s ID, which may be given at the base of the individual pages, are for information about the latest revision of that page(s) only.

S The LTT production process ensures that the Training Manual contains a complete set of all necessary pages in the latest finalized revision.

A380 71−80

ATA 71−80

ENGINE RR TRENT 900

FOR TRAINING PURPOSES ONLY!

Lufthansa Technical Training

ENGINE

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ATA 71 POWER PLANT TRENT 900 FOR THE AIRBUS A380−840 Rolls−Royce has developed the high thrust Trent family to meet the strong market demand for heavyweight, long range Aircraft, and its design exploits proven advance technology to provide a low−risk route to high power. The engine for the Airbus A380−840 is designated Trent 900. The Trent 900 benefits from the experience of the Trent 700 in the Airbus A330, the Trent 500 in the A340−500/600 and the Trent 800 in the Boeing 777. Reliability is ensured by the use of high technology components and keeping operating temperatures close to RB211 experience. The unique Rolls−Royce three−shaft configuration, a high bypass ratio and enhanced component efficiencies contribute to improved fuel consumption and overall efficiency.



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POWERPLANT EXTERNAL DIMENSIONS The diagram opposite shows the powerplant external dimensions in imperial and metric, it is the same for all thrust variants of the Trent 900. Ground Clearance S Inboard - 1.05m to 1.25m / 42in to 49.2in S Outboard - 1.90m to 2.27m / 74.4in to 90in Leading Particulars Take off thrust (S.L. Static)


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LP System N1 Indication − IP System N2 Indication − HP System − N3 Indication − Flat Rated − Temperature

Trent 970−84 − 78 304 lbs Trent 970B−84 − 75 152 lbs Trent 972−84 − 76 750 lbs Trent 972B−84 − 80 211 lbs Trent 977−84 − 83 835 lbs Trent 977B−84 − 80 780 lbs Trent 980−84 − 84 098 lbs Single Stage Fan 5 Stage Turbine 8 Stage Axial Flow Compressor Single Stage Turbine 6 Stage Axial Flow Compressor Single Stage Turbine ISA + 15 °C

By−pass ratio − Overall Pressure Ratio at Take−off

8.12:1 41.7:1

Powerplant length Powerplant diameter Fan Diameter Dressed Engine Weight

329in/8.36m 152.5in/3.87m 116in/2.95m 14 190lb/6 437kg

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Direction of rotation shafts: LP Counter−clockwise viewed from rear IP Counter−clockwise viewed from rear HP Clockwise viewed from rear Shaft speeds (100%) N1 = 2 900 rpm N2 = 8 300 rpm N3 = 12 200 rpm

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Figure 2 FRA US/T

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71 DANGER AREAS OF THE ENGINE WORKING AREA Engine Not Running Even if the engine is not running, the area is still dangerous and the personnel has to obey the precautions, which are given to operate an engine safely.

WARNING:

Engine Running To enable personnel safety when he has to act exceptionally on a running engine, the power level must be kept to the minimum necessary by setting throttle control levers to the IDLE position. The restricted areas are: S the intake suction area: in a radius of 4.5 m (15 ft), S the exhaust danger area: a corridor of 30_ from the exhaust nozzles to 70 m (230 ft) afterwards. To work on the engine safely, you must use the entry corridors located at the engine outboard side 1.3 m (4 ft) aft of the air intake cowl.

KEEP ALL PERSONS OUT OF THE DANGER AREAS DURING ENGINE OPERATION. CLEAN THE RAMP IF THERE IS SNOW, ICE, WATER, OIL OR OTHER CONTAMINATION OR MOVE THE AIRCRAFT TO A LOCATION THAT IS CLEAN. MAKE SURE THAT ALL PERSONS ARE SAFE BEFORE YOU START THE ENGINE. MAKE SURE THE PERSONS IN THE COCKPIT CAN SPEAK TO ALL PERSONS NEAR THE DANGER AREA DURING ENGINE OPERATION. OBEY ALL OF THE GROUND SAFETY PRECAUTIONS FOR THE ENGINES. THE ENGINES CAN PULL PERSONS OR UNWANTED MATERIALS INTO THEM AND CAUSE SERIOUS INJURIES OR DAMAGE TO EQUIPMENT

To work on the inboard engines, the outboard engines must be shut off first. Human factor points:


NOTE:

WARNING:

BE CAREFUL WHEN YOU DO WORK ON THE ENGINE PARTS AFTER THE ENGINE IS SHUTDOWN. THE ENGINE PARTS CAN STAY HOT FOR ALMOST 1 HOUR.

WARNING:

UNDER NORMAL CONDITIONS, EXCEPT IN THE ASSISTED MANUAL START SEQUENCE, THERE IS NO NEED AND IT IS NOT ALLOWED TO PERFORM MAINTENANCE TASKS ON A RUNNING ENGINE.

WARNING:

DO NOT GO NEAR AN ENGINE THAT IS IN OPERATION ABOVE LOW IDLE. IF YOU DO, IT CAN CAUSE AN INJURY. GO NEAR AN ENGINE IN OPERATION THROUGH THE ENTRY CORRIDORS ONLY.

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8.9 m (29 ft)

4,5 m (15 ft) 1,3 m (4 ft 3 in)


70 m (230 ft)

30 TO 548.6 m (1800 ft) AFT OF EXHAUST NOZZLES

30 ° INTAKE SUCTION DANGER AREA MINIMUM IDLEWPOWER EXHAUST DANGER AREA

INTAKE SUCTION DANGER AREA MAX TAKE−OFF POWER

ENTRY CORRIDOR

EXHAUST DANGER AREA

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71 MAJOR UNITS The propulsion system is comprised of the following items: S Air inlet cowl S Left and Right fan cowl doors S Engine, associated fairings, front and rear mounts S Exhaust nozzle assembly including the Thrust Reverser S Pylon mounted − left and right thrust reverser halves (inboard engines) or Fan Exhaust Duct (outboard engines).



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71 ACCESS DOORS AND PANELS There are a number of access doors and panels around the engine to give access for maintenance and servicing.



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Lufthansa Technical Training FOR TRAINING PURPOSES ONLY!


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ENGINE COWLING DESCRIPTION Fan Cowl Opening The fan cowl doors can be opened for maintenance purposes on the engine. The unlatching sequence is carried out from the latch access panel located at the split line between the two fan cowl doors. Unlocking of the four latches is done in a defined sequence: L4 first, L1, L3 and L2 at the end. Once the fan cowl doors are unlocked, the opening is done from the fan cowl P/B control switches installed on the air intake cowl, at the RH and LH sides of the engine. The maintenance personnel must push and hold the UP switch until the fan cowl door has reached the desired position. The HORs (Hold Open Rods) are automatically locked. When a HOR is locked the green indicator is visible in the full open position. Then the maintenance personnel must push the DOWN switch to hold the cowl on the HORs. The fan cowl doors have two open positions: S intermediate position of 40 degrees, S full open position of 50 degrees. The fan cowl doors can be directly opened from zero to the full open position. NOTE:

There are two flag indicators to know the HOR state: − red indicator, unlocked between 0º and 40º positions, − No indicator, locked on 40º position and unlocked between 40º and 50º positions, − green indicator, locked at 50º position.

CAUTION:

MAKE SURE THAT THE WIND SPEED CONDITIONS ARE NOT MORE THAN 45 KNOTS.

CAUTION:

BEFORE YOU FULLY OPEN THE FAN COWLS, MAKE SURE THAT SLATS ARE RETRACTED AND THAT THEY CANNOT MOVE TO PREVENT FROM POSSIBLE INTERFERENCES.

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Fan Cowl Closing At the end of maintenance tasks on the engine, the fan cowl doors have to be closed to put the aircraft back into operation. First of all, the maintenance personnel must push the UP switch momentarily and operate the release lever on the HOR to manually unlock it. When the HOR is unlocked, the red indicator is visible. Then he has to push and hold the DOWN switch until the fan cowl door closes completely. The locking of the four latches is done in a defined sequence: S L2 first, S L3, S L1, S and L4 at the end. Once the latches are locked, the latch access panel has to be closed.

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71 FAN COWL OPEN SEQUENCE


FAN COWL CLOSE SEQUENCE

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MAINTENANCE Preservation of the Powerplant Cautions: CAUTION:

YOU MUST DO ALL THE APPLICABLE PRESERVATION PROCEDURES WHEN YOU PUT AN ENGINE INTO STORAGE. IF YOU DO NOT, CORROSION AND GENERAL DETERIORATION OF THE CORE ENGINE AND THE FUEL SYSTEM CAN OCCUR.

YOU MUST NOT KEEP THE ENGINE IN STORAGE FOR TOO LONG. THE TIMES GIVEN IN THIS PROCEDURE ARE THE MAXIMUM FOR WHICH THE ENGINE CAN BE PRESERVED. IF THE TIME THE ENGINE IS IN PRESERVATION IS TO BE EXTENDED, YOU MUST DO THE FULL PRESERVATION PROCEDURE AGAIN. IF THESES PROCEDURES ARE NOT FOLLOWED, DAMAGE TO ENGINE CAN OCCUR The preservation procedure protects the RR TRENT 900 against corrosion, liquid and debris entering the engine and atmospheric conditions during periods of storage and inactivity. The time during which the engine will be stored, and the climatic conditions of storage are shown in a chart. This chart also gives the preservation procedures, which must be done in different conditions and for the different storage times. Refer to the AMM (Aircraft Maintenance Manual) for specific storage requests. To find the applicable preservation procedure you have to: S find the climatic condition in which the power plant will be stored, S find the time during which the power plant will be stored, S compare this data with the chart and make the decision as to which preservation procedures must be done.


CAUTION:

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Before a power plant is put in storage, these basic procedures must also be done: S clean and examine the power plant, S make sure the power plant is dry, S clean the power plant if a fire extinguisher has been used on it. For powerplants stored on−wing, desiccant must be used for protection. According with the conditions and the time of storage the procedure can also composed of: S Preservation of the main line bearings, S Inhibit the engine fuel system, S Attach the transportation covers, S Remove the engine and install it in an MVP bag.

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71 ENGINE ATTACHMENT Description The engine is core mounted and attached to the Aircraft pylon by: S Front Mount S Thrust Links S Rear Mount The engine mounts transmit the engine loads and thrust to the Aircraft pylon.



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71 ENGINE MOUNTS Purpose The mounts support the weight of the engine and transmit loads to the Aircraft structure. Front Mount The engine front mount is installed on the top of the intermediate case and attaches to the Aircraft pylon with six tension bolts. The front mount transmits the following loads to the Aircraft pylon: S Vertical S Side Thrust Links The thrust links transmit the thrust from the intermediate case to the underside of the pylon just forward of the rear mount attachment.


Rear Mount The engine rear mount is installed on top of turbine exhaust case and attaches to the Aircraft pylon via a pylon adapter beam with four tension bolts. The rear mount transmits the following loads to the Aircraft pylon: S Vertical S Side S Torsion

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71 ENGINE DRAINS Description The drains system collects and discards unused fuel and other fluids that can leak from certain engine units and from certain engine areas. The drain system collects leakage from the following systems: S Fuel S Oil S Hydraulic


Fuel System Drains A drains tank is installed on the right side of the LP compressor case, just above the HMU, and it collects fuel from the fuel manifold when the engine is shut down on the ground. The contents of the drains tank are drawn back into the main fuel system during subsequent engine running via a self−consuming drains system, consisting of a float valve and an ejector valve, which is located in the base of the drains tank. A float within the tank prevents the ingress of air into the system when the level falls. Should the tank become full an overflow pipe carries surplus fuel to the drains mast. Drain lines take fuel from the following components to the drains mast: S Fuel pump mounting pad S Variable stator vane actuators (VSVA) S Fuel drains tank overflow

Other Drains There is a pipe from the lower splitter fairing in Zone 2, to allow drainage overboard in the event of leakage or water ingestion. This drain exits through a hole in the C−duct latch access panel between latches 1 & 2. The turbine case drain is provided to drain any residual fuel left in the turbine area following a wet crank or start attempt when the engine fails to light up. The drain pipe exits through a hole in the C−duct latch access panel just to the rear of latch 6. A duct incorporated within the interservice bifurcation panel provides drainage from Zone 3 through a hole in the C duct latch access panel between latches 1 and 2 There are also forward and rear pylon drains which drain fluid overboard through holes in the latch access panel just to the rear of latch 6.

Oil Drains Drain lines take oil from the following components to the drains mast: S Oil tank filler scupper S Air starter mounting pad S Variable Frequency Generator (VFG) Hydraulic System Drain lines take hydraulic fluid to the drains mast from the inboard and outboard hydraulic pump mounting pads and pump seal cavity.

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71 Drain System Leakage Rates To be sure that an engine operates correctly, the leakage rates at drain mast have to be monitored, checked and measured. The leakage rates for each system have to be within the acceptable limits specified by the engine manufacturer. If this is not the case, further troubleshooting is necessary to identify the source of the leak.



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71 DRAINS MAST AND BREATHER OUTLET The drains mast and breather outlet are attached to a bracket on the rear face of the external gearbox. The drains mast is on the split line between the two fan cowl doors. The breather outlet from the centrifugal breather and other drains are annotated on the drains mast.



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Drains Mast

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71 DRAINS TANK Purpose To prevent the formation of coking deposits within the fuel spray nozzle manifold drains system to give increased HMU and float valve/ejector valve reliability.


Drains Tank Location The drains tank is installed on a bracket on the lower right side of the fan case, between the Fuel Oil Heat Exchanger (FOHE) and the Hydro Mechanical Unit (HMU).

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71 DRAINS TANK OPERATION Unlike previous RB211 / Trent designs there is no dedicated drain line from the fuel spray nozzle manifold. When the HMU drains valve is opened, fuel is drained directly from the main HP fuel line. When the engine is shut down, or after failure to start on the ground, fuel is drained from the fuel manifold. As fuel flows into the tank air is released through the outlet tube. After a number of failed starts, the tank can become full of drained fuel; this fuel is then discharged through the outlet tube to the drains mast. During normal operation, fuel in the drains tank lifts the float valve and moves it to the open position. During engine starting LP fuel flows through the ejector, this will lower the fuel pressure in the ejector to less than that in the tank and the non−return valve opens. This allows fuel to be removed from the tank and routed to the inlet side of the LP pump. When the fuel in the tank falls to a certain level the float valve closes preventing air being introduced into the LP fuel supply.



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Drains Tank Operation 14|71 |L3

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71 PYLON ELECTRICAL DISCONNECTS There are 18 separate harness electrical connectors between the engine / nacelle mounted components and the pylon. The connectors are keyed to correctly align the connector with its mating receptacle and to prevent cross connection. The powerplant harnesses are colour coded by having braids of different colours, known as tracer colours. These are used to identify the harness and follow its route. They also assist in identifying the FADEC systems harnesses from those of other systems. The illustration opposite shows the harness numbers and the pylon connectors to which they attach. It also shows the units which are connected by each harness.



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POWER PLANT

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71 PYLON ELECTRICAL RECEPTACLES & CONNECTORS VFG Cable and Zone The illustrations below show the following electrical disconnects: The VFG power cables junction block on the upper left side of the fan case. The receptacles and harness connectors above the left side of the engine core.


Fan Case to Pylon The illustrations below show the receptacles and harness connectors above the left side of the fan case. There are two groups of connections in this area.

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ELECTRICAL CONNECTOR 5013VCA


VFG CABLE AND ZONE ELECTRICAL CONNECTION



FAN CASE TO PYLON ELECTRICAL CONNECTION

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ATA 72 ENGINE MAIN ROTATING ASSEMBLIES Description The three rotating assemblies comprise: S Low Pressure (LP) compressor (fan) connected by a shaft to a five−stage turbine. S Intermediate pressure (IP) compressor connected by a shaft to a single stage turbine. S High Pressure (HP) compressor connected by a shaft to a single stage turbine. Roller bearings and ball (location) bearings support each shaft. The external gearbox is driven from the HP shaft through an internal gearbox and an intermediate (step−aside) gearbox.



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POWER PLANT ENGINE

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ENGINE MAIN BEARING ARRANGEMENT The LP and IP rotor assemblies are each supported by three bearings. The HP rotor is supported by two bearings. Two types of bearings are used in this engine, ball bearings for shaft location and roller bearings providing shaft radial support whilst allowing axial thermal movement. The bearings are located in 4 bearing housings. The location bearings for all three shafts are positioned in the intermediate case module. The front bearing housing contains the LP compressor and IP compressor roller bearings. The Internal gearbox contains the three thrust or location ball bearing assemblies. The HP/IP Turbine bearing housing contains the HP turbine and IP turbine roller bearings. The Tail Bearing Housing (TBH) contains the LP turbine roller bearing.



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TRENT MODULAR BREAKDOWN The Trent engine consists of eight modules as follows: S Module 01 (31) − LP Compressor Rotor S Module 02 (32) − IP Compressor S Module 03 (33) − Intermediate Case S Module 04 (41) − HP System S Module 05 (51) − IP Turbine S Module 06 (61) − External Gearbox S Module 07 (34) − LP Compressor Case S Module 08 (52) − LP Turbine The numbers in parentheses are the ATA numbers relating to modules, as used in the Engine Manual. The fan blades are non−modular items but can be considered as part of module 01 (31). The modular construction gives several important benefits: S Decreased turn−round time for repair S Lower overall maintenance costs S Reduced spare engine holdings S Maximum life achieved from each module S Savings on transport costs S Ease of transport and storage S On−wing test capability after any module change The engine is completed by the addition of various non−modular items and systems e.g. fuel, oil etc. Modules 01, 02, 03, 04, 05 and 08 form the core engine module.

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LP COMPRESSOR Description The LP compressor consists of the fan disc and fan shaft. The fan blades and annulus fillers, are non−modular but considered to be included in this module. Fan Disc The fan disc is a titanium disc with axial ”dovetail” slots for blade fitment. Each blade is held in the disc with a shear key. The disc incorporates a drive arm that connects to the rotor shaft with a curvic coupling. The disc also incorporates annulus filler location lugs as integral features. LP Compressor Shaft The LP compressor (fan) shaft connects to the fan disc through a curvic coupling that provides accurate location. The coupling is secured by a ring of bolts, which thread into captive nuts on the LP compressor roller bearing inner race, which is secured to the shaft by an interference fit in addition to the bolts. The bearing race also incorporates the front bearing housing oil seal and a phonic wheel for measurement of LP speed. The shaft connects to the LP turbine shaft through a helical spline coupling. A failsafe shaft is fitted inside the LP compressor shaft and secured to the LP turbine shaft by a collar and nut.


LP Compressor Blades The 24 wide chord titanium fan blades incorporate an inner platform with a dovetail feature for location in the disc. The blades are retained axially in the disc by a shear key. Annulus Fillers There are 24 aluminium annulus fillers located between each fan blade, which provide an aerodynamic profile at the base of each blade. The annulus fillers are installed onto the fan disc lugs and the located by a dowel into the rear spinner rear flange.

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POWER PLANT ENGINE

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SPINNER ASSEMBLY Description The spinner assembly directs air into the hub of the fan and has three main parts: S Spinner S Fairing S Rear Spinner Spinner The air intake spinner is made of glass reinforced plastic (GRP) material. The spinner is painted with a white spiral marking (to indicate fan rotation in poor lighting conditions) and has a rubber tip to prevent ice buildup. The spinner attaches to the rear spinner with 18 bolts and is located on the rear spinner by 3 timing dowels. 9 of the attachment bolts secure 9 support brackets, which are located by 2 dowels on the spinner flange. There is a P−seal forward of the flange which seals against the inner surface of the fairing to prevent moisture ingress. The spinner weighs 10.52 Kg (23.2 lbs)


Fairing The fairing smoothes the airflow across the flange, located between the spinner and rear spinner assemblies. It is made of composite material and attached with 9 screws to the support brackets on the spinner flange. The fairing weighs 2.4 Kg (5.3 lbs) Rear Spinner The rear spinner attaches to the fan disc with a bolted rear flange. There is also a balance ring on the rear flange, which may contain balance bolts, which are used to balance the assembly during module build. The rear spinner weighs 21.32 Kg (47.0 lbs) On the outer surface, adjacent to the rear edge, is a circumferential ring of 60 counter sunk bolts positions. These contain either standard bolts or trim balance bolts. The trim balance bolts (one Part No.) are installed when the LP rotor requires balancing during service.

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FAN BLADE ASSEMBLY Description The LP compressor has 24 wide−chord, hollow, titanium fan blades, incorporating low speed swept fan aerodynamics for efficiency and noise. The assembly consists of the following parts: S fan blade S shear key S slider assembly S annulus filler The fan blades fit into dovetail slots in the LP compressor disc. Each blade is axially located by a shear key, which fits into a slot in the disc. A rubber strap on the base of the blade dovetail holds the shear key on the blade. A slider assembly fits in the dovetail slot at the end of each blade and ensures that the shear key is located in the slot in the disc. The annulus fillers provide an aerodynamic profile between adjacent fan blades. They are manufactured in aluminium and incorporate retention lugs, which mate with the disc lugs for location. They also incorporate a rubber strip on both sides, which abut the airfoil surface of the fan blade. Axial retention of the annulus fillers is provided by the rear spinner assembly, which locates each annulus filler by a dowel through the rear flange.



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POWER PLANT ENGINE

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Fan Blade

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IP COMPRESSOR Description The IP compressor module is an eight stage axial assembly consisting of four main sections: S Front bearing housing S The IP compressor stage 1 - 4 case S The IP compressor stage 5 - 8 case S The IP compressor rotor


Front Bearing Housing (FBH) The front bearing housing includes a hub, which locates the LP and IP compressor bearings and an oil sump, also the LP and IP shaft speed probes. Connected to the hub are the engine section stator vanes (ESS) or fixed inlet guide vanes. The vanes are welded together as one unit and there are lugs on the outer ring. These lugs are connected to the FOGV torsion ring to make the FBH/OGV joint. This FBH/OGV joint holds the LP compressor case to the core engine. The electrical cables, from the shaft speed probes, pass internally through the ESS vanes. Other vanes contain tubes to supply oil to and from the roller bearings. Behind the ESS vanes are the variable inlet guide vanes.

IP Compressor Rotor The IP compressor rotor is an assembly of eight titanium rotor discs, in between the discs of stages 1, 2 & 3 there are spacers that have interstage seal fins. The discs at stages 1 to 6 have axial dovetail slots into which the rotor blades are installed. Retaining plates and lock plates keep the blades in position. At stages 7 and 8 the blades are installed in circumferential dovetail slots. These blades are locked in position with nut and screw lock assemblies. The IP front stubshaft is attached to the stage 1 disc with bolts, the forward end of the stubshaft has a phonic wheel for IP speed measurement. The stage 6 disc incorporates a drive arm with a curvic coupling to which the rear stubshaft is attached. Splines in the stubshaft engage with splines on the IP turbine shaft.

IP Compressor Stage 1 - 4 Case The stage 1 to 4 case is connected to the FBH at the front and to the stage 5 to 8 case at the rear. The case is divided into two semi−circular titanium half cases. The stage 1 and 2 vanes are variable with spindles on the outer surface, which are connected by levers and unison rings to the VIGV/VSV operating mechanism. The stage 3 and 4 vanes are fixed and located in T slots around the inner circumference of the half cases. Between the stator vane positions on the inner surface there are abradable linings, located opposite to the rotor blade tracks. IP Compressor Stage 5 - 8 Case The IP compressor case is flanged and bolted to the rear of the stage 1 to 4 case and is made of steel and contains stages 5 to 8 of the compressor. The case is divided into two semi−circular half cases. The stage 5 to 8 vanes are made of nickel alloy and installed in T slots around the inner circumference of the half cases. The stage 8 stator vanes are also known as the IP compressor outlet guide vanes (OGV‘s).

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INTERMEDIATE CASE The intermediate case is one of the major structural parts of the engine and made from two titanium cylindrical casings, which are welded together. In the rear half, behind the weld, there are ten equally spaced radial struts, which support an inner structure. The IP and HP location bearings and the internal gearbox are attached to the inner structure. Two lugs on the rear case, above the radial struts, transmit engine thrust through struts to the airframe pylon. The front part of the intermediate case has a stronger area at the top, which includes lugs for the attachment of the front engine mount. Above and below the center−line there are symmetrical positions for the installation of the A frame struts. The two A frame struts on each side of the case align with the installation point on each side of the LP compressor case. Below the engine horizontal center line on the intermediate case, there are borescope access holes, which align with related holes in the compressor cases. The radial struts, which have an aerofoil shape, are hollow. Some of the vanes contain tubes, which supply oil to and from the internal gearbox. The external gearbox drive shaft, which transmits power to the External Gearbox (EGB), is in one of the struts. Other struts supply compressor air to cool the HP/IP and LPT bearing chambers and seal the EGB accessory mount pads. The front part of the intermediate case is installed around the Stage 5 to 8 case and is connected to a flange around the middle of the stage 1 to 4 case. The rear part of the intermediate case is installed around the front part of the HP compressor case. The rear of the intermediate case is connected to the combustion outer case. There is also a bayonet connection from an internal flange at the rear of the intermediate case to the HP compressor case. Inner and outer walls make an annulus, through which the air flows from the IP compressor to the HP compressor.



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HP SYSTEM Description The system comprises: S HP compressor S Combustion chamber and outer case S HP turbine

HP Turbine The HP turbine is a single stage disc connected to a mini disc to the rear of the HP compressor drum. On the rear of the disc there is a stubshaft, which is inertia bonded to the disc. The disc has fir tree roots into which fit the turbine blades. Adjacent to the casing rear flange is a turbine case cooling (TCC) air manifold.


HP Compressor The HP compressor rotor is a six−stage assembly. Stages 1 to 4 are made of heat resistant alloy discs welded together to form one drum. The stage 5 disc is also heat resistant alloy. The stage 6 disc and rear cone are made of heat resistant alloy and welded together. The first stage blades are made of titanium and installed in axial dovetail slots and are locked with retaining plates. Stages 2 to 6 are made of heat resistant alloy and installed in circumferential dovetail slots and locked with nuts and screws. The heat resistant alloy cone, which tapers rearwards is inertia bonded to the rear of the stage 6 disc. At the rear of this cone is a mini disc to which the HP turbine is connected. The HP compressor case is an assembly of six flanged, cylindrical casings bolted together. The flanged joints are also the location for the rotor path abradable linings. There are slots in this assembly for the installation of the stator vanes. The stage 6 stator vanes are also the HP compressor outlet guide vanes (OGVs). These are installed at the entrance of the combustion chamber inner case. Combustion Chamber and Outer Case The outer case is flanged and bolted to the rear of the intermediate case and to the front of the IP turbine module. There are 20 openings through which the fuel spray nozzles are installed. There are also two igniter plugs installed through bosses in the combustion outer case. The combustion chamber is fully annular and consists of a tiled liner that is located inside the combustion chamber inner case. At the front of the inner case are the HP compressor outlet guide vanes (OGVs) and at its rear are the HP turbine nozzle guide vanes (NGVs).

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IP TURBINE Description The IP turbine case houses the IP turbine and IP NGVs, LP turbine stage 1 NGVs and the HP/IP bearing housing. The front flange bolts to the combustion outer case and the rear flange bolts to the front flange of the LP turbine module (52). The IP turbine NGVs are hollow. In alternate NGVs there is a strut that is attached to the turbine case by a bolt. The inner end of each strut is connected to the structure that holds the HP/IP bearing support assembly. Through some of the other NGVs are tubes to supply oil to and from the bearings and IP 8 cooling air to cool the housing. The IP turbine is a single stage turbine assembly. At the hub of the disc a drive arm extends rearwards, which connects to the IP turbine shaft and stub shaft using taper bolts The IP turbine shaft runs forward and is connected to the IP compressor stub shaft with helical splines. The IP stubshaft runs forward to engage with the IP turbine roller bearing. The disc has fir tree roots into which fit the turbine blades. Adjacent to the rear flange is a turbine case cooling (TCC) air manifold and location bosses for fourteen thermocouples. To the rear of the turbine blades are the LP1 NGVs.



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LP TURBINE Description The LP turbine has five discs which are bolted together to form a drum. The stage 4 disc acts as the drive arm and attaches to the turbine shaft with a curvic coupling. Also attached to the drive arm on the rear face is a stub shaft that connects the LP turbine to the LP roller bearing in the tail bearing housing to provide radial support. The stub shaft also connects to a phonic wheel shaft assembly for LP turbine shaft speed measurement. The discs have fir tree roots into which fit the turbine blades. The LP turbine case is a one−piece cylinder flanged and bolted between the IP turbine case at the front, and the exhaust outer case at the rear. Around the case is a cooling duct through which cooling air flows. On the inner surface between the NGV locations there are seal segments which touch the turbine blade shrouds. In front of each stage of turbine blades there is a stage of NGVs. The first stage of NGVs, which are hollow, are installed as 3 vane sets in the outlet from the IP turbine case. One vane in fourteen of the sets contains an EGT thermocouple and one set includes an overheat detector and one set includes a borescope access hole. Stages 2, 3, 4 and 5 NGVs are hollow and are installed in the LP turbine case. At the inner ends of the NGVs are honeycomb liners, which touch the fins of the interstage seals between the rotor discs. The LP turbine shaft goes forward through the center of the IP shaft and connects with the LP compressor shaft with splines. The tail bearing housing support structure includes a hub that is held concentric in an outer case by 14 radial hollow vanes. Some of the vanes contain tubes that supply oil to and from the bearing housing. There is also a supply of IP 8 air to cool and seal the bearing. One of the vanes has a pressure inlet in the leading edge to measure LP turbine outlet pressure (P50). LP turbine outlet pressure is used for health monitoring. The front flange of the case is attached with bolts to the rear flange of the LP turbine case. At the rear flange to the primary exhaust nozzle around the case are two flanges to increase the strength. Attached to these flanges, at the top, is the rear mount.

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EXTERNAL GEARBOX Description The external gearbox is a one−piece aluminium gearcase. It is installed on the lower part of the LP compressor case. The gearbox assembly transmits power from the engine to provide drives for the accessories mounted on the gearbox front and rear faces. During engine starting the gearbox also transmits power from the air starter motor to the engine. The gearbox also provides a means of hand turning the HP rotor system for maintenance purposes. The gearbox is driven from the HP rotor via a transmission system, consisting of an Intermediate gearbox (step−aside gearbox), an external gearbox drive shaft (radial drive) and lower bevel gearbox. The drive shafts for the installed accessories are sealed by non−contact air blown labyrinth seals fed with IP8 air. All the accessory interfaces are protected by a drains system. Components Installed on the Front Face S Dedicated Alternator S Air Starter Motor S Hand turning point S 2 Hydraulic Pumps Components Installed on the Rear Face S Variable Frequency Generator (VFG) S Lower bevel gearbox S Oil Pumps S Centrifugal Breather S LP/HP Fuel Pumps S Hydro−Mechanical Unit (HMU)

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LP COMPRESSOR CASE Description The LP compressor casing assembly consists of three main sections: S Front Fan Case S Rear Fan Case S Fan Outlet Guide Vanes


Front Fan Case The containment case (front) and the center case are manufactured from titanium and are welded together to form the front fan case. The containment case has circumferential stiffening ribs (3 off), which provide reinforcement in the fan track region where additional energy absorption is required in the event of an LP compressor blade release. The front case has the following linings attached to the inner surface: S Acoustic panels (4) S Attrition lining S Ice impact area S Acoustic perforate skin Rear Fan Case The rear fan case is made from a titanium honeycomb structure. Two titanium supports (A frames), located on the horizontal centerline, connect the rear case to the core engine. On the rear outer edge of the case, there is a “V“ groove, which provides axial location of the thrust reverser. There is an opening in the left side of the case for the Variable Frequency Generator (VFG) Air Cooled Oil Cooler. There is also a large opening at BDC for the external gearbox drive shaft. Fan Outlet Guide Vanes (OGV‘s) The OGV outer ring is attached at the rear of the front case with bolts. The 52 OGV‘s are hollow titanium vanes filled with blue filler. The vanes are installed at equal distance around the circumference and the inner ends are welded to an inner ring.

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ENGINE CORE FAIRINGS Description To ensure a smooth airflow over the parts of the gas generator not covered by the thrust reverser halves, six removable fairings are fitted around the front part of the IP compressor case. Each fairing panels are a sandwich construction of titanium inner skin and perforated titanium outer skin with a nomex honeycomb core. The outer skin is perforated for noise attenuation. Two ventilation inlet holes are provided, one in each of the upper panels and two ventilation outlet holes, one in each of the lower panels. The front edge of each fairing is attached to the LP compressor OGV torsion ring with bolts secured in floating anchor nuts. The rear edge is attached to mounting brackets on the rear support diaphragm with bolts secured in floating anchor nuts. UPPER SPLITTER FAIRING Purpose To smooth the fan airflow into the thrust reverser halves and to provide a position for the fan air pressure rake (P160).


Description The upper splitter fairing is a carbon and glass composite fairing installed between the fan case and the intermediate case support structure. The P160 probe rake is installed inside the fairing with its six measuring heads projecting into the fan duct through holes in the leading edge of the fairing. LOWER SPLITTER FAIRING Purpose To smooth the fan airflow around the external gearbox driveshaft (radial drive) into the thrust reverser halves. Description The lower splitter fairing is a carbon and glass composite fairing installed forward of the external gearbox driveshaft assembly, between the fan case and the intermediate case support structure.

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FAN BLADE CLEANING Purpose To maintain the efficiency of the fan it is necessary to clean the fan blades and fan outlet guide vanes (OGV s) at regular intervals.


Description The procedure is fully described in the AMM 72−00−00 and is briefly described below: Follow all applicable Warnings and Cautions. Note: Depending upon the outside air temperature the washing fluid is a mixture of demineralized water, washing fluid (OM−1070) and monopropylene glycol (OM - 1076). Follow the AMM procedure for the applicable ratios. S Use a clean lint−free cloth soaked in the cleaning solution to clean the LP compressor blades. Makesure you apply the cleaning solution to the front andthe aft of the blades, and that the blade to becleaned is at bottom dead center. S Let the cleaning solution stay on the surface of theblades for 15 minutes. S Use a clean lint−free cloth soaked in demineralized or distilled water to remove the cleaning solution fromthe surface of the blades S Examine the blades for dirty areas S If they are not sufficiently clean, repeat the cleaning procedure again S Repeat this process for the fan OGV s. NOTE: 1. It is important that the fan blades are cleaned at bottom dead center to avoid any dirt migrating into the blade dovetail root area. 2. Most the dirt tends to stay on the suction face (rear) of the fan blade and particular attention should be given to this area. 3. Mix the washing fluid at regular 30 minute intervals

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INSPECTION OF LPC BLADE & ANNULUS FILLERS (AMM 72−31−41) WARNING:

Annulus Filler Inspection: Examine the annulus fillers for the following: S Cracks S Bends S Distortion S Nicks S Scores S Dents S Missing or split air seals If the annulus fillers are removed then the hooks and ribs should also be checked for nicks and dents. Cracks, bends and distortion are not allowed. Refer to the AMM for all other damage limits.

YOU MUST MAKE SURE THAT THE APPLICABLE COVERS ARE INSTALLED TO THE REAR OF THE ENGINE. THE MOVEMENT OF AIR THROUGH THE ENGINE CAN CAUSE THE LP COMPRESSOR TO TURN VERY QUICKLY AND CAUSE INJURY.

Preparation: Before carrying out the Inspection carry out the following: S Put a suitable access platform in a safe position S Put a protective rug into the air inlet cowl. (Make sure the red warning flagcan be seen externally of the air intake). S Install the Immobiliser - LP compressor rotor to prevent movement


Fan Blade Inspection The blade airfoil surfaces should be inspected for the following types of damage: S Cracks S Blade tip & adjacent airfoil surface heat discolouration S Arc−burns S Scratches & dents S Nicks S Blade bends

NOTE:

Annulus fillers that are rejected should be replaced with components that are the same weight or almost the same weight.

Cracks and arc−burns are not permitted and the affected blades must be replaced. Refer to the Aircraft Maintenance Manual limits for all other damage. NOTE:

NOTE:

The blade is divided into separate areas with different limits for each.

NOTE:

In addition to the normal limits for blade bends, there are fly on limits − the blade must be replaced within 125 hours or 25 flights (whichever occurs first).

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Figure 32 FRA US/T

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REMOVAL /INSTALLATION OF THE SPINNER & FAIRING (AMM 72−35−41) WARNING:

YOU MUST MAKE SURE THAT THE APPLICABLE COVERS ARE INSTALLED TO THE REAR OF THE ENGINE. THE MOVEMENT OF AIR THROUGH THE ENGINE CAN CAUSE THE LP COMPRESSOR TO TURN VERY QUICKLY AND CAUSE INJURY.

Preparation: S Put a suitable access platform in a safe position S Put a protective rug into the air inlet cowl. (Make sure the red warning flag can be seen externally of the air intake). S Install the Immobiliser (HU44211) - LP compressor rotor to prevent movement

Installation Procedure The installation procedure is the reverse of the removal procedure but you must make sure of the following points. 1. Align the timing pin on the spinner with the hole on the rear spinner 2. Torque all bolts to the value stated in the AMM. 3. Make sure all equipment is removed and the aircraft is put back to the correct configuration.

Removal Procedure: The component weights are as follows: fairing 2.40 Kg (5.3 lb) spinner 10.52 Kg (23.21 lb) 4. Using a temporary marker make an alignment mark across the fairing, spinner, rear spinner and annulus filler 5. Remove the attaching screws and remove the fairing. 6. Remove the bolts and brackets securing the spinner 7. Install the guide pins (HU44265) Make sure the groove points up (this is to catch the spinner when it is released from the support ring). 8. Install four of the removed bolts in the four extraction bushes and turn the four bolts in equal increments to release the spinner 9. Carefully remove the spinner from the guide pins 10.Put the spinner rear edge down on to an applicable flat surface.


NOTE:

CAUTION:

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Figure 33 FRA US/T

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REMOVAL /INSTALLATION OF THE REAR SPINNER (72−35−41) Removal Procedure: NOTE: The spinner weights 21.32 Kg (47.0 lb) Make a record of the positions of any compensation balance weights that are installed on the balance flange Install the lifting tool handles (HU44445) on the front flange of the rear spinner Hold the rear spinner and remove the attaching bolts and washers Install the guide pins (HU44265), making sure the groove points up. Install four of the removed bolts in the four extraction bushes and turn the four bolts in equal increments to release the rear spinner Remove the rear spinner from the guide pins Put the rear spinner rear edge down on to an applicable flat surface


Installation Procedure The installation procedure is the reverse of the removal procedure but you must make sure of the following points. Install the lifting tool handles (HU44445) on to the front flange of the rear spinner Align the timing pin on the rear spinner with the timing pin hole in the LP compressor disc Torque all bolts to the value stated in the AMM.

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REMOVAL / INSTALLATION OF THE ANNULUS FILLER (72−31−41) Removal Procedure 4. Using a temporary marker identify the location of each fan blade and each annulus filler 5. To remove the annulus filler, pull the annulus fillers forward to disengage the hooks from the LP compressor disc, then turn the annulus filler in the direction of its curve to clear the blades 6. Remove the two annulus fillers on each side of the blade to be removed.

Installation Procedure 1. Make sure all grease and debris has been removed from the seals and mating blade aerofoil surfaces 2. Lubricate the rubber seals with 1 part compressor washing fluid (OMat 1070) mixed with 4 parts water Engine Oil can be applied if core washing detergent is not available 3. Install the annulus fillers in their initial positions NOTE:

A maximum of 5 replacement annulus fillers can be installed without a change to the positions of the full set. If new annulus fillers are installed, the moment weight of each replacement must be no more than +10/-10 grams of the removed filler. 4. Make sure the lugs of the annulus filler are fully engaged in the lugs of the LP compressor disc. 5. Make sure the annulus fillers are aligned at the forward end and that the rear is located below the rear air seal NOTE:

The information on the annulus filler including serial number, part number and weight, is found on the underside at the rear.


NOTE:

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REMOVAL/INSTALLATION OF THE FAN BLADE (72−31−41)

Installation Procedure:

WARNING:

YOU MUST USE APPLICABLE GLOVES WHEN YOU HOLD THE FAN BLADES. THE LEADING EDGES OF THE BLADES CAN CAUSE INJURY.

WARNING:

YOU MUST MAKE SURE YOU CAN HOLD THE WEIGHT OF THE COMPONENT BEFORE YOU REMOVE /INSTALL IT. IT IS HEAVY AND CAN CAUSE INJURY TO PERSONS AND DAMAGE EQUIPMENT.

CAUTION:

YOU MUST MAKE SURE THE BLADES DO NOT TOUCH ADJACENT BLADES AS DAMAGE CAN BE CAUSED IF THE BLADES TOUCH.

NOTE:

The LP Compressor Blade weighs 15.2 Kg (33.5 lb)


Removal Procedure 1. Turn the LP rotor so that the blade to be removed is at Bottom Dead Centre (BDC) and install Immobilizer HU44079 to prevent movement of the out of balance fan assembly. 2. Using extracter HU29255 & adapter HU37594 remove the chocking pad and slider 3. Lift the blade to disengage the shear key then carefully pull the blade forward to remove it. 4. Record the radial moment weight of the blade.

BEFORE INSTALLING THE BLADE ALL UNWANTED MATERIAL MUST BE ROVED FROM THE BLADE DOVETAIL AND THE GROOVE IN THE DISC. THE DRY FILM LUBRICANT SHOULD BE INSPECTED AND REPAIRED AS NECESSARY . IF YOU DO NOT DO THIS N1 VIBRATION CAN OCCUR. 1. If a different blade is being fitted then the Moment Weight Difference (MWD) must be calculated. 2. If the MWD is between +80 and −80 oz.in then the installation can proceed. 3. If the MWD is more than +80 and −80 oz.in, then the procedure should be followed to remove the blade opposite to the initial blade removed. 4. Install the blade into the slot until the shear key engages. 5. Put the slider assembly into the opening of the disc groove above the blade, push it rearward then fully install using a nylon faced mallet. On completion a vibration survey & fan trim balance is required, unless: A. You have replaced no more than 3 blades and the MWD is between +8 and −8 oz.in of the blade it replaces. B. You have replaced no more than 5 annulus fillers and the weight difference is between +10 and −10 grams of the annulus filler it replaces. CAUTION:

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FAN TRIM BALANCE Reason for the Job: Some repair work, including fan blade replacement, can affect the balance of the Low Pressure (LP) Compressor. The balance of the fan can also change with time as the engine wears. A fan that is not balanced causes engine vibration. Trim Balance Methods There are two methods of fan trim balance in the AMM: S .The One−Shot Trim Balance S Trial Weight Trim Balance. The one−shot method uses the data recorded by the Engine Monitoring Unit (EMU) during flight or ground runs and gives the necessary information in order for the trim balance weights to be installed in the correct positions to reduce the level of vibration of the fan assembly. Flight data should be used where possible, particularly if the fan vibration has been changing with time. Ground data is normally used if components on the fan have been changed or repaired since the last flight. The trial weight method is used if the one−shot method is not giving good results and fan vibration remains high. Occasionally some engines exhibit different vibration characteristics to the majority of engines and generic coefficients cannot be used.

NOTE:

There are 60 positions where trim balance bolts can be installed. The hole positions are numbered counter−clockwise, when you look at the engine from the front. Their numbers start from the asterisk that identifies hole position No.1.

Description: There is only one part number for trim balance weights. When required, the standard bolt is removed and replaced by a trim balance weight. The trim balance weights can be identified by the part number on the bolt head, when installed in the rear spinner. Removal of a standard bolt and installation of a trim balance bolt increases the mass of the assembly by 13.32 g (0.470 oz.) Mass of standard bolt = 12.36 g (0.436 oz.) Mass of trim balance bolt = 25.68 g (0.906 oz.)


NOTE:

Fan Trim Balance Weights The fan trim balance weights are installed on the rear spinner outer circumference near the rear edge. The bolt holes contain either standard bolts or trim balance bolts. All trim balance bolts are the same weight and have the same part number (the part numbers of the bolts are vibro−engraved on the bolt head).

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BALANCE-WEIGHT ASSEMBLY

STANDARD BOLT ASSEMBLY

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BORESCOPE ACCESS PORTS Description To inspect the gas path of the engine there are many borescope access ports provided as follows: S IP Compressor - 4 ports S HP Compressor - 4 ports S Combustion Chamber - 6 ports S HP turbine - 2 ports S IP turbine - 2 ports S LP turbine - 5 ports On the turbine section some ports are used to inspect HP/IP or IP/LP stage 1. There are a total of 21 borescope access ports, all of which are located on the right side of the engine except for the combustion chamber ports which are located radially around the combustion case.


NOTE:

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IP COMPRESSOR BORESCOPE ACCESS Borescope Plug Removal: The procedure that follows is the same for the blanking plugs at positions. IP3S, IP5S, IP7S . IP3S, IP5S, IP7S: Remove the two retaining bolts and using impact extractor HU29255 and adapter HU51166, remove the blanking plug. IP1S Remove the two retaining bolts and remove the blanking plug. Borescope Plug Installation: On completion of the inspection carry out the following actions: Clean the mating faces of the blanking plugs and the IP compressor case (AMM Task 70−20−01−100−802) Use a brush to apply a thin layer of Omat 4−62 anti−seize compound to the location surface of the plug end and the mating faces of the blanking plug and IP compressor case. Put the blanking plug into position in the IP compressor case and install the bolts. YOU MUST NOT USE THE BOLTS TO PULL THE BLANKING PLUGS INTO POSITION. IF YOU DO, YOU CAN CAUSE DAMAGE TO THE PLUG AND ENGINE. Torque the bolts to the figure given in the AMM


CAUTION:

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HP COMPRESSOR BORESCOPE ACCESS Borescope Plug Removal The procedure that follows is the same for the blanking plugs at positions − HP inlet, HP1S, HP2S. A different extractor adapter is used for HP5S blanking plug. HP inlet, HP1S, HP2S: Remove the two retaining bolts and using impact extractor HU29255 and adapter HU51166, remove the blanking plug. HP5S Remove the two retaining bolts and using impact extractor HU29255 and adapter HU28499, remove the blanking plug. Borescope Plug Installation: Clean the mating faces of the blanking plugs and the HP compressor case (AMM Task 70−20−01−100−802) Use a brush to apply a thin layer of Omat 4−62 anti−seize compund to the location surface of the plug end and the mating faces of the blanking plug and HP compressor case. Put the blanking plug into position in the HP compressor case and install the bolts. NOTE:

On the HP5S blanking plug, install a new face seal on the plug before installation.

YOU MUST NOT USE THE BOLTS TO PULL THE BLANKING PLUGS INTO POSITION. IF YOU DO, YOU CAN CAUSE DAMAGE TO THE PLUG AND ENGINE. Torque the bolts to the figure given in the AMM


CAUTION:

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COMBUSTION CHAMBER BORESCOPE ACCESS Borescope Plug Removal / Installation S Remove the HP compressor exit (T30) thermocouples (AMM Task 77−33−12−000−801) S Remove the bolts and using Impact Extractor HU29255 and Adapter HU28499 remove the combustion borescope blanking plugs. S Remove and discard the face seals from the blanking plugs. S Carry out inspection. S Install the HP compressor exit (T30) thermocouples (AMM Task 77−33−12−400−801). S Clean the mating faces of the blanking plugs and the combustion outer case (AMM Task 70−20−01−100−802). S Install new face seals to the blanking plugs. S Apply with a brush a thin layer of anti−seize compound (Omat 4−62) to the mating faces of the blanking plugs and the combustion outer case. S Fit the blanking plugs into position. CAUTION: YOU MUST NOT USE THE BOLTS TO PULL THE BORESCOPE BLANKING PLUGS INTO POSITION. IF YOU DO NOT OBEY THIS INSTRUCTION, DAMAGE TO THE PLUG AND/OR ENGINE CAN OCCUR S Torque the bolts to figure given in AMM


CAUTION:

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HP TURBINE BORESCOPE ACCESS HP NGV Borescope Plug Removal/Installation S Remove the bolts, the blanking plate and the borescope access blanking plug. S Make sure the face seal has been removed with the cover − Remove and discard the face seal S Clean the mating faces of the blanking plug and the combustion outer case (Task 70−20−01−100−802) S Apply with a small bristle brush a thin layer of anti−seize compound (OMat 4−62) to the plug thread S Install the HP NGV borescope blanking plug in the combustion outer case Align the plug end into its location by moving the central rod at the hexagonal end of the plug S Torque the HP NGV blanking plug to figure given in the AMM S Clean the mating faces of the blanking cover and the combustion outer case & install a new face seal on the cover S Apply with a small bristle brush a thin layer of anti−seize compound (OMat 4−62) to the mating faces of the cover and the combustion outer case S Put the cover into position on the combustion outer case and install the bolts S Torque the bolts to the figure given in the AMM

IP Turbine Borescope Plug Removal/Installation S Remove the IP Turbine borescope blanking plug S Clean the mating faces of the blanking plug and the IP turbine case (Ref Task 70−20−01−100−802) S Apply with a small bristle brush a thin layer of anti−seize compound (OMat 4−62) to the threads and the mating faces of the blanking plug S Put the IP turbine blanking plug in the IP turbine case. S Torque the HP NGV blanking plug to figure given in the AMM LP Turbine Borescope Plug Removal/Installation S Remove the LP borescope blanking plugs. S Clean the end and mating faces of the LP blanking plug and the IP/LP turbine cases (AMM Task 70−20−01−100−802). S Apply with a small bristled brush a thin layer of anti−sieze compound to the location surface of the plug end and the mating faces of the blanking plug and IP turbine case. S Fit the LP blanking plugs in the turbine case. S Torque the borescope plug to figure given in the AMM



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Figure 42 FRA US/T

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TURNING THE LOW PRESSURE (L.P.) SYSTEM (AMM 72−00−00−860−801) WARNING:

YOU MUST BE CAREFUL WHEN YOU DO WORK ON THE ENGINE PARTS AFTER THE ENGINE IS SHUT DOWN. THE ENGINE PARTS CAN STAY HOT FOR ALMOST 1 HOUR.

WARNING:

YOU MUST NOT TOUCH HOT PARTS WITHOUT APPLICABLE GLOVES. HOT PARTS CAN CAUSE INJURY. IF YOU GET AN INJURY PUT IT INTO COLD WATER FOR 10 MINUTES AND GET MEDICAL AID.

WARNING:

MAKE SURE THE APPLICABLE COVERS ARE INSTALLED TO THE REAR OF THE ENGINE. THE MOVEMENT OF AIR THROUGH THE ENGINE CAN CAUSE THE L.P. COMPRESSOR TO TURN VERY QUICKLY AND CAUSE INJURY.

WARNING:

YOU MUST USE APPLICABLE GLOVES ON YOUR HANDS WHEN YOU HOLD THE LP COMPRESSOR BLADES. THE LEADING EDGES OF THE BLADES CAN CAUSE AN INJURY

SAFETY PRECAUTION Make sure engine has been shutdown for at least 5 minutes.


Turn the LP System You must go into the air intake cowl to turn the L.P. system which can be turned by hand. Procedure: S Position a suitable access platform in a safe position and install the Exhaust Nozzle and Thrust Reverser Covers S Position a suitable access platform in a safe position at the Engine Air Intake Cowl. And install the inlet protective rug into position in the air intake cowl. Make sure red warning flag of the mat can be seen externally of the intake cowl. S Enter the intake cowl. and turn the L.P. compressor with your hand. When task is complete ensure all equipment tools and fixtures are removed.

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Figure 43 FRA US/T

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TURNING THE INTERMEDIATE PRESSURE (IP) SYSTEM (AMM 72−00−00−860−802) ATTENTION: Warnings and Cautions Observe all Warnings and Cautions in the AMM. Turn the IP System The variable inlet guide vanes are normally at the fully open position when the engine is shut down. If they are not fully open then the following procedure should be used to so that the IP system turning tool can be installed. If you do the procedure on an inboard engine, do the deactivation of the thrust reverser. YOU MUST MAKE THE THRUST REVERSER UNSERVICEABLE (INSTALL AND SAFETY THE INHIBITION DEVICE) BEFORE YOU DO WORK ON OR AROUND THE THRUST REVERSER. IF YOU DO NOT INSTALL AND SAFETY THE INHIBITION DEVICE YOU CAN CAUSE ACCIDENTAL OPERATION AND/OR DAMAGE TO EQUIPMENT.


WARNING:

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Procedure S Position a suitable access platform in a safe position and install the Exhaust Nozzle and Thrust Reverser Covers S Position a suitable access platform in a safe position at the Engine Air Intake Cowl. And install the inlet protective rug into position in the air intake cowl. Make sure red warning flag of the mat can be seen externally of the intake cowl. S Drain the variable Stator Vane Actuator (VSVA) fuel tubes at the interface with the winged bib into a clean container. S Remove the applicable gas generator fairings to get access to one of the VSVA‘s. S Install the VSV tool HU43122 onto the crankshaft and turn in an anti−clockwise direction to the fully open position S Note: some more fuel may come out of the fuel tubes when the VSVA‘s are moved. S Note: The VSVA‘s and mechanism will go back to the closed position during the next engine start or wet motor. S Remove the VSV tool HU43122 from the crankshaft S Install the gas generator fairings removed for access S Install the fuel tubes to the winged bib and torque the end fittings to 6.1m.daN (44.98 lbf.ft) (Task 70−51−00−910−801) S Access to the IP rotor is from the engine intake reaching through the LP Compressor (fan) blades. S Install the immobiliser (TBD) to prevent movement of the LP Compressor Rotor. S Carefully put the turning tool (HU43985) through the LP compressor blades, inlet guide vanes and variable inlet guide vanes to turn the IPC stage 1 rotor blades. S Push the turning tool against the leading edges of the 1st stage IP compressor blades to turn the IP system as required Do the fuel & oil leak check on the fuel tubes.

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Figure 44 FRA US/T

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TURNING THE HP SHAFT The HP rotor provides the drive to the external gearbox and this is utilised for turning the rotor. Procedure Observe all Warnings and Cautions S Remove the bolts and washers and the blanking plate from the front of the gearbox S Remove and discard the seal ring S Carefully install the turning tool (HU43923) into the gearbox and attach with the slave bolts S Use an applicable wrench to turn the turning tool. This will turn the HP system through the external gearbox YOU MUST NOT EXCEED THE TURNING TORQUE VALUE GIVEN IN THE AMM. IF YOU DO NOT OBEY THIS INSTRUCTION DAMAGE TO THE ENGINE AND/OR TOOL CAN OCCUR. S On completion of the turning operation, carefully remove the turning tool. S Install a new seal ring on the blanking plate. S Put the blanking plate into position on the gearbox and install the bolts and washers Torque the bolts to the figure given in the AMM.

CAUTION:

After performing the handcrank procedure, it is an idle leak test to perform.


NOTE:

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Figure 45 FRA US/T

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RADIAL DRIVE SHAFT REMOVAL/INSTALLATION (AMM 72−61−43) Removal Procedure The procedure is contained in the AMM but the main points are as follows: S Observe all the relevant safety precautions S On the OMT, get access to the Power Distribution Control management pages and Open, safety/lock and tag the relevant circuit breakers. S Open the fan and fan exhaust cowls S Remove the lower splitter fairing S Remove the bolts & segments and disengage the lower shroud from the input drive bevel housing S Remove the bolts and washers and disconnect the upper shroud from the intermediate gearbox housing S Disconnect the driveshaft from the driven bevel gearshaft: − Remove the driveshaft attachment bolts and nuts − Turn the coupling half a spline on the gearshaft − Move the coupling and drive shaft adapter up the gearshaft S Carefully remove the driveshaft, shrouds, adapter and coupling from the engine Keep the driveshaft, adapter and coupling together as a set they are a balanced set and identified by the same S/No. S Inspect the weir seals on the driveshaft and coupling and repair as necessary


NOTE:

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Installation Procedure The installation procedure is the reverse of the removal procedure but the main points are as follows: S Lubricate the splines of the driveshaft, adapter and coupling with clean oil (OMat−1011) S Loosely assemble the driveshaft, upper & lower shrouds and install new seal rings on the upper & lower shrouds S Install the coupling & adapter on the driven bevel gearshaft & move to the highest point S Keeping the shrouds retracted, move the top of the driveshaft up until it is around the driven bevel gearshaft, then align the bottom of the driveshaft with the driving bevel gearshaft and lower into position S Turn the drive shaft and align the mark on the rim with the mark on the adapter, move the adapter down to engage the splines S Connect the driveshaft to the driven bevel gearshaft − Move the driveshaft and adapter up until just below the lowest groove on the gearshaft − Move the coupling down and align it‘s inner splines with the lowest groove on the gearshaft. Turn the coupling to align it‘s mark with the driveshaft & adapter & install the bolts & nuts - torque the nuts S Connect the upper and lower shrouds. NOTE:

30 |72 |L3

After installing the upper shroud it is necessary to use pressing tool HU43381 to push the lower shroud into the input drive bevel housing.

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ATA 73 ENGINE FUEL & CONTROL FADEC SYSTEM

FADEC FUNCTIONS:

Introduction A Full Authority Digital Engine Control system (FADEC), together with the aircraft systems, provides control for engine starting, shut down, power management and engine instrumentation The FADEC system is made of sub−systems working together to form a closed loop control system, maintaining efficient engine operation. The two channel Engine Electronic Controller (EEC) uses embedded software to control functions. It also has segregated and duplicated electrical circuits for engine sensors, actuators and digital data busses to aircraft systems. FADEC is used for engine control of the following: S Fuel Metering Valve S Minimum pressure and shut−off valve S VSV actuators S Handling bleed valves S Ignition S Starting: starter control valve and pneumatic starter S Turbine Case Cooling S Hydraulic pump off−load solenoid (request to A/C system) S Thrust Reverser (request to A/C system),

S Control engine start - pneumatic starter sequence, ignition, fuel & hydraulic pump off−load (as necessary). S Control fuel and airflow to provide steady state and transient response for all environmental conditions. S Schedule engine power levels as necessary for aircraft operation. S Schedule thrust reverser deploy and stow control S Provide limit protection for N1, N2, N3, & P30 (plus EGT during ground automatic start) S Provide HP, IP & LP turbine tip clearance control S Shut−off fuel in the event of an N1 or N2 overspeed or LP shaft breakage S Shut−off or limit fuel flow (as permitted by the aircraft) in the event of thrust control malfunction S Provide auto−relight (ignition) if a flame−out occurs S Provide recovery if an engine surge occurs S Provide instrumentation, engine and control data to the aircraft for control computers, cockpit displays, maintenance and data recorders.



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FADEC POWER SUPPLY ON GROUND General Architecture The FADEC (Full Authority Digital Engine Control) system accepts signals from the various aircraft sub−systems and the engine sensors. These signals allow the FADEC to provide all the necessary features to control the engine, command stow and deploy of the thrust reverser and to provide engine data to the aircraft. The system is composed of: S the EEC (Engine Electronic Controller), S the EMU (Engine Monitoring Unit). The EEC is the FADEC central unit, which is a full authority, dual channel, digital electronic control unit, interfacing with the aircraft and engine control system components. The EMU monitors engine vibration and engines condition. The inputs received from the EEC and various engine and environmental sensors are analyzed by the EMU, which generates a report on the engines condition and identifies irregular engine data. For maintenance purposes, the FADEC system can be energized from the ENGine FADEC GrouND PoWeR P/BSW located on the overhead maintenance panel. The EIPM (Engine Interface Power Management) computer achieves the power supply command. Supply on Ground The power supply of the FADEC systems is controlled by the EIPM computer, which supplies the electrical power from the aircraft to the FADEC systems. When the engine is not running, the EEC gets its 115 VAC power supply from the AC BUS 2 and the AC EMER BUS. The EMU is supplied in 115 VAC from the AC BUS 2. The EIPM computer 1(2) itself is supplied in 28 VDC from the DC BUS 1(2). During on−ground maintenance operations, setting the FADEC GND PWR P/BSW to ON allows the EEC to be energized for 10 minutes. The EEC will stay permanently energized if the EEC INTERACTIVE mode is set through the CMS (Central Maintenance System) during the 10 minutes. Releasing out the FADEC GND PWR P/BSW cuts the EEC power supply.

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Figure 48 FRA US/T

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ELECTRONIC ENGINE CONTROLLER (EEC) Location The Electronic Engine Controller (EEC) is located on the upper left side of the fan case at approximately the 10 o‘clock position. Function The main function of the EEC is to control the engine through all ground & flight modes and environmental conditions.


Physical Description The EEC is bolted through 4 anti−vibration mounts at each corner of the EEC housing, to the mount brackets on the fan case. The EEC is grounded and protected against Electro Magnetic Interference (EMI). The unit has two almost identical housings, which contain the EEC channels, A and B. Each control housing contains the power supply/input circuits, pressure sensors and EEC channel circuits. The two EEC channels are isolated from each other. The power supply/input circuits regulate power for each channel of the EEC from the aircraft and dedicated generator inputs. Each channel is provided with a stable DC input. There are 17 electrical receptacles on the EEC housing, 9 on the channel A housing and 8 on the channel B housing. They connect to the mating connectors from the aircraft and engine systems. They are keyed to prevent incorrect fitment. The Data Entry Plug (DEP) receptacle is located on the Channel B housing at the top of the EEC. The EEC harnesses are colour coded, yellow stripes -ChA, green stripes - ChB.

Functional Description The EEC is a microprocessor controlled digital unit, which has two channels of operation, identified as Channel A and Channel B. Each channel is supplied with inputs from the aircraft, FADEC system and cockpit sources. Each channel can monitor and control the operation of the engine using torque motors, solenoids and relays and transmit engine data to the aircraft. The EEC also maintains and supplies data for fault analysis and output to other systems on the aircraft. One channel is the control computer (channel in control) while the other channel is the stand−by computer. The control computer can access the input interfaces of the stand−by computer and would stay in control if a related input becomes defective. If there is a failure of the control computer circuits or power supply, then control would be given to the stand−by computer, which then becomes the control computer. The channel in control is normally alternated on each engine run to make sure the circuits are used and to minimise the risk of dormant faults. During start, between starter cut−out and idle, the EEC will select a channel change using the following selection procedure (in priority): S If one channel has defects then the channel with no defects will get control. S If both channels have defects, the channel in control when the defects are found, will stay in control.

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Figure 49 FRA US/T

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DATA ENTRY PLUG (DEP) Location The Data Entry Plug (DEP) is located on the channel B housing of the EEC at the top and fastened to the engine fancase by a lanyard. Function The EEC has been designed to control all possible configurations of the engine, regardless of individual characteristics. The function of the DEP is to supply the specific engine related data for EEC operation. Description The DEP is a dual channel memory device providing storage for Engine specific performance and configuration information. The DEP consists of a plug and housing, which contains two EEPROM (Electrically Erasable Programmable Read Only Memory) devices located inside the plug, one for each channel of the EEC. Data Stored in the DEP Both DEP EEPROMS are programmed with identical data: S Engine Serial Number S Engine Ratings Selection S TPR/Thrust Trim Relationship S EGT Trim S Idle Trim


NOTE:

TPR Trim The necessary TPR trim is calculated during the engine test to make the TPR indications (at the cockpit) the same for all engines of the same build standard. And changes the calibration of the engine thrust to TPR relation. This relation can be different for each engine because of the manufacturing tolerances. The data stored in the DEP gives the EEC the level of trim that is necessary for the engine. EGT Trim The EGT trim factors the actual engine EGT to a lower value for display in the cockpit. The EGT trim is calculated from data obtained during the engine manufacturers type test to align approved EGT levels with the cockpit indications. Idle Trim The EEC can trim the idle speeds for minimum and approach idle, as necessary, for the aircraft operation. The data stored in the DEP gives the EEC the trim levels that are necessary for this function.

The data in the EEPROM can be changed as required by the use of a test set.

Engine Serial Number The engine serial number is stored in the DEP so that the aircraft can identify engine health data transmitted from the EEC. Engine Rating selection The EEC is programmed with all possible engine ratings. The data stored in the DEP lets the EEC make the selection from memory of the applicable ratings for the aircraft operation.

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Figure 50 FRA US/T

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DEDICATED ALTERNATOR Location The unit is installed on the external gearbox front face and driven by direct drive from the HP shaft (N3).

NOTE:

The primary source of N3 speed for vibration monitoring is transmitted from the EEC Channel A to the Engine Monitoring Unit (EMU).


Purpose The purpose of the Dedicated Alternator is to provide the main source of power to the EEC and provide a speed reference signal of the HP shaft speed (N3). Description The EEC dedicated alternator supplies three−phase power for each EEC Channel during engine operation. The alternator has four independent windings, two isolated three−phase outputs to operate the control electronics and two single−phase outputs to supply the N3 speed for monitoring, control and overspeed sensing. A satisfactory power output is available to the EEC from the alternator at N3 speeds higher than approximately 8 percent. At N3 speeds between 5 and 8 percent the power supply to the EEC is from the alternator and the 115V AC aircraft stand−by power. The alternator is the assembly of a rotor and a stator. The rotor is a cylinder, which contains a set of permanent magnets (below the surface). It is assembled to the related output shaft on the gearbox module. The stator is an outer cover, which contains two electrical windings in an aluminium stator housing. The rotor is aligned with the windings in the stator housing when the two parts are assembled to the gearbox module. An electrical current is magnetically induced in these windings when the rotor is turned. Two electrical connectors (Ch A & Ch B) are attached to the bottom side of the stator. The harness routing is to the EEC where they connect to their related EEC Channels. When the engine HP shaft turns it causes the gears in the external gearbox module to turn. This causes the alternator rotor to turn. An electrical alternating current then flows through the stator windings and alternator output harnesses. The frequency of these voltages is in proportion to the N3 shaft speed. At engine speeds higher than 8 percent N3, the output from the alternator only is sufficient for the EEC to use (as regulated by the EEC power supply circuits). FRA US/T

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Figure 51 FRA US/T

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ATA 73 ENGINE FUEL & CONTROL; ATA 77 ENGINE INDICATING SHAFT SPEED MEASUREMENT Introduction There are three primary rotors in the engine known as the LP (Low Pressure), IP (Intermediate Pressure) and HP (High Pressure) rotors. These rotate independently of each other and consequently are measured independently and shown as a percentage equivalent (N1, N2 and N3 rotor speeds) on the ECAM displays.


Component Location The following components are fitted in the system: S The LP shaft phonic wheel (60 teeth) is installed to the rear of the roller bearing inner race. S The IP shaft phonic wheel (60 teeth) is installed on the IP compressor front stubshaft. S Four LP speed probes installed in the front bearing housing S Four IP speed probes installed in the front bearing housing. S Dedicated alternator installed on the front face of the external gearbox. Description N1 & N2 shaft speeds are measured using probes that interact with phonic wheels. The output from the speed probes is sent to Channel A and Channel B in the EEC. Two speed probes on each shaft output to Channel A and the other two speed probes on each shaft output to Channel B. N3 speed is supplied by the dedicated alternator, which is turned by the gearbox and HP rotor. There are two separate single phase N3 speed windings in the dedicated alternator which provide the N3 speed to both channel A and channel B of the EEC. The EEC uses these speed inputs to facilitate speed monitoring, engine control and overspeed sensing. The EEC sends digital N1, N2, and N3 signals to the Aircraft for indication. In the unlikely event of total loss of speed signals, the EEC generates a synthesised N1 and N2 to support cockpit indication and N3 to maintain transient control.

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ENGINE PROTECTION SYSTEMS Introduction The Protection System is incorporated into the EEC and provides hardware to perform the following functions: S LP & IP Rotor Overspeed protection S LP Turbine Overspeed protection S TCM (Thrust Control Malfunction) protection LP & IP Rotor Overspeed System The EEC monitors the LP & IP compressor shaft speeds (N1 & N2). If the measured values are above the defined limits, overspeed of the engine is detected and the engine is automatically shut down. Turbine Overspeed System (TOS) The EEC compares the LP compressor speed with the LP Turbine speed. If the speed difference is more than the limit, it is an indication of a shaft breakage and the engine is automatically shutdown.

System Operation Each channel of the EEC has a hardware protection system incorporated in it, which is separate from the other EEC control functions. Comparators are used to determine if an overspeed or TCM have exceeded the threshold values. If set, a fuel shutoff or reduction command is sent to both protection PAL‘s (Programmable Array Logic) from the channel that has detected the condition. The first set of this channels torque rail enable switches are closed. The protection PAL determines if there is a request from the opposite channel for fuel shutdown or reduction and if so, closes the second set of enable switches. Combinational logic is then used to set the current command required for fuel reduction or shutdown to the protection motor in the Hydromechanical Unit (HMU). In normal operation both sets of enable switches need to be closed before the Protection system outputs to the protection motor in the HMU. In degraded operation (power supply or processor failure), shutdown or fuel reduction can be activated by one set of enable switches.


Thrust Control Malfunction (TCM) The EEC monitors the engine thrust, both TPR (Turbofan Power Ratio) and N1 speed. If the actual thrust or N1 speed exceeds the commanded values in excess of the limits, the TCM protection system will operate and will either reduce engine power or shut the engine down automatically, depending upon the aircraft speed and altitude. The Flight Controls Primary Computer (PRIM) provides a discrete signal hardwired to the EEC which permits engine shut down.

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P20/T20 PROBE Description The P20/T20 probe is installed inside the air intake cowl at 150 to right of top dead centre when viewed from rear. The probe measures both engine intake pressure and temperature. T20 Temperature T20 is measured by two independent platinum resistance elements. A small amount of air passes over the elements, whilst the rest of the air passes straight through the probe. The two elements are wired one to each channel of the EEC. The system performs compensation for probe self heating effects and the change to measured temperature caused by the probe heater. The EEC also carries out fault detection on the compensated values.


P20 The pressure signal offtake is just above where the main airstream flows through the probe. A pipe passes through the body to the pressure connector on the base plate and a single pipe connects the probe to the transducer in the EEC. P20 is measured by a single transducer, situated in channel A of the EEC. Its output is available to channel B via cross channel communication. The P20 input is filtered to prevent noise degradation of the EEC performance and also subjected to range checks. Probe Heater An electrical de−icing heater element is configured around the probe powered by Aircraft 115V supply. The EEC selects probe heat on and off dependent upon the following: Probe Heat selected ON if: Aircraft is in flight and N1>10% Or Aircraft is on ground and engine is producing thrust with N3>45% and N1>10% Probe Heat is selected OFF if: Aircraft is in flight and engine is not producing thrust, with N1 < 10% Aircraft is on ground and engine is not producing thrust, with N3<45% or N1<10%. FRA US/T

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Figure 54 FRA US/T

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P20/T20 Probe

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EXHAUST GAS TEMPERATURE (EGT) Introduction The EGT is the temperature of the gas at the inlet to the LP turbine. The thermocouples generate an electrical voltage proportional to the temperature at the thermocouple. The thermocouples are connected in parallel in two groups of 7 and an average value of the 7 thermocouples is sent to each channel of the EEC. The left side thermocouples are connected to EEC Ch A and the right side thermocouples are connected to EEC Ch B. The signal is transmitted to the cockpit to be displayed on the upper ECAM display unit.




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EXHAUST GAS TEMPERATURE (EGT) THERMOCOUPLE Description The EGT indicating system uses 14 thermocouple assemblies to measure EGT (T44). The thermocouples are installed in the stage 1 LP turbine nozzle guide vanes (LP1 NGV), through a transfer tube between the turbine case and the LPT1 NGV. The transfer tube isolates the turbine case from the hot exhaust gas in the LPT1 NGV. Each thermocouple assembly consists of two sheathed elements positioned at two different immersion depths within the LP1 NGVs. The outputs from the two elements are balanced, paralleled and brought out to a common pair of terminals to form a single thermocouple unit. The thermocouple assemblies are connected using wires, one made from Nickel Chromium (Chromel) and the other made from Nickel Aluminium (Alumel). Each of these harnesses is connected to a terminal block. Two electrical harnesses, one for Channel A and the other for Channel B connects the terminal blocks to the EEC. The received signal is trimmed by the EEC from data in the Data Entry Plug (DEP), changed to a digital form and transmitted to the aircraft for display by the ECAM system and the Engine Monitoring Unit (EMU). EGT is used for the following: S As a parameter for TPR S Engine condition monitoring S Engine starting S Cockpit indication




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ENGINE MONITORING UNIT Purpose The purpose of the Engine Monitoring Unit (EMU) is to carry out vibration and condition monitoring for the engine. The EMU receives inputs from the EEC and various engine and environmental sensors. It analyses the data from these inputs and provides reports on both normal and abnormal engine condition. Location The EMU is located on the upper left side of the fan case forward of the EEC.


Introduction The EMU contains two processing units, both contained in a fire resistant box: The Signal Processing Module (SPM) & the Main Processing Module (MPM) The EMU has 4 operating modes as follows: S Initialisation Mode − when power is supplied, various tests and operations are carried out to check EMU health. S Normal Mode − when EMU is fully operational S Extreme Mode − when EMU operates outside a specified temperature range. Some SPM & MPM functions are changed. S Maintenance Mode − when engine is not running, allows maintenance staff to download data and carry out software reprogramming.

Main Processing Module (MPM) The MPM receives digital signals from the SPM and EEC. The MPM carries out the following functions: S Software reprogramming − New & updated programs are entered into the MRM through the Up & Down data loading system. S EMU Built−in Test (BIT) function which it collates and sends to the EEC. S Output EMU hardware and software standard. S It performs on−engine analysis of engine performance and health and reports on any irregular engine data.

Signal Processing Module (SPM) The SPM receives inputs for vibration, shaft speeds, engine pressures and oil contamination. The 3 sources of inputs are: S Direct inputs from environmental sensors S Environmental sensor inputs via the EEC S EEC processed data via the CAN data bus NOTE:

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ENGINE MONITORING UNIT (EMU) INTERFACE SPM Inputs The SPM inputs include the following: S Vibration indication from dual transducer S Oil debris monitoring (EMCD) S Engine Pressures: P20, P25, P50, P160 S Shaft Speeds (N1, N2, & N3) S Engine Performance parameters S Engine Control parameters (TPR, fuel flow etc) S EEC Bite data S Aircraft sourced data (altitude, flight phase etc)


SPM Outputs The SPM outputs the following: S Processes and stores Fan Trim Balance data S Carries out broad level built−in test (BIT) of the EMU functions during start. S Sends digital data to the MPM for monitoring and analysis S Sends engine vibration data to the EEC through the CAN bus. The EEC then outputs this data via the AFDX for cockpit display. Sends oil debris alerts to the EEC through the CAN bus. The EEC then sends this data via the AFDX for reporting in the cockpit.

MPM Outputs Input data is analysed by a multi−sensor data fusion system, known as a “Q“ system. Computational methods are then used to identify abnormal engine data and produce related reports as follows: S Engine Novelty Reports showing abnormal engine data. S EEC/Engine Incident Event Summaries and snapshot data S Fan Damage Report S EMU BITE data S The processed data is output via the CAN to the EEC, which the EEC addresses and sends out via the AFDX for: S Status indication on the ECAM S Storage on the aircraft e.g. Engine Novelty Reports S Output through ACARS e.g. Engine Incident Event Summaries & Snapshot data.

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VIBRATION TRANSDUCER

T25 THERMOCOUPLE

The vibration transducer is mounted on the right side of the intermediate case on the end of the No. 2 vane on the upper right side. The vibration transducer is a dual output accelerometer. It contains two peizo−electric crystal stack elements, each with a mechanical load of electrically insulated seismic mass. Each element has a mineral insulated electrical lead, which connects them to an engine harness. The harness connects the transducer to the Engine Monitoring Unit (EMU). The vibration signals to the EMU are used in two ways: S The engine vibration is sent to the EEC, which sends the signal to the ECAM for cockpit display. S The EMU uses the signals to do on−board analysis to give information on engine performance, general health and any irregular engine data.

The input parameter T25 provides a measure of IP compressor outlet temperature. T25 is used solely for engine condition monitoring. One single element thermocouple provides an input to Channel B and the processed temperature is available to Channel A. The T25 thermocouple is installed on the right side of the intermediate case in the No. 2 vane.




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T30 THERMOCOUPLE The input parameter T30 provides a measure of HP compressor outlet temperature. There are two single element thermocouples per engine. The two separate signals are input into each channel of the EEC. Under normal operating conditions the EEC averages the two signals. If one signal is lost, the EEC will use the other signal. T30 is used for: S Engine condition monitoring S Detection of rain/hail ingestion S Engine starting The T30 thermocouples are installed in two of the combustion borescope ports.




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ENGINE MASTER CONTROL OPERATION General The ENGine MASTER lever located on the center pedestal interfaces with the fuel system and the FADEC system. Note that the engine FIRE pushbutton also acts on the LP fuel valve. On the fuel system, the ENGine MASTER lever acts on the LP valve and MPSOV (Minimum Pressure Shut−Off Valve). On the FADEC system, the ENGine MASTER lever is used for the starting mode selection and the EEC (Engine Electronic Controller) memory reset purposes.


Low Pressure fuel valve And Airframe Shut Down Solenoid The MASTER lever is directly hardwired to the airframe shut down solenoid of the HMU. It controls also the Low Pressure fuel valve through the engine master switch relay. Setting the switch from the ’ON’ to the ’OFF’ position directly energizes the airframe shut down solenoid then the MPSOV moves to the close position. After one minute, the power off relay de energized the solenoid in order to avoid heat dissipation into the HMU. This gives the independent authority to close the MPSOV regardless of the EEC command.

ENGine MASTER and EEC reset Moving the MASTER lever from ”ON” to the ”OFF” position, closes both channel reset discrete contacts, resetting both EEC channels; all data stored in the EEC memory will be cleared. ENGine MASTER FAULT light The amber ’’FAULT” light installed located on the ENGine MASTER lever indicates a disagreement between the MPSOV poition and its commanded position. The Master lever FAULT light is managed by the EIPM, based on the digital data received from the related EEC via the IOM.

ENGine MASTER and netwok Interface The MASTER Lever is directly hardwired to each channel of the EEC. Then each channel sends its own discrete signal via the EEC internal data bus to the other channel. This signal is used to keep the MASTER Lever position readable into the EEC in case of AFDX failure. The MASTER lever is also hardwired to the IOM and interfaces with the EEC through the ADCN. The MASTER Lever uses the ADCN signals as source to arbitrate in case of disagreement between network signals or discrete signals into the EEC. The MASTER Lever signal acts on the metering valve servo valve of the HMU, which is the second device to turn on or off the MPSOV.

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ATA 76 ENGINE CONTROLS THROTTLE CONTROL ASSEMBLY COMPONENT DESCRIPTION General The TCA (Throttle Control Assembly) design is based on a modular concept. It is composed of 4 independent assemblies (two inboard assemblies and two outboard assemblies), each one dedicated to one engine. Each assembly is composed of: S A housing, S A throttle lever, S A thrust reverser lever (inboard assemblies, engines 2 and 3), S A/THR instinctive disconnect push button (outboard assemblies, engines 1 and 4), Electrical connectors.

Sensing Devices The primary function of the TCA is to sense the commands and to generate electric signals. This positional information is received by several A/C systems. The throttle control lever sensing devices are composed of 4 independent groups of 2 resolvers, and 4 independent groups of 3 potentiometers. The thrust reverser lever deployed order (inboard levers only) is provided by means of a switch (one per lever). These 2 switches signals are obtained through one track of potentiometers. Throttle Levers and Thrust Reverser levers Detents Points and Stops Thrust reverser lever detent point and stops are located as follows:

Modulation of Engine Thrust Except during A/THR mode, control of the forward thrust of each engine shall be achieved by modulation of the related throttle lever position. The throttle levers can only be moved manually. The throttles move over a sector divided into three areas separated by unique positions. The rating selection is achieved by setting the thrust levers in the pre−determined detent point, which divide the sector. The four throttle levers can be moved independently. Each detent point gives the limit mode for each engine rating. Reverse Mode Control of the reverse thrust of either engine 2 or 3 is achieved by modulation of thrust reverser levers fitted on the throttle lever inboard assemblies. Control of the stow/deploy sequence is achieved when the thrust reverser levers are in reverse area. As in forward mode, the thrust reverser levers can be moved independently. When the throttle levers are not at idle, the thrust reverser levers are mechanically locked in the stowed position.

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Inboard and Outboard Assemblies There are internal mechanical features that are installed into each inboard and outboard assembly, which are: S The Artificial force feel device (friction force), S The soft detent device, related to several thrust settings, S The interlock mechanism (inboard assemblies only). There are internal electrical features that are installed into each inboard and outboard assembly, which are: S The 3 potentiometers and one switch (inboard assemblies) which are installed inside the TTU (Throttle Transducer Unit), S The 2 resolvers, S A/THR push button switch (outboard assemblies), S The electrical connectors. Interlock Mechanism This mechanism is implemented only on inboard assemblies. The purpose of the interlock mechanism is to prevent thrust reverser levers movement from the stowed position if one of the throttle levers is out of the forward idle position. The interlock also has the following functionalities: S Prevent the throttle lever movement forward or backward from idle position if reverse lever is raised. S Automatic recall of the throttle lever to IDLE when thrust reverser lever is moved. S Automatic recall of the thrust reverser lever to neutral when the throttle levers are moved from IDLE and when thrust reverser levers are positioned between 0_ and 10_.

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THROTTLE CONTROL ASSEMBLY INTERFACES Genreal Modulation of engine thrust and selection of the thrust limit mode functions are achieved using throttle position lever sent by resolvers and potentiometers. For each throttle lever: S Each of the two resolvers transmits the angle information of the throttle lever to the A and B channels of the EEC (cross−communicated to the other channel). The EEC supplies 6 VAC power to the resolvers. S Each potentiometer transmits the angle information of the thrust lever to each PRIM. The PRIM (primary system) supplies 10 VDC power to the potentiometers For each thrust reverser lever: The reverse position switch sends a discrete signal via the EIPM (Engine Interface Power Management) to control the ETRAC (Engine Thrust Reverser Actuator Controller) power. In addition of the two resolvers signal, the EEC receives, via the AFDX (Avionics Full Duplex Switched Ethernet) network, three digital throttle angle values coming from the three PRIMs. PRIM potentiometer information is used to consolidate resolver signal selection.




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Throttle Position Selection Logic To measure the Throttle position, the EEC has 5 sources of Throttle angle measurement: 2 Resolvers (one analog signal per channel, cross−communicated to the other channel). 3 Potentiometers signals (sensed by Flight Controls Primary Computers) received from AFDX Network. Based on the 5 sources of throttle position, the EEC does the following logical selection: Resolver and Potentiometers signals are all validated by the EEC (range & consistency checks). The resolvers are selected if they are both validated and agreed by each other (digital information from potentiometers are disregarded). When both resolvers are in disagreement, then the potentiometers are used as a referee to identify which resolver has failed. Then, the EEC selects the valid resolver. If there is a disagreement between a single resolver and the potentiometers, then the potentiometers are selected (via AFDX). Instinctive Disconnect Push Button Interface The disengagement of the A/THR function can be done manually through action on the instinctive disconnect push buttons on the throttle levers. Both Instinctive Disconnects (A/THR disengagement) are directly hardwired to each EEC. The EEC receives also this signal as an AFDX information from the PRIM. The FADEC Autothrust Function is inhibited until the next EEC reset if the Autothrust Instinctive Diconnect signal is asserted continuously for more than 15 seconds. The ESS BUS and DC 2 BUS supply 28 VDC power to the instinctive disconnect pushbutton.

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ATA 77 ENGINE INDCIATING ENGINE POWER PHILOSOPHY General The engine thrust is the result of several cockpit settings. To meter the fuel flow, according to its own laws, the EEC (Engine Electronic Controller) takes into account: S The throttle control levers positions, S The AFS (Auto Flight System) commands, S The KCCU (Keyboard and Cursor Control Unit) take−off data input by the flight crew. The command signals and other relevant input signals are processed within the EEC. Output EEC control signals are transmitted to the engine HMU (Hydro Mechanical Unit) to be converted in fuel flow and through the ACUTE (Airbus Cockpit Universal Thrust Emulator) for the indication of the thrust parameters. The EEC sends to the CDS (Control and Display System) the thrust that must be indicated via the Aircraft AFDX (Avionics Full Duplex Switched Ethernet) network. Manual thrust and Autothrust Two thrust setting philosophies are used in order to obtain the required thrust, manual and automatic modes. In the manual mode, the EEC receives a command signal from the TRA (Throttle Resolver Angle), to set the thrust. Alternatively, when the A/THR (AutoTHRust) is activated, the EEC sets the thrust by taking into account: S The N1 target from the AFS and, S The TRA, for thrust limitation and to set the thrust limit mode. During Take−Off the A/THR function is engaged but not active. Memo Thrust Mode This is a transitive mode of thrust control between the autothrust mode and the manual mode.

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When the autothrust mode is deactivated and the throttle levers are set on the max continuous or max climb detent points, the EEC will enter the memo thrust mode. In this mode the EEC prior to the exiting autothrust mode locks the thrust demand. This is to prevent potential thrust step changes, which may occur when reverting from autothrust to manual mode. Thrust Setting: TPR Mode and N1 Mode There are two EEC internal thrust laws to meter the fuel flow. The ”TPR” (Turbofan Power Ratio) law is the normal operating mode to compute the thrust. The selected parameters for TPR thrust control are: S P20/T20: LP compressor inlet pressure/ temperature. S P30: Combustor inlet pressure. S EGT (Exhaust Gas Temperature): Low pressure turbine inlet temperature. The N1 law is activated as a back−up mode if the TPR mode fails. In manual or in A/THR modes, the EEC dedicated to each engine adapts the metered fuel flow to set the thrust. The EEC prevents the thrust from exceeding the limit related to the throttle lever position in both manual and automatic modes. The EEC controls the engine to an N1 reversionary schedule as a result of cockpit command (ALTerNate mode push button) or loss of TPR parameters. There are two forms of N1 reversionary control: S Rated N1 Reversionary Mode: The TPR command is converted into a N1 command. The EEC calculates N1 command using a simple comparison table ”N1 versus TPR” and the engine is controlled using this N1 command. S Unrated N1 reversionary mode: The EEC sets the forward idle detent position equal to idle N1 and the max take−off detent position equal to ”red line” N1. The EEC then uses a graphical comparison table such that the N1 versus TRA profile is equivalent to the TPR versus TRA profile. The engine is then controlled using this N1 command. Either in TPR or in Rated or Unrated N1 mode, the manual mode or A/THR mode can be achieved. NOTE:

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THRUST CONTROL FUNCTION OPERATION Air Data Selection Logic Engine signals (P0 and P20/T20) are compared by the EEC to the 3 independent ADIRUs (Air Data Inertial Reference Units) signals (PS, PT, TAT) via the Aircraft AFDX network to be used as inputs for the air data selection logic. Engine P0 is measured by a single transducer, which is situated in the EEC. The transducer measures the pressure P0 air pressure from Zone 1 as under−cowl pressure environment. The Engine air data selection logic has the input of each of the three parameters (P0, P20 and T20) of the EEC compared with each of the three ADIRU parameters, which are: S Ps (static pressure) equivalent to engine P0, S Pt (total pressure) equivalent to engine P20, S TAT (Total Air Temperature) equivalent to engine T20 The 3 ADIRUs plus the 4 EECs give a total of 7 available sources that are compared and validated through the AFDX network, to compute the TPR or N1 command. To make sure that the engine thrust symmetry or N1 symmetry and the selection between the TPR and the N1 mode are related to the availability of air data inputs to the EEC. TPR actual Calculation TPR actual is derived from the P20, P30, T20 and EGT parameters. P20/T20 probe is installed in the Engine air intake forward of the fan. The probe is electrically heated to prevent ice formation. P30 (measure of the HP Compressor exit pressure) is used into the EEC to calculate the TPR and to schedule fuel to the burners. 14 EGT thermocouples (low pressure turbine inlet temperature) supply a gas temperature measurement. This temperature measurement is also used to compute theTPR. The value of TPR is calculated using the following relationship: ǸTGT TPR + P30 P20 ǸT20

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Reversionary Thrust Control The Reversionary Thrust Control gives a backup control in the event that the FADEC System can no longer support theTPR control. The reversionary thrust control mode has the following settings: Rated reversionary thrust control, which is selected when there is not enough valid signals to calculate a TPR thrust setting demands. Unrated reversionary thrust control, which is selected when there are not enough valid parameters available to calculate the TPR thrust setting demands. Rated Reversion Thrust Control Rated reversion is used when it is not possible to calculate an engine TPR actual, but theTPR command can still be derived and so rated N1 is derived from the TPR command. The rated reversionary thrust control N1 command is calculated as the product of T20 and TPR command and calculated mach number. The EEC selects rated reversionary thrust control when one or more of the following conditions are true: S EGT measurement has been confirmed as Invalid. S Selection of model P30 has been confirmed as Invalid. S P30 measurement has been confirmed as Invalid. S TPR measurement has been confirmed as Invalid. S TPR control loop upward run−away is detected. S P30 pipe fault detection has been confirmed. S P30 pipe freezing has been detected.

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Unrated Reversion Thrust Control The unrated reversionary thrust control N1 command is selected as the reversionary thrust control N1 command when it is not possible to calculate a TPR demand. The unrated reversionary thrust control N1 command is scheduled as a function of TRA position and altitude. The EEC selects unrated reversionary thrust control when one or more of the following conditions are true: S P0 signal has been confirmed as Invalid. S P20 signal has been confirmed as Invalid. S T20 signal has been confirmed as Invalid.


Autothrust Control The AFS interfaces with the FADEC to give the autothrust function, including the Alpha Floor protection. The autothrust function can be engaged or disengaged according to the logic implemented in the PRIMs computer. When engaged, the function is either active or inactive. Once engaged and active, the EEC uses the airframe N1 target to set the engine power level. In normal mode, even if the A/THR sends an N1 target to the engine, the THR is computed from the TPR. Autothrust is operative in the TPR and ALTerNate (N1) modes. The autothrust function can be engaged if the engines are not in the same mode (TPR or N1). The PRIM accepts the engine in ALTerNate N1 Unrated mode.

Autothrust Function Engagement / Disengagement / Activation The engagement of the autothrust function can be accomplished manually or automatically. The autothrust function can be engaged manually through the A/THR push button of the FCU (Flight Control Unit). The autothrust function is automatically engaged when throttles are set in the take off detent (it is associated to the engagement of the TAKE−OFF / GO AROUND mode of the autopilot) or when the Alpha Floor protection is activated. The disengagement of the autothrust function can be achieved: S Manually via the instinctive disconnect pushbuttons located on the throttle control levers (normal operation), S Manually through the FCU autothrust push−button (if already engaged), S Automatically when all (4) throttle control levers are selected at Idle, S Automatically when all (2) thrust reverser levers are selected to reverse, S Automatically when more than 1 engine is not in A/THR mode, S Automatically in case of more than 1 Engine failure, S Automatically in case of failure seen by the AFS. In case of autothrust disengagement, each Engine is controlled in manual mode, or in memo mode in the case of involuntary disconnection. When autothrust is engaged it can be: S Active: throttle control levers between IDLE and CLB (or MCT (Maximum Continuous Thrust) with one engine failure) and at least one throttle at or below CLB (with no engine failure). Thrust is controlled by the A/THR function. S Inactive: if all throttles are above CLB (or above MCT with one Engine failure). Thrust is controlled by the throttle position. ALPHA FLOOR: Autothrust Activation In case of Alpha FLOOR detection the A/THR mode is automatically activated and commands the TOGA thrust, regardless of the throttle lever position.

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Cockpit Thrust Display (ACUTE) ACUTE (AIRBUS Cockpit Universal Thrust Emulator) is a percentage indication of thrust. The ACUTE function calculates percentage parameters from engine command and thrust feedback parameters, for transmission to the airframe and subsequent cockpit display. The parameters are: S THR Limit, S THR Actual, S THR Command, S THR REF (Throttle), S THR Idle, S THR Max. THR WML: Thrust windmilling is the THR achieved when engine in Wind Milling (0%). THR 100: Thrust 100 is the THR achieved when Throttle at TOGA and Bleed Off (100%). THR IDLE: Low−end of grey sector, corresponds to the THR achieved when the engine is operating at Idle. THR MAX: High−end of grey sector agrees with the THR achieved when throttle at TOGA detent. THR Actual: raw engine thrust corrected by engine drag. The parameters THR100, THR Limit, THR Actual, THR Command, THR Throttle, THR Idle, THR MAX, are sent to the airframe CDS through the AFDX network. When operating in Unrated N1 mode, the EEC THR parameters output sent to the airframe CDS are not computed.

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Thrust Limit Modes and Thrust Rating Limit The thrust limit modes (Max CLB, Derated Climb, Derated Take−Off, FLEXible take off, MTO, MCT, GA, FlexGA), are calculated to show the engine thrust setting mode from which the THR LIMIT (THRust Limit) value is computed. The selected thrust limit mode is shown in the cockpit beside the thrust limit value. THR Limit value in N1 mode is the value of THR Limit calculated as derived for N1 mode.

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FADEC ARCHITECTURE & INTERFACE DESCRIPTION FADEC Overview The FADEC (Full Authority Digital Engine Control) system, together with aircraft systems, gives the control for engine starting, shut down, power management and engine indicating. The FADEC system is controlled and monitored by an EEC (Engine Electronic Controller). The EEC is a dual channel digital unit. The EEC reads inputs from the aircraft and the engine systems and provides engine control and cockpit indications. Output data from the channel A of the EEC is sent to the EIPM (Engine Interface Power Management) computer via an ARINC 429 bus, for Back−up purpose. The EEC sends also a N1 ANALOG speed back−up signal directly wired to the EIPM. The N1 speed value is then forwarded to the IOM through ARINC 429 bus. Each channel of the EEC receives its own TRA (Throttle Resolver Angle) analog signal from the Throttle Control Assembly, independently from the AFDX network through two dedicated resolvers. The A/THR (Autothrust) instinctive disconnect discrete signal is directly hardwired to each channel of the EEC. Both A/THR instinctive disconnect P/Bs are used by the flight crew to disengage the A/THR mode on all engines. The EEC exchanges signals and data with the EMU (Engine Monitoring Unit). The EMU analyses data from engine sensors such as pressure sensors, accelerometers, tachometers and electrical magnetic chip detectors. The EMU gives a report on the engine condition and identifies irregular data. Some processed data are sent from the EMU to the EEC for cockpit display. EEC acquires the following discrete signals from cockpit panels, through 4 IOM’s (Input Output Module) and the ADCN (Avionics Data Communication Network): S Rotary selector CRANK/NORM/IGN START position S ENG MAN START p/b switches status S ENG ALTN MODE p/b switch status (for N1 Back−up mode) S MASTER lever position ON/OFF (one per engine), to initiate the Engine Starting sequence (in Automatic Start) or to turn the fuel on (in Manual Start or Wet crank). A discrete signal is directly hardwired from the MASTER lever to each channel of the EEC for EEC reset, and to keep the MASTER Lever position in case of AFDX failure. FRA US/T

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The MASTER Lever FAULT light is managed by the EIPM, based on digital data received from the related EECs (own and opposite), via the 4 IOMs. EIPM sends a discrete signal to the ENG MASTER lever for fault light

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FADEC Overview

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EEC Aircraft Interfaces EEC has digital interfaces, analog and discrete inputs/outputs. EEC Digital Interfaces The four EECs have digital interfaces with the aircraft systems through the ADCN. The IOMs transmit cockpit commands (Master Lever, Rotary Selector, N1 P/B, Man Start P/B) and the EIPM and the AICU (Anti−Ice Control Unit) data to the EEC. The EIPMs receive, via the IOM, engine status data (speed, starting, shutdown, reverse inhibition, reverse locked...) from the EEC own and opposite data busses. The AICU (Anti−Icing Control Unit) receives, via the IOM, data on engine running, maximum take off/go around, flex take off and derated take off limit mode selected. It sends WAI/NAI (Wing Anti−Ice/Nacelle Anti−Ice) status to the EEC for engine thrust modulation. The EBAS (Engine Bleed Air System), the PADS (Pneumatic Air Distribution System) and the OHDS (Overheat Detection System) are hosted in CPIOM−A. Those systems receive data from the EEC concerning: S Engine status (engine starting and running, reverse operation) and starting information for Pack closure, S Engine pressure & temperatures (P0, P30, T30), S Starter control valve position, S Burst duct detection (to OHDS). The EBAS, the PADS and the OHDS send data to the EEC concerning: S Bleed configuration status, S HP/IP Command, S Bleed manifold Pressure, S Cross−bleed valves position, S APU isolation valve. The AVS (Avionics Ventilation System), the AGS (Air Generation System) and the CPCS (Cabin Pressure Control System) are hosted in CPIOM−B. The AVS receives data from the EEC concerning the Engine Status (engine running/ not running). The CPCS receives data from the EEC concerning the Engine running, the engine take−off power and the N1 speed. The AGS receives

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engine starting information for the closure demand of the pack valves from the EEC. The FWS (Flight Warning System) is hosted in CPIOM−C: It receives engine failures warning annunciation and engine status (speed, starting, shutdown, reverse operation) from the EEC via the ADCN and from EIPM in back up with ARINC 429 bus. The WBBC (Weight and Balance Back−up Computer), hosted in CPIOM−C, receives data on fuel used from the EEC. The ATC (Air traffic Control) is hosted in CPIOM−D1 and receives data from the EEC on engine status (engine running, not running...). The FQMS (Fuel Quantity Management System) is hosted in CPIOM−F: It receives Fuel used data from the EEC and sends fuel temperature data to the EEC. The LGERS (Landing Gear Extension Retraction System) is hosted in CPIOM−G: It sends wing and body landing gears status (for flight/ground status computation) to the EEC. The DSMS (Doors and Slides Management System) receives engine running data from the EEC. The EPGS (Electrical Power Generating System) receives data from the EEC on the engine status (MASTER lever OFF, engine Start/crank sequence active) and on engine speeds N3. The PRIM (Flight controls and Guidance Computer) receives Engine status (speed, starting, shutdown, reverse operation) and autothrust feedbacks (actual thrust, commanded thrust, thrust limits...) from the EEC. It sends to the EEC: S Autothrust command, S Autothrust engaged and active signals, S Alpha floor protection, S Throttle position (for consolidation of EEC signals), S T/O mode selection input (Flex temperature, derated T/O levels, Derated Climb levels), S TCM (Thrust Control Malfunction) permission discrete command, S Wheel speed (provision).

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The AESS (Aircraft Environment Surveillance System) receives Engine running, selected take off power and thrust data from the EEC. The ADIRS (Air data and Inertial reference System) receives engine running data and engine Ps, Pt, TAT (for consolidation of ADIRU internal parameters) data from the EEC. The ADIRS sends air data parameters (Ps, Pt, TAT, M), probe heat status (pitot, L/H static, TAT, AOA), and CAS (calibrated airspeed) to the EEC. The ECB (APU controller) receives start sequence signal for the APU boost from the EEC and sends APU availability signal (for bleed configuration determination) to the EEC. The SFCC (Slat/Flap Control Computer) sends slat/flap configuration (for approach selection) to the EEC. The CDS (Command and Display System) receives from the EEC: S Engine Primary parameters (THR, N1, EGT), S Engine secondary parameters (N2, N3, FF (Fuel Flow), Fuel Used, Oil Quantity, Oil Temperature, Oil Pressure, Vibration levels), S Engine status (speed, starting, shutdown, reverse operation). The ACMS (Aircraft Condition Monitoring System) and the FDIAF (Flight Data and Interface/Acquisition Function) are hosted in the CDAM (Centralized Data Acquisition Module). Those systems receive engine data for performance and trend monitoring, engine manufacturer’s reserved parameters, and EMU advanced Maintenance reports. The FDRS (Flight data Recording System), linked on the CDAM, receives data from the EEC concerning: S PS3, Regulated Pressure, Engine bleed demand/K factors, S N1/TPR limit, N2, N3, S Each thrust reverser position, and throttle / power lever position, S EGT, oil quantity, Engine Vibration, oil temperature, and oil pressure, S HP/LP fuel valve, S Fuel flow, and derated take−off, S Position engine relight indication, S Thrust command, S Engine warning (each engine vibration), S Thrust/Power on each engine.

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The CDAM transmits its data to the OMS (Onboard Maintenance System). The OMS application interfaces with ADCN through the SCI (Secure Communication Interface) data.

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ATA 73 ENGINE FUEL & CONTROL EEC ANALOG AND DISCRETE INPUTS/OUTPUTS The EEC has direct interfaces with aircraft systems and cockpit controls. It receives and sends analog and discrete data. Control of the engines is achieved by modulation of a throttle lever angle. The TCA (Throttle Control Assembly) receives the excitation current for the resolvers from each channel of the EEC. The TRA (Throttle Resolver Angle) of the throttle lever position is transmitted in analog signals to each channel of the EEC. A discrete signal from the MASTER lever is directly hardwired to each channel of the EEC, for EEC reset function. The activation of the A/THR instinctive disconnect P/B is used to disengage the A/THR mode on all engines. An A/THR instinctive disconnect discrete signal is directly hardwired to one EEC channel (internally cross−wired) as well as to the Flight Controls Computers (PRIMs) and to the Flight Warning System. In order for the Engine Control System to protect against TCM (Thrust Control Malfunction), an independant discrete signal from the aircraft is directly hardwired to each EEC. The purpose of this independant input is to authorise the EEC to shut the engine down if it has detected an uncommanded and uncontrollable thrust excursion, which may affect the aircraft controllability. The TCM protection signal is set by the Flight Controls PRIM (PRIMary Computer). Each EEC receives from the Airframe hardwired discretes indicating position on the aircraft. These discretes are directly hardwired from the Pylon jonction box to the EEC. A N1 speed back−up signal will be made available at the aircraft level. The analog signal is wired directly from the N1 sensor on the engine to the EIPM computer. The N1 back−up indication is used to keep as a minimum the N1 display available under the following cases: S AFDX network failure, S Complete loss of EEC, S Complete loss of AFDX busses on the engine. When engine speed is detected to be higher than 50% N3 (for Trent 900 Engine), both EEC channels set the engine running discrete output.

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However, only the output from the channel A is planned to be acquired on the aircraft side by the IOMs 1 and 2, the Emergency Power Center, the PEPDC (Primary Electrical Power Distribution Center) and the HSMU (Hydraulic System Monitoring Unit). For engine in−flight wind milling restart purposes, the EEC has the possibility to depressurize both hydraulic pumps on the engine. To achieve this function, both EEC channels are able to switch one output ground/open discrete signal that commands the depressurization of both engine−driven hydraulic pumps.

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EEC COMMAND AND SENSOR INTERFACES The FADEC has to perform engine control and monitoring. The DEP (Data Entry Plug) is a dual channel serial memory device providing storage for engine specific performance and configuration information. The DEP is a plug and housing, which is fastened to the engine by the use of a lanyard. The data entry plug is only programmed with the applicable data for the engine on which it is installed. It cannot be removed and then installed to a different engine unless it is programmed for that engine. The data entry plug is programmed with the data that follows: S Turbofan power ratio trim, S Engine rating selection, S EGT trim, S Engine serial number, S Idle trim. Note: If the DEP and the engine do not have the same data the engine will not operate normally. The fuel flow XMTR (Transmitter) continuously monitors the fuel flow to the combustion system. The XMTR supplies analog signals to the EEC that are in proportion to the mass fuel flow rate. The EEC uses these signals to calculate the flow rate and the quantity of fuel that has been used. The EEC then transmits this data for display in the cockpit. The Fuel filter differential pressure switch indicates to the EEC if the fuel filter is coming clogged. The oil quantity XMTR is installed through an opening in the center of the top face of the oil tank. The EEC uses this signal for display in the cockpit. The oil pressure XMTR senses the difference between supply and scavenge oil pressures. One XMTR per each channel of the EEC supplies an oil pressure indication. The oil temperature thermocouples are installed at the top of the scavenge oil filter housing. The system uses the thermocouples that are sensitive to temperature changes. An oil temperature signal is sent through the EEC to the aircraft indicating system. The filters differential pressure switches (supply and scavenge) compare the difference between upstream and downstream pressure for their related filters.

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Dual vibration XDCR (Transducer) signal and magnetic chip detector signal are computed by EMU, which monitors engines performance and trend, engines vibration. Both channels of the EEC have a T20 thermocouple analog input. The T25 thermocouple sends, to the EEC, the signal of the TAT (Total Air Temperature) at the IP compressor exit. This signal is used for health monitoring purposes. The signal is input to both channels of the EEC. The T30 signal is obtained by two single element thermocouples mounted at different radial positions around the engine. Each thermocouple sends a signal to the related channel of the EEC. T30 is the TAT at the HP compressor. The EGT (Exhaust Gas Temperature), or TGT (Turbine Gas Temperature) is derived from 14 double element thermocouples mounted in the nozzle guide vanes. The thermocouples are wired in parallel by two leads, one in alumel and one in chromel. Each pair of leads is connected to each channel of the EEC. The TCAF (Turbine Cooling Air Front) probe converts the IP turbine disc−cooling air temperature at the front of the disc into an electrical signal. The TCAR (Turbine Cooling Air Rear) probe converts the IP turbine disc−cooling air temperature at the rear of the disc into an electrical signal. TCAF and TCAR thermocouples are used to provide IP Turbine disk overheat detection. A single thermocouple, mounted on the IP/LP TCC flange, sends to the EEC a temperature signal from Zone 3 nacelle, used for condition monitoring purposes. The N1C and N1T shaft speeds are derived from engine pulse probes. The probes provide a sinusoidal frequency voltage proportional to the LP compressor and LP turbine shaft speed rotation. The N1 Compressor speed signal is sent to each channel of the EEC, as a main parameter for thrust limitation and N1 mode back−up computation. The comparison between N1C and N1T speeds is used to give a LP Turbine overspeed protection. The N2 shaft speed signal is derived from engine pulse probes. The probes supply a sinusoidal fraquency voltage proportional to the IP shaft speed rotation. The N2 speed signals are used for engine control functions and are used by the ROS (Rotor OverSpeed) protection. The N2 speed signal is sent to each channel of the EEC.

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The N3 shaft speed signal, used within the EEC, is derived from the PMA (Permanent Magnetic Alternator). The outputs from the PMA are at a frequency proportional to the N3 shaft speed and send an N3 speed signal to each channel of the EEC. The P0 signal is input to channel B. The P20 probe sends to the EEC, the signal of the TAT (Total Air Temperature) at the engine air intake. The EEC automatically selects the P20 probe heater to prevent ice on the probe air inlets. The P20 signal is input to channel B. The HP compressor pressure signal called P30 is split inside the EEC to give a pressure tapping to a transducer in each channel. The ratio P30/P20 is used for the TPR thrust computation. The IP Compressor pressure signal called P25 is input to the EMU for condition monitoring purposes. The P50 signal is an exhaust gas pressure signal, which is split into the EMU to give a pressure tapping to a transducer in each channel of the EEC. A fan exit signal called P160 is used for condition monitoring purposes and is input to the EMU. The EEC controls the starting system during the engine start sequence, the EEC opens the starter control valve to operate pneumatic starter from either APU air, cross−bleed air or an external air source. The EEC receives feedback from the starter control valve position switch. The EEC supplies the two ignition units (A and B) of the Ignition system with 115 VAC aircraft power. The EEC controls the fuel flow to the combustion system. The control elements are: S The Metering valve, which controls the rate of fuel flow (the EEC receives feedback from an LVDT), S The PROT MOTOR, which has three positions (STBY, TCM, OVSP), S The MPSOV, which can stop the flow and cause an engine shutdown in case of an overspeed (the feedback is given by the MPSOV switch) S The VSV controller, which supplies fuel to the VSV actuators (the EEC receives feedback from a LVDT located on the VSV actuators).

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In the engine air system, the EEC controls the operation of eight valves: S Three IP 8 bleed valves, S Three HP 3 bleed valves, S The TCC (Turbine Case Cooling) valve. To prevent an engine surge condition, bleed valves controlled by solenoids are independently supplied with electrical power from the EEC. During cruise condition, the EEC fully opens the TCC valve to supply LP compressor air to the external surface of the turbine cases. This causes a smaller clearance between the cases and the tips of the HP and IP/LP turbine blades to increase turbine performance. There is no feedback of the bleed and the TCC valves. The EEC controls the ETRAC (Electronic Thrust Reverser Actuation Controller) through an ARINC 429 BUS. The Thrust Reverser Power Unit sends an inhibition signal to the EEC through the ETRAC. The EEC receives feedback from the TLS (Tertiary Locking System) proximity SNSRs, from the RH and LH side proximity SNSRs and from the RH and LH side cowl resolvers. The EEC controls the hydraulic pump off−load solenoids (channel A for EDP1 and channel B for EDP2) to depressurize the hydraulic system during an in−flight start. The EEC receives feedback from the engine anti−ice protection system for bleed status demand. The EEC channel B only monitors the RAIV (Regulated Anti Ice Valve) position by means of a High Pressure Switch.

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EIPM ARCHITECTURE & INTERFACE DESCRIPTION Architecture There are two EIPMs (Engine Interface Power Management Units) per aircraft, one unit per two engines with dedicated and separated boards and processor per engine. EIPM1: S Board A: ENG 2 S Board B: ENG 4 EIPM2: S Board A: ENG 3 S Board B: ENG 1 The EIPM, installed in the avionics bay, controls and delivers electrical power supply from aircraft towards engine systems. The basic function of the EIPM is to control and monitor the electrical power supply to: S EEC (Engine Electronic Controller), S COS (Cowl Opening System) via the PCPU (Primary COS Power Unit), S P20T20, S ETRAC (Electronic Thrust Reverser Actuation Controller), S EMU (Engine Monitoring Unit), S Ignitors. The EIPM converts N1 analog signal in ARINC 429 bus, for back−up. From a DSI (Discrete Signals Input) group the EIPM generates, for aircraft interface purpose, DSO (Discrete Signals Output) group.

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The EIPM exchanges also data via ARINC 429 with: S The OMS (Onboard Maintenance System) through the SCI (Secure Communication Interface). The OMS and the SCI are in the NSS (Network Server System), S The CDS (Control and Display System) via the CPIOM (Core Processing Input Output Module) of the FWS (Flight Warning System) as a back−up output, S The IOM (Input Output Module), S The EEC Channel A (back−up), S The SSPCs (Solid State Power Controllers) in the SEPDC (Secondary Electrical Power Distribution Center) via the IOM and the ADCN (Aircraft Data Communication Network).

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EIPM INTERFACES Electrical Power Supply Control Logics Each electrical power supply AC2 BUS (EEC, Igniters, ETRAC, P20T20, EIPM) is controlled by SSPCs. In case of EIPM failure or loss, the EEC channels are fail−safe power supplied. In the EIPM each power supply is controlled by relays, which are controlled by the electrical power supply control logics. Interface Control Logics The EIPM proceeds to the control and monitoring of the DSI and DSO groups. The EIPM computes the ”oil low press and ground” signal based on the acquisition and combination of discrete signals from the LGRDCs (Landing Gear Remote Data Concentrator) and Oil Low Press switch, and ARINC signal from IOM (SCI, EEC). The EIPMU sends (via discrete signal) the ”oil low press and ground” signal to other users (IOM, CIDS (Cabin Intercommunication Data System), FCDC (Flight Control Data Concentrator)). The EIPM controls the second line of defense of the Thrust Reverser system, only according to states of inputs of the LGRDC status, reverse switch position, and ARINC bus. This second line of defense is authorized via discrete output RCCB (Remote Control Circuit Breaker) command. The EIPM monitors the T/R second defense line authorization via the RCCB monitoring input. The RCCB Command function is only available for engine 2 and engine 3 (inboard engine). The EIPM controls the power supply of the Cowl Opening System. The fan and Thrust Reverser (or Fan Exhaust) cowl opening is done via the PCPU supplying electrical actuators. The power supply of the COS is only available when the aircraft is on ground and with engine not running. EIPM also uses ARINC data to manage COS application. By default, manual cowls opening is inhibited and carried out by the function of COS. The power supply to the COS is cut in case of action on the ”handful fire−break” of the associated engine.

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When the FADEC (Full Authority Digital Engine Control) ground power P/B is activated, the EIPMU electrically powers the EEC channels for five minutes (maintenance only) if no OMS interactive mode. If the ENG FIRE P/B SW is activated, the EIPM cuts−off the electrical power supply to EEC channels for isolation purpose. The EEC sends the MASTER lever FAULT light (Boolean information) to the EIPM. The EIPM generates a power supply discrete signals to turn the Engine FAULT light on, on the Master Lever. The EIPM acquires N1 speed in an analog form and transmits it via ARINC 429 to the IOM and the FWC. This information is used as a back up information of the N1 speed from the EEC via the ADCN.

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EIPM & FADEC POWER SUPPLY DESCRIPTION General The FADEC (Full Authority Digital Engine Control) has two computers: S EEC (Engine Electronic Controller), S Engine Monitoring Unit (EMU). The power supply of the EEC can be processed into two different manners: S By the airframe power supply (115 VAC) that comes from the EIPM (Engine Interface Power Management), S By the EEC dedicated alternator, also called PMA (Permanent Magnet Alternator). EEC is normally powered by its own power supply (PMA), when engine is running. EMU is supplied by airframe power supply (115 VAC) WARNING:


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DO NOT SET THE MASTER LEVER TO THE ”ON” POSITION WITH THE ENGINE ROTARY SELECTOR ALREADY IN THE ”IGN/START” OR ”CRANK” POSITION. ENGINE RISKS TO BE STARTED OR CRANKED

Airframe Power Supply EIPM is powered in 28 VDC DC1 BUS. The EEC receives power from two airframe 115VAC buses through the EIPM control logic function. The supply line from the emergency bus is connected into channel A of the EEC and the line from the Airframe AC2 bus is connected with channel B of the EEC. In an emergency situation (following loss of all variable frequency generators), only the emergency bus from airframe will operate. The airframe power supply is available on ground and in flight and shall be used by the EEC for its ground tests, ground engine starting, and in flight starting when engine speed is below 8% N3, or in case of PMA failure. EEC Dedicated Alternator Power Supply (PMA) The PMA has a Stator and Rotor that supply two independent three−phase power windings to the EEC (1 per channel). A mechanical drive from the Engine gearbox is used to rotate the PMA Rotor. The interface is required to power the EEC in all Engine running conditions.

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When the engine speed is above 8% N3, the PMA will deliver the electrical power necessary for the EEC to achieve its functions including in−flight starter assist or wind−milling engine starting. NOTE:

Between 5% and 8% of N3 the power supply to the EEC is shared between airframe power and PMA power.

NOTE:

one single phase is also dedicated to N3 sensing.

FADEC Power Supply Aircraft Power−Up At aircraft Power−up or EIPM initialization, the EEC and the EMU will be powered as detailed below: Channel A will be powered if with the airframe 115 VAC Emergency bus is available, Channel A & B and the EMU will be powered for 5 minutes if the full airframe electrical network is available. Engine Mode Selector With engine not running: S When you set the ENG START rotary selector to ”CRANK” or ”STAR/IGN” position, the EEC and the EMU are permanently supplied, S When you set the ENG START rotary selector to the ”NORM” position the power supply of the EECs and the EMU is cut off Engine Master Lever With engine not running: S Each ENG MASTER lever in the ”ON” position supplies permanently the related EEC and EMU. On the ground, airframe 115 VAC will be removed from the EEC and EMU 15 minutes after selection of the ENG MASTER lever from the ”ON” to ”OFF” position. This will not occur in flight.

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Engine FADEC Ground Power For maintenance operation, with the ENG FADEC GrouND PoWeR P/B selected to the ”ON” position and the EEC interactive mode not instigated, airframe 115 VAC will be cut off for 10 minutes. The airframe 115 VAC power will be cut off immediately by selecting the ENG FADEC GrouND PoWeR P/B to the ”OFF” position or by returning the ENG rotary selector to the ”NORM” position (with the ENG MASTER lever to the ”OFF” position). Engine Fire Push−Button In the case of fire, in flight or on ground, airframe 115 VAC power will be cut off immediately following operation of the ENGine FIRE P/B SW. EIPM−Failure In the event of EIPM failure, airframe 115 VAC power will be permanently available to the EEC whenever the airframe electrical network is powered.


EEC Dedicated Alternator Failure If the EEC dedicated alternator winding for the EEC channel in control becomes defective, there will be an EEC channel change over if the second winding is healthy. If both alternators power supply is lost, the EEC will be supplied by the airframe 115 VAC through the EIPM.

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FADEC TEST Tests The OMS (Onboard Maintenance System) is used for the test of two main computers of the power plant system: S EEC (Engine Electronic Controller) S EIPM (Engine Interface Power Management) These tests are launched from the OMS HMI (Human−Machine Interface) using the OMT (Onboard Maintenance Terminal), OIT (Onboard Information Terminal) or GMAT (Ground Maintenance Access Terminal).


EEC To reach the ”TEST SELECTION” page, you must select the ATA and the system to test in the ”ATA SELECTION” page. Then you select the channel in the ”SIDE SELECTION”. The EEC gives the following interactive tests, reports, engine procedures, specific functions:

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Tests AUDIBLE TEST OF THE IGNITERS The EEC cannot detect the operation of the igniter during the test. Even if the test is OK, the result is indicated; you have to make sure that you hear sparks from the ignition system on the engine. VARIABLE−STATOR−VANES SYSTEM TEST The engine will be dry cranked during the test. YOU SET THE CONTROLS AS SPECIFIED IN THE PROCEDURE DISPLAYED ON THE OMS, THE DRY CRANK WILL START IMMEDIATELY. TEST OF THE P20T20 PROBE HEATER CAUTION:

THE P20T20 PROBE WILL BE ENERGIZED FOR 5 SECONDS AND BECOMES HOT DURING THIS TEST. Make sure that not cover, cap or plug is installed on the P20T20 probe. HYDRAULIC PUMP OFFLOAD TEST The engine will be dry cranked during the test. CAUTION:

YOU SET THE CONTROLS AS SPECIFIED IN THE PROCEDURE DISPLAYED ON THE OMS, THE DRY CRANK WILL START IMMEDIATELY. IN THIS TEST YOU MUST LOOK TO SEE IF THE HYDRAULIC PRESSURE INCREASES AND DECREASES AT THE APPLICABLE TIMES. HARNESS TEST This test monitors the Full Authority Digital Engine Control (FADEC) system for 15 minutes and looks for faults while you shake the harness. THRUST REVERSER CYCLING TEST


CAUTION:

NOTE:

Reports The reports are the same for the two channels of the four EEC. EGT EXCEEDANCE REPORT SHAFT−SPEED THE STATUS OF AIRCRAFT HARDWIRED INPUTS Engine Procedures The engine procedures are the same for the two channels of the four EEC. FAN TRIM BALANCE ENGINE CORE WASHING BLEED−VALVE TEST SCHEDULING CAUTION:

THE ENGINE MUST BE STARTED TO PROVIDE THE AIR PRESSURE TO OPERATE THE BLEED VALVES WHEN COMMANDED BY EEC

Specific Functions The specific functions are the same for the two channels of the four EEC. ENGINE RUNNING SIMULATION Engine run discrete signal simulation. The engine is not started for this test RESET FUEL USED

This test is only done onto the EEC of the inboard engines.

THRUST REVERSER WILL BE ENERGIZED AND MOVED DURING TEST. MAKE SURE THAT THE THRUST REVERSER AREA IS CLEAR AND CLEAN OF PERSONS AND TOOLS OR OTHER ITEMS. Make sure that the thrust reverser is not in the inhibited position. Move the throttle lever to reverse idle within 50 seconds, then move the throttle lever to forward idle within 50 seconds. WARNING:

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EIPM To reach the ”TEST SELECTION” page, you must select the ATA and the system to test in the ”ATA SELECTION” page. Then you select the channel dedicated to an engine in the ”SIDE SELECTION” page. The EIPM gives the following interactive tests, reports and Specific function: Tests The tests are the same for each EIPM GROUND POWER LIGHT ENGINE LIGHT FAULT Reports The reports are the same for each EIPM DISCRETE INPUTS REPORTS DISCRETE OUTPUTS REPORTS PIN PROGRAMMING REPORTS Specific Function OIL LOW PRESS AND GROUND On the EIPM 1 ENG 2 and the EIPM 2 ENG 3 there are two other specific functions: THRUST REVERSER 3*115 V / 25 KW POWER SUPPLY REVERSE SECOND LINE OF DEFENSE WILL BE DEACTIVATED; POSSIBLE REVERSE DOORS ACTIVATION CAN OCCUR. ETRAC MANUAL POWER SUPPLY


WARNING:

WARNING:

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ELECTRONIC THRUST REVERSER ACTUATION CONTROLLER (ETRAC) WILL BE POWER SUPPLIED; POSSIBLE REVERSE DOORS ACTIVATION CAN OCCUR.

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FUEL SYSTEM INTRODUCTION The function of the system is to receive fuel from the Aircraft tanks and deliver conditioned metered fuel into the combustion chamber for combustion. The fuel system is divided into: S Fuel Control S Fuel Supply Low Pressure Pump (LPP) The LPP is a single stage centrifugal pump that receives fuel from the Aircraft system and ensures satisfactory pressure to both HP pump inlets. Fuel Oil Heat Exchanger (FOHE) The heat exchanger is used to transfer heat between the engine oil and the LP fuel from the LPP.

Fuel Flow Transmitter The transmitter provides a signal of fuel flow to the EEC for onward transmission to the cockpit for display. HP Fuel Filter Located in the inlet to the fuel manifold to prevent blockage of the fuel spray nozzles. The HP fuel filter is a cleanable 250 micron filter Fuel Manifold & Fuel Spray Nozzles (FSN) The fuel manifold is an assembly of flexible hoses at equal distances around the combustion outer case. The manifold distributes the fuel to the 20 FSNs that provide the necessary atomisation of fuel into the combustion chamber.

LP Fuel Filter The filter removes contaminants from the LP fuel before it passes to the HP system. It is a 40 micron non−cleanable filter. A differential pressure transducer (set at 5psid) provides an indication of impending filter blockage. A by−pass valve operates at 25 psid to allow unfiltered fuel to the HP pump.


Main High Pressure Pump (HPP) The HP pump is a spur gear type pump and provides high pressure fuel to the HMU. Servo High Pressure Pump (HPP) The Servo HP pump is a spur gear type pump and provides high pressure fuel to the HMU for use in the Variable Stator Vane (VSV) system. Hydromechanical Unit (HMU) The HMU interfaces with the EEC to control fuel flow for all normal & emergency conditions through the fuel metering, overspeed and fuel shut−off valves. The shut−off valve can also be operated by electrical signals from the fuel control switch in the cockpit

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FUEL SYSTEM SCHEMATIC & CONTROL Direct Control Inputs The engine master switch outputs to the following: S ON/OFF command directly to the LP Shutoff Valve S ON/OFF command to the IOM‘s S OFF command only to the airframe shutdown solenoid in the HMU. The master switch input to the IOM is sent to both channels of the EEC via the ADCN. The throttle resolvers input the throttle position to each channel of the EEC. The auto thrust system also inputs thrust requirements to the EEC when auto thrust is selected. The PRIMs also input to the EEC to provide a discrete signal when parameters allow engine shutdown, during a thrust control malfunction (TCM). Control Outputs The EEC controls the engine fuel system through the HMU using the following devices: S Metering valve servo valve S Protection Torque Motor S VSV Controller Fuel System Inputs to the EEC The HMU has the following feedback devices: S An LVDT which provides positional feedback to the EEC of metering valve position S A dual proximity probe providing Minimum Pressure Shut Off valve (MPSOV) position to the EEC. A differential pressure (dP) transducer provides an indication of partial blockage of the LP fuel filter. The Fuel Flow Transmitter provides a mass flow indication to the EEC.

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Figure 84 FRA US/T

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FUEL PUMP Purpose The fuel pump receives fuel from the Aircraft and pressurises it sufficiently to ensure: S Adequate pressure for fuel powered actuators S Good atomisation of fuel at the FSNs


Location The low pressure fuel pump and the high pressure fuel pumps (2) are combined as one single assembly mounted on the rear face of the external gearbox Description Fuel from the aircraft flows into the inlet of the centrifugal impeller type LP pump. The LP pump compresses any fuel vapour back into solution and increases the fuel pressure by centrifugal action to approximately 175 psid (at max speed). The LP fuel is supplied to the FOHE. The LP pump also supplies fuel through a filter to the fuel drains tank ejector. There are two high−pressure fuel pumps (HPP) which are both positive displacement spur gear type pumps. The main HP pump increases the main fuel pressure to approximately 1725 psid. (at maximum speed) and supplies fuel to the HMU for fuel delivery and control. The servo HP pump increases the fuel pressure to approximately 1825 psid. (at maximum speed), and supplies fuel to the HMU for use by the Variable Stator Vane (VSV) system. A full flow relief valve is fitted within the pump to prevent overpressurising the pump casing, which opens at 2250 psid. The relief valve returns HP fuel back to the HP pump inlet. There is also a relief valve on the servo pump, which limits maximum pressure to 2350 psid. The pump is driven from the external gearbox and is bolted to the rear face of the external gearbox.

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Figure 85 FRA US/T

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FUEL OIL HEAT EXCHANGER (FOHE)

LP FUEL FILTER

Purpose To transfer heat from the engine oil to the fuel to prevent ice formation in the fuel.

Purpose To remove contaminants from the fuel before passing into the high−pressure system.

Location The FOHE is mounted on the right side of the fan case below the oil tank.

Location Attached to the FOHE assembly and mounted on the right side of the fan case.

Description Heat is transferred from the oil to the fuel in the core of the FOHE. The oil flow is made slower by many baffle plates around the steel tubes through which the fuel is flowing. The slower oil flow enhances the exchange of heat between the oil and fuel. The LP fuel filter housing is part of the same LRU and fuel flows directly from the FOHE into the fuel filter.

Description The LP fuel filter is a 40−micron non−cleanable element located in a cylindrical aluminium cast housing. The filter housing is connected by a bolted flange midway along it is length, to the FOHE housing. The filter is held in place in the housing, sealed by a spring−loaded pressure plate reacting against the filter housing end cap that is bolted in position. In the event of a partial blockage of the filter, a differential pressure transducer (5 psid.) will provide a cockpit indication. If the filter becomes blocked, a by−pass valve opens at 25 psid to allow unfiltered fuel through to the HP pump.




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Figure 86 FRA US/T

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HYDROMECHANICAL UNIT (HMU) The HMU is installed on the rear face of the external gearbox. Fuel flow between the pump and HMU is via internal passages in the gearbox. Purpose To control the fuel flow to the fuel spray nozzles and combustion chamber from electrical inputs received from the following: S EEC S Overspeed Protection System (OPS) S Cockpit engine master switch.


Location Bolted to the rear face of the external gearbox.

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HYDROMECHANICAL UNIT (HMU) Operation The EEC controls four servovalve torque motors in the HMU, which gives the following functions: S Fuel Metering Valve control S Fuel high−pressure (HP) control S Overspeed & protection S Fuel shut−off control S VSVA control The HMU works on a constant pressure drop principle and varies the fuel flow to the combustion chamber by varying a bypass return flow back to the inlet of the HP pump. The Pressure Regulating Valve (PRV) senses any changes in the pressure drop across the metering valve and opens or closes to maintain a constant pressure drop.

Protection Torque Motor The protection torque motor is operated by the EEC in the following circumstances: S During normal engine shutdown on the ground S N1 and N2 overspeed S LP turbine overspeed S Thrust Control Malfunction (TCM) BITE checks are also carried out during engine start and shutdown as follows: S TCM BITE carried out during start S Overspeed BITE carried out during shutdown During a TCM the fuel flow can be reduced or engine shut−down commanded. The shutdown permission is provided by the aircraft PRIM.

Metering Valve Controlled by the EEC via electrical inputs through the metering valve torque motor. A change in the metering valve position changes the pressure drop, which is sensed by the pressure regulating valve (PRV) to effectively change the fuel flow to the combustion chamber by changing the bypass return flow. The LVDT on the metering valve provides positional feedback to the EEC.

VSVA Controller Torque Motor See Section 8.

Airframe Shutdown Solenoid The airframe shutdown solenoid is energised when the MASTER switch is moved to the OFF position. This changes the reference pressure at the PRV, which reduces the pump discharge pressure and the metered fuel pressure reduces to a low value and the Minimum Pressure & Shut−Off Valve (MPSOV) spring forces the valve closed, resulting in drop tight shutoff of fuel flow to the engine. A dual proximity probe provides MPSOV position to the EEC.

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Figure 88 FRA US/T

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HMU REMOVAL/INSTALLATION (AMM 73−21−52) ATTENTION: Warnings & Cautions Make sure you observe all the applicable warnings and cautions given in the AMM HMU Removal (73−21−52−000−801) The AMM procedure is briefly described below: S Open the fan cowl doors S Drain the fuel from the drain point on the FOHE S Disconnect the electrical connectors on the HMU and put blanking caps on the harness connectors and the HMU. S Put a container in place to collect the remaining fuel S Disconnect the fuel tube connections at the HMU. S Support the weight of the HMU and remove the bolts securing the HMU to the external gearbox module. S Carefully remove the HMU from the gearbox. S Blank all remaining openings.

HMU Installation S Remove the blanks from the HMU and external gearbox S Examine the HP fuel main outlet, HP fuel servo outlet and the LP spill inlet in the external gearbox module. Make sure these openings are clean and clear of unwanted objects. S Install new seal rings in the grooves on the external gearbox module S Put the HMU in position and loosely install the bolts and washers. S Torque the bolts in the sequence given in the AMM S Connect the fuel tubes to the HMU and torque the connectors S Remove the blanking caps and connect the electrical connectors S Carry out a leak check of the HMU installation and FOHE drain point S Put the aircraft back to its initial configuration

The seals located in the grooves on the rear face of the external gearbox can be difficult to remove. Do not insert sharp objects into the groove as this can cause damage and subsequent leaks from the mating face. The correct method of removing the seal is to carefully lift the edge of the seal using special tool (TBA) then pull the seal out of the seal groove with a pair of pliers. S If you think there has been a release of material from the HMU, clean the HP fuel filter (73−11−42−100−801)


NOTE:

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Figure 89 FRA US/T

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HMU SHUTDOWN SEQUENCES There are 3 independent methods of shutting the engine down as follows: Normal Shutdown When the MASTER lever is moved to the OFF position the AF shutdown solenoid is energised. The pressure downstream of the metering valve is ported to return. The Minimum Pressure & Shutoff Valve (MPSOV) spring forces the valve closed, resulting in drop−tight shutoff. After shutdown is commanded, the EEC recognises that shut off has occurred and signals the Metering Valve (MV) Torque Motor (T/M) to latch the MV in the shutoff position. If the aircraft is on the ground, EEC commands the Protection T/M, which ports HP fuel to the spring side of the MPSOV and also to the manifold drain valve. The high pressure signal puts the drain valve to the drain position, which remains in the open position until the HP fuel signal from the Protection Motor is removed. Note: If shutdown occurs in the air, the EEC does not energise the protection motor and the manifold drain remains closed.

Thrust Control Malfunction (Fuel Reduction) The EEC protection system, utilising the aircraft input, commands the Protection torque motor to the TCM position (−40 mA). This provides a spill of fuel back to HP pump return via the TCM orifice. This results in the pump discharge pressure now being referenced to a combined spill from both the Pressure Regulating Valve and the TCM orifice. The greater spill causes a reduction in fuel flow.

Unusual Event Protection Shutdown The protection motor is commanded to its overspeed position by the EEC for any of the following conditions: S LP compressor shaft overspeed S IP compressor shaft overspeed S LP shaft failure S Thrust Control Malfunction Movement of the protection motor to its overspeed position causes a servo port to open. The MPSOV spring will then close the valve and shut off fuel to the engine. This also commands the manifold drain valve to Open. When the MPSOV closes and the manifold drain valve opens, it lets the remaining pressure in the engine force the HMU fuel discharge into the drains tank. This will cause rapid shut down of the engine. For Thrust Control Malfunction, the EEC must receive the aircraft permission discrete input, in order for shut down to occur. If the shutdown discrete is not received from the aircraft a fuel reduction will be activated as described below. NOTE:

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NORMAL SHUTDOWN Step 1

Step 2

AF Shutdown Solenoid Engised. MV Downstream pressure opend to spill. MPSOV spring forced vale closed.

Step 3 - (On Ground only)

EEC recognises S/D has occured. Selects MV T/M to Shutoff position. MV latched in shutoff postion.

EEC commands Protection T/M and HP fuel ports to spring side of MPSOV conforming closure and to Manifold Drain valve to position open

THRUST CONTROL MALFUNBCTION (FUEL REDUCTION)

OVERSPEED SHUTDOWN

EEC commands Protection T/M to TCM position (-40 mA): Flow path opened to spill via TCM orifice. Pump discharge pressure noe referenced to combined spill between PRV and TCM orifice.


EEC Commands Protection T/M: Ports MV downstream pressure to spill. Ports HP fuel to spring side of MPSOV. MPSOV closes rapidly.

Reduced in Fuel Flow.

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FUEL FLOW TRANSMITTER Purpose To provide fuel flow and fuel usage indications in the cockpit. Location The fuel flow transmitter is located in the fuel line between the HMU and HP fuel filter and is attached to brackets on the bottom of the fan case at the rear at approx 5 o‘clock position viewed from the rear.


Description The transmitter sends analogue pulse signals to the EEC that are in proportion to the engine mass fuel flow rate. The flowmeter is connected to one channel of the EEC and crosswired between channels. The EEC uses the signals to calculate the engine mass fuel flow and fuel usage and sends this data on the ARINC 429 data bus for display on the cockpit System Display (SD).

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Figure 91 FRA US/T

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HP FUEL FILTER Purpose To filter HP fuel prior to entry into the primary fuel manifold Location Attached to the inlet of the fuel manifold, on the core engine at the underside.


Description The filter is a 250 micron element housed in a cast casing. The element is secured in the casing by a retained bolt. The element is re−usable.

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Figure 92 FRA US/T

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FUEL MANIFOLD Purpose To deliver HP fuel to the fuel spray nozzles. Location Fitted around the combustion outer case.


Description The fuel manifold is an assembly of flexible hoses at equal distances around the combustion outer case. The manifold distributes the fuel to the 20 FSNs that provide the necessary atomisation of fuel into the combustion chamber. The fuel manifold assembly is divided into 5 parts: S The inlet manifold S The right−hand rear fuel manifold S The right−hand forward fuel manifold S The left−hand rear fuel manifold S The left−hand forward fuel manifold The inlet manifold has the fuel inlet from the HP fuel filter and the other manifolds are connected to the inlet manifold. When the engine is shutdown (or does not start) on the ground, the fuel in the manifold is drained back through the inlet connection. This allows fuel to drain from the manifold via the HMU to the drain tank. Fuel is not drained from the fuel manifold when the engine is shutdown in the air.

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Figure 93 FRA US/T

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FUEL MANIFOLD INSPECTION (AMM 73−11−43−200−802) Observe all Warnings and Cautions given in the AMM. Procedure: The procedure is fully covered in the AMM and briefly described below. S On the OMT, get access to the Power Distribution Control management and open, safety the applicable circuit breakers. S On the inboard engines make sure that the thrust reverser is made unserviceable for maintenance. S Open the Fan cowls and fan exhaust cowls S Clean the fuel manifold (AMM 70−20−01−100−801) S Dry the fuel manifold with a lint free cloth S Using a light source and mirror examine the fuel manifold assemblies. Examine the following areas: S Fuel manifold tube brackets, clips and their related nuts and bolts. Replace damaged, loose or worn parts. S Fuel manifold tube end connectors and unions. Refer to the AMM for applicable limits on nicks. S The condition of the silicon on the fuel manifold fire sleeve for the following damage: − Torn − Split − Cut − Cracked or missing material − Chafed (where the woven fibreglass fire sleeve cannot be seen - accept) S Torn, Split, Cut and Chafed silicon on the fuel manifold sleeve where the woven fibreglass sleeve can be seen − reject and replace the applicable section of the fuel manifold. S On completion return the aircraft back to its initial configuration.

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Figure 94 FRA US/T

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FUEL SPRAY NOZZLES (FSN) Purpose To deliver the correct fuel/air mix to the combustion chamber. Location Fitted through openings in the outer combustion case into the head of the combustion chamber. Description There are 20 fuel spray nozzles (FSNs). They are cast body fabrications of simplex air spray design. Fuel is delivered to the FSN then through the body (feed arm) to the swirl chamber head for atomisation and air mix before entry into the combustion chamber. The fuel enters the swirl chamber and is partially atomised, HP compressor delivery air passes into the rear of the swirl chamber mixing with the swirling fuel. The air/fuel is swirled further by a series of vanes before exiting the swirl chamber nozzle.


Weight Type Distributors The weight type distributor valve fits inside the feed arm to control the individual fuel delivery pressure, to match all the FSNs output during low flow conditions i.e. engine start. There are two different weight assemblies in the distributor vales installed. S Two weight assemblies in position 8 and 12 with identical weight and S 18 other weight assemblies in the other positions which have all the same weight.

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DISTRIBUTOR WEIGHT

WEIGHT ASSEMBLY LOCATIONS 8 AND 12


WEIGHT ASSEMBLY 18 LOCATIONS (DOES NOT INCLUDE LOCATIONS 8 AND 12)

Figure 95 FRA US/T

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LP FUEL FILTER REMOVAL/INSTALLATION (AMM 73−11−41) ATTENTION: Warnings & Cautions Make sure you observe all the applicable warnings and cautions given in the AMM Removal Procedure S The procedure is described in AMM 73−11−41 and briefly described below: S On the OMT, get access to the Power Distribution Control management and open, safety the applicable circuit breakers. S Open the right fan cowl door S Drain fuel from the drain point on the Fuel Oil Heat Exchanger into a container S Loosen and disengage the captive bolts from the fuel filter housing and carefully lower the end cover with the filter element together.


NOTE: The four bolts and washers stay attached to the end cover. S Discard the filter element S Remove and discard the seal ring S Put a cover on filter housing S Installation Procedure S Remove the cover from the filter housing S Inspect the inner area of the filter housing and make sure it is clean and clear of unwanted material Install a new seal ring in the groove on the end cover S Put a new filter element into position on the end cover. Make sure that the bonded seal at the end of the filter element, engages with its location in the end cover S Carefully install the filter element and end cover into position in the filter housing. Attach the end cover to the filter housing with the captive bolts S Torque the captive bolts to the value in the AMM S Close the applicable circuit breakers through the OMT. S Carry out a leak check of the filter installation and drain point S Return the aircraft back to its initial configuration FRA US/T

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Figure 96 FRA US/T

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INHIBIT THE ENGINE FUEL SYSTEM If the aircraft or the engine is to be placed into storage for periods in excess of 30 days, then the engine fuel system may be required to be inhibited to protect the fuel system internal components. The storage conditions are given in the AMM Task 71−00−00−600−802−A.

5. 6.

7.

8.


Procedure: The procedure is given in AMM Task 71−00−00−600−806 and briefly described below. Follow all Warnings & Cautions given in the AMM for your own safety and the safety of others. 1. Using the Onboard Maintenance Terminal (OMT) get access to the Power Distribution Control Management pages and open, safety/lock and tag the applicable circuit breakers as shown in the AMM. 2. Get access to the engine and open the fan cowl doors. A. On the inboard engines make sure that the thrust reverser is made unserviceable for maintenance. 3. Drain the fuel from the following engine fuel system component drain points: A. a) FOHE, fuel pump, HMU, HP fuel supply, VSV extend and retract tubes 4. Inhibit the engine fuel system: A. Disconnect the engine LP fuel supply tube at the pylon (Let the fuel drain into a clean container). B. Connect the adapter HU41792 to the engine LP fuel supply tube. C. Prepare the inhibiting rig

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YOU MUST NOT PRESSURIZE THE RESERVOIR ABOVE 100 PSIG AS THIS CAN CAUSE DAMAGE AND INJURY TO PERSONS. Connect the delivery hose to the adapter on the LP fuel supply tube. Install new seal rings on the drain plugs removed in 3. (Note: Do not install a new seal ring on the HMU drain plug as it will be necessary to install this temporarily). Supply mineral oil (OMat 1024) to the fuel system. A. Make sure the pressure gauge on the shows 50 psi (3.45 bar) and with clean containers below the drain plugs, open the shut off valve on the rig. B. When fuel free mineral oil flows from the drains, install the drain plugs in the following order: C. FOHE, fuel pump, HMU, HP fuel supply & VSV fuel supply tube drain points D. Wait 30 secs for the VSV actuators to fill then close the shut−off valve on the rig. E. Remove the drain plug on the HMU and install the adapter HU80277 and connect the other end of the adapter to the HP fuel supply tube drain point F. Open the shut−off valve on the rig and supply mineral oil to the fuel system until a fuel free flow comes from the LP turbine drain tube. Disconnect the adapters and install the LP fuel supply tube and drain plugs. Install new seal rings, torque load and safety where necessary all disconnected points.

WARNING:

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ATA 77 ENGINE INDICATING ENGINE & FADEC SYSTEMS OPS/CTL & IND (RR) General Let’s see the general view of the A380 cockpit.




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Figure 98 FRA US/T

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Engine Control Panels Location The engine control panels are located on the overhead panel: S The EIPM 1 & 2 reset circuit breakers, S The ENGINE FADEC GND PWR panel, S The ENGine FIRE panel, S The ENG MANual START panel with the ALTN MODE P/BSW, S The ENG START panel. An A/THR (autothrust) P/B is located on the glareshield and on the pedestal there are: S The THROTTLE CONTROL LEVERS with the instinctive disconnect P/Bs and the reverser levers on the engine 2 and 3 only, S The ENG MASTER levers panel.




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Figure 99 FRA US/T

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EEC Powering / Depowering By setting the ENG START selector switch to the CRANK or IGN START position with engines not running the FADEC is powered. The corresponding indication is clearly displayed on the EWD and the amber ”XX” which were displayed due to the absence of information are replaced by the main thrust parameters. Also when you press the FADEC GND PWR P/BSW to ON, the FADEC parameters are clearly displayed on the EWD and the amber ”XX” are replaced by the main thrust parameters. If the FADEC GND PWR P/BSW remains ON, this means that the FADEC is powered for 10 min, except if OMS in interactive mode. In this case the FADEC stays automatically energized as long as you are in the related EEC tests menu. After these 10 min, on FADEC GND PWR panel, the ON legend goes off automatically. The MASTER LEVER could be used for powering the FADEC. In this case, the dedicated FADEC will be powered for 15 minutes, but it is not recommended on A/C, because there is a risk of the engine to be started if the rotary selector has been forgotten in IGN START position. If you do so, observe quick ON and OFF action, because you do not have to forget that when the MASTER LEVER is set to ON the LP fuel SOV is controlled to open. If the EIPM reset switches are pulled for maintenance purposes, the dedicated EECs are fail−safe powered. The ENGINE FIRE P/BSW also has effects on the EEC powering/depowering. When this P/BSW is released out, it has effects on the various aircraft systems such as: S FADEC power supply is cut, S The hydraulic system: closure of the fire shut−off valves in order to isolate the hydraulic pumps from the reservoir, S The electrical system, S The bleed system, S The fuel system: closure of the engine LP valve.


NOTE:

CAUTION:

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WITH ENG FIRE PUSH BUTTON RELEASED OUT, THE FIRE EXTINGUISHERS ARE ARMED. DO NOT PRESS ON THE AGENT P/BSWS.

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Figure 100 FRA US/T

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Engine Parameters Display Among the main parameters displayed on the EWD, you have the thrust mode indication (CLB, MCT, TO). The N1 and EGT indications are given as numerical values. The ECAM ENGINE page can be called by the selection of the ENG key on the ECAM Control Panel (ACP) located on the pedestal. The secondary engine parameters, which are shown on the ECAM ENGINE page, are: S IP rotor speed (N2), S HP rotor speed (N3), S Fuel Flow (FF), S Engine oil quantity (OIL QTY), S Engine oil temperature (OIL TEMP), S Engine oil pressure (OIL PRESS), S Engine rotor vibration levels (N1, N2, N3), S Nacelle temperature (NAC) from 0 to 500_C.




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Figure 101 FRA US/T

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Throttle Control Levers The Throttle Control Levers can be individually moved and manually only, they are used to adjust the aircraft forward thrust. The range of the levers movement is divided into 4 detent points: 0 (IDLE), CL for climb, FLX/MCT and TO/GA. When you supply an engine FADEC for example, by pushing the FADEC GND PWR P/BSW to ON and when you move the corresponding Throttle Control Lever, you will see the displacement of the cyan circle on the EWD THR indicator. The thrust mode corresponding to the lever position is also displayed on the upper part of the EWD. On the Throttle Control Leverl, are the reverser levers to control the deployment or the stowing of the reversers and adjust the reverse thrust. These thrust reversers are installed on the engine 2 and 3 only. The reverser levers move between two detents: IDLE and full reverse thrust. Two A/THR instinctive disconnect P/BSWs (red) are located on the Throttle Control Levers (engines 1 and 4 only). They direct input to all EECs in order to disconnect the A/THR function as soon as one of them is pushed.


A/THR P/B An A/THR P/B is located on the FCU (Flight Control Unit) section of the glareshield. When it is pressed, three green lines on this P/BSW illuminate. The green lines go off when any instinctive disconnect P/B is pressed or when you press the A/THR P/B again.

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Figure 102 FRA US/T

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POWER PLANT AIR

A380 75

ATA 75 ENGINE AIR SYSTEM ENGINE AIRFLOW CONTROL INTRODUCTION Description The engine compressor system is designed to produce high pressure ratios in the higher RPM range in which the engine normally operates. In the lower RPM range the airflow through the IP and HP compressors becomes unstable especially during acceleration and deceleration. It is therefore necessary to have airflow control devices to provide stable compressor airflow, during starting and lower power operation. The EEC controls the airflow control system. IP Compressor Airflow Control The IP compressor airflow control system consists of: S Variable Inlet Guide Vanes (VIGVs) at the inlet to compressor S Two stages of Variable Stator Vanes (VSVs) S Three bleed valves at stage 8 The VIGVs and VSVs control the angle of the air supplied to the first three stages of the IP compressor. The angle of the VIGVs and VSVs is changed to adapt to different conditions of compressor operation and helps to prevent compressor stall/surge conditions.


HP Compressor Airflow Control The HP compressor airflow control system consists of three bleed valves at stage 3. IP & HP Bleed Valves At lower engine speeds the bleed valves are open bleeding some of the compressor airflow into the by−pass duct to prevent stall/surge conditions. The bleed valves are closed at higher engine speeds to provide full airflow through the IP and HP compressors. All the bleed valves are two position valves only and are either open or closed.

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VIGV/VSV CONTROL SYSTEM Operation The IP compressor VIGV/VSV system consists of the following units: S VIGV/VSV Control Valve in the HMU S Two VIGV/VSV Actuators S VIGV/VSV Actuating Mechanism The EEC uses IP shaft speed, IP shaft acceleration, LP shaft speed, altitude and T20 to schedule the VSV position. When these conditions change during acceleration or deceleration the EEC will send a signal to the VIGV/VSV controller in the HMU. The controller responds by directing HP servo fuel pressure to position the actuators. The actuators are extended at low speeds and retract as the IP shaft speed increases The two actuators are each connected to a crankshaft assembly located at the 3 o‘clock & 9 o‘clock positions. From each crankshaft assembly, there are 3 output rods to the VIGV and VSV unison rings. The unison rings are connected to the vane operating levers and as the unison ring moves, they change the angular position of each vane. Linear Variable Differential Transducers (LVDT), located inside the actuators, send signals back to the EEC confirming the position of the actuators.

Engine Shutdown When the engine is shutdown, the EEC directs the VSV system to open the VIGVs & VSVs. This allows the IP rotor to be turned without the requirement to disconnect fuel lines and manually move the vanes to the open position. During start, the EEC returns the VSV system back to the normal operating schedule.


Transient Control The control system modifies the actuator position schedule through all transient conditions, including engine acceleration, deceleration, reverse thrust operation and in the unlikely event of engine surge, transiently increasing handling margins. Failsafe Control If the IP shaft speeds are not available, control is attempted using LP shaft speed, whilst the fuel control system brings the engine speed to idle. If a failure of the electrical supply occurs, the system is designed to retract the actuators to the high−speed position. This minimises the risk of an overspeed event that might otherwise occur if the actuators extend too fast relative to fuel flow control.

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VIGV/VSV ACTUATORS Purpose The two identical actuators provide the muscle force to move the VIGV/VSV mechanism to required position. Location Mounted on brackets attached to the IP compressor case and the intermediate casing at approximately the 3 & 9 o‘clock positions. The left actuator is just above the engine centreline, the right actuator just below the engine centreline.


Description The actuators are powered by fuel pressure from the VSV actuator control in the HMU. There are fuel lines to the extend and retract sides of the actuator. There is also a fuel drain line to collect fuel that leaks past the actuator seals. Each actuator has a single channel LVDT that provides a signal of actuator position to the EEC. The left actuator LVDT provides a signal to channel A, the right actuator provides a signal to channel B. The EEC channel in control only uses the input from it is own LVDT. If that signal is lost, it will then use the input signal from the other channel.

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COMPRESSOR BLEED VALVE SYSTEM IP / HP Bleed Valve Control Solenoids The EEC controls the bleed valves through the bleed valve solenoid units (2). The two units are located on the upper left and right sides of the IP compressor case and control the bleed valves on their respective sides. Each solenoid unit has an HP3 air inlet connection and two electrical connectors to the EEC Channels A and B. Each solenoid is dual wound, with control from each EEC channel. IP Bleed Valves There are three poppet type IP compressor stage 8 bleed valves, each controlled by a solenoid valve switching HP3 servo air. The bleed valves are spring−loaded in the normal open position with the engine not running. After starting, with the engine at idle or above, sufficient air pressure will build up at the valve inlet to close the bleed valves against the spring force. When the solenoid is energised, the servo air system is pressurized and the bleed valve opens. When the valve solenoid is de−energized, the servo air system is vented and the bleed valve closes. HP Bleed Valves There are three poppet type HP compressor stage 3 bleed valves, each controlled by a solenoid valve switching HP3 servo air. Operation of the control valve solenoid and bleed valve is the same as that for the IP bleed valves.

Failsafe Control If a failure of the electrical supply occurs, the system is designed for the handling bleed valves to automatically close (high speed position). This failure mode makes sure that the engine internal air pressure distribution does not adversely affect turbine cooling.

All the bleed valves bleed air into the by−pass duct when in the open position.


NOTE:

Fault Annunciation The EEC can carry out continuity checks between the EEC and the bleed valve controllers and will set a fault message for failure of continuity. However, there is no feedback to the EEC to confirm that the bleed valve has operated correctly. If a bleed valve is not operating it will show itself by either of the following: Valve open when it should be closed this will bleed air from the compressor at the higher rpm range and will show an increase in TGT. This may be observed by the aircrew, but will certainly show itself on condition monitoring as a step change. Valve closed when it should be open - this is likely to show itself during starting with a tendency to cause hung/hot starts. A bleed valve scheduling test can be carried out on the ground, with the engine running at idle. The EEC commands each of the bleed valves open & closed and reports any faults by monitoring changes in engine conditions.

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BLEED VALVE SOLENOIDS Purpose To control the opening and closing of the three HP3 and three IP8 bleed valves on commands from the EEC. Location One four−solenoid unit is installed on the left side of the IP compressor case. The other four−solenoid unit is installed on the right side of the IP compressor case. Access by opening the C−ducts and removing the left and/or right upper core fairing.


Description The two bleed valve solenoid units consist of eight independently operated solenoid valves in total. (3 for the IP8 bleed valves, 3 for the HP3 bleed valves, 1 for the turbine case cooling valve and 1 for the NAI shut−off valve. There is one pneumatic connector (HP3) and two electrical connectors on each unit, these supply electrical power and air to the solenoids. Each solenoid has two coils, one is connected to EEC channel A, the other to channel B. The outlets pneumatically connect the solenoids with the bleed valves.

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CHANNEL A & B ELECTRICAL CONNECTORS


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IP AND HP BLEED VALVES Location The three IP8 and three HP3 bleed valve locations are shown opposite. The IP8 Bleed Valves are numbered and positioned as follows viewed looking forwards: S No.1 − Top right S No.2 − Bottom right (blanking plate fitted to casing) S No.3 − Bottom left S No.4 − Top left The HP3 Bleed Valves are numbered and positioned as follows viewed looking forwards: S No.1 − Top right S No.2 − Bottom right S No.3 − Bottom left



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COOLING & SEALING INTRODUCTION The engine is internally cooled with air supplied by the IP and HP compressors. This air is also used to seal bearing chambers to prevent internal leakage of oil. Air that is supplied by the IP compressor is taken from stages IP5 and IP8. Air that is supplied by the HP compressor is taken from stages HP3 and HP6. Parts of the engine, which are at different pressures, are isolated from each other by labyrinth seals. The temperature of the cooling air around the IP turbine disc is monitored by the turbine overheat detection system.



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VARIABLE STATOR VANES SYSTEM TEST CAUTION:

THIS TEST IS CARRIED OUT WHEN THE ENGINE IS BEING DRY CRANKED.

Description The test is carried out through the On−board Maintenance Terminal (OMT). The engine is dry cranked to provide sufficient fuel pressure to move the VSV actuators when commanded by the EEC. The EEC commands the VSV actuators to move between the closed and open positions and monitors the feedback signal from the LVDT‘s (Linear Variable Differential Transducer) in each actuator to ensure they move to the commanded position within a specified time. The test takes approximately 90 secs. If the N3 speed does not increase to a satisfactory value in 60 secs, the test is aborted. During the test a dry crank time indicator appears on the OMT screen and this is replaced by a time indicator for the test, once the N3 speed has reached a satisfactory value. On completion, a message appears on the screen to command: S Rotary selector switch to NORM position S Engine Manual Start pushbutton to OFF S Remove the starter air The Test result is then displayed. If the test failed then you should check for related Maintenance Messages.

Procedure: Obey the Instructions shown on the Onboard Maintenance Terminal Make a decision on which EEC channel you need to set during the test. Make sure the ENG MASTER lever is set to off Make sure the ENG START rotary selector is in the NORM position On the OMT, on the OMS home page, select: A. SYSTEM REPORT/TEST B. ATA 73 C. EEC D. The applicable EEC Start the test On completion of the test, put the aircraft back to its initial configuration.

NOTE: 1. 2. 3. 4.

5. 6.



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BLEED VALVE TESTS SCHEDULING CAUTION:

THIS TEST IS CARRIED OUT WHEN THE ENGINE IS BEING GROUND RUN AT GROUND IDLE POWER.

Description The test is carried out through the On−board Maintenance Terminal (OMT). The engine is started to provide the air pressure to operate the bleed valves when commanded by the EEC. The test can be enabled for HP bleed valves only or can be enabled for both the IP & HP bleed valve scheduling test. The EEC commands each bleed valve in turn (HP only or IP&HP valves) and monitors the engine parameters during the test. If the EEC does not see a change in engine parameters when the bleed valve is operated between open & closed, then a fault message will be set.



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TURBINE CASE COOLING SYSTEM (TCC) Purpose The turbine case cooling system uses fan air to cool the HP, IP and LP turbine cases to maintain the turbine casings within satisfactory temperature limits. It also controls the HP, IP and LP turbine casing thermal growth and consequently controls the turbine blade running tip clearances, which improves the turbine efficiency.


Location of Units The solenoid control valve is part of the Bleed Valve Solenoid pack on the left side of the intermediate case in Zone 2. The TCC Valve assembly is located on the left side of the engine at the horizontal centreline, to the rear of the PRSOV in the air offtake system in Zone 3.

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TURBINE CASE COOLING SYSTEM (TCC) Description The TCC valve assembly is a single line replaceable unit (LRU) interfacing with the fan bypass air offtake and turbine cooling manifolds. The valve is a poppet type similar to the IP and HP handling bleed valves and is controlled by a solenoid valve switching HP3 servo air. Operation The TCC valve is spring−loaded in the closed position. When the solenoid is energized, the servo system is pressurized and the bleed valve opens. When the control valve solenoid is de−energized, the servo air system is vented and the spring force closes the TCC valve. The control valve solenoid is connected to both channels of the EEC and is driven by a signal from either channel A or B. In Take−Off Conditions The TCC valve is in the closed position preventing flow to the HP and IP turbine cooling manifolds. However, a smaller amount of fan bypass air is still allowed to flow into the LP turbine cooling manifold, thereby maintaining cooling to the LP turbine case.


In cruise conditions The TCC valve is opened allowing fan bypass air to flow into the HP, IP & LP turbine cooling manifolds. This contracts the combustion outer case, HP/IP turbine case and LP turbine case, reducing HP, IP and LP turbine blade tip clearances and thereby maintaining engine performance. Failsafe Control If a failure of the electrical supply occurs, the system is designed for the TCC solenoid to be de−energised, the servo line vented and the TCC valve closed by spring pressure.

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TCC MANIFOLD AND COOLING DUCT Description The manifold and cooling duct assembly consists of the following items: S manifold inlet (HP and IP turbine) S manifold − 2 halves S 4 access panels S 4 sections of LPT cooling duct



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TURBINE OVERHEAT DETECTION SYSTEM Purpose The turbine overheat detection system monitors the temperature of the HP3 cooling air at the front and rear of the IP turbine, and the IP8 cooling air at the rear of the seal panel. A “TURBINE OVHT“ warning occurs on the E/WD in the cockpit if the temperature is more than the overheat limit. Location The two thermocouple probes are located on the IP turbine case. The front thermocouple assembly fits through the inside of one of the IP nozzle guide vanes and is located to the left of top dead centre. The rear thermocouple fits through the inside of one of the LP1 nozzle guide vanes and is located on the right of bottom dead centre. Description The turbine overheat detection system has two thermocouple probes which monitor the temperature. Each probe has two thermocouple elements, one sends a signal to EEC channel A, the other sends a signal to channel B. If the temperature is more than the overheat limit, the EEC sends a signal to the Aircraft via the AFDX outputs. The EEC will send a signal to the Aircraft when: S Both elements in the same thermocouple indicate the overheat limit S One element indicates the overheat limit and the other element in the same thermocouple has a fault.


NOTE:

If one element in the front thermocouple assembly and one in the rear thermocouple assembly indicate the overheat limit the EEC will not signal an overheat.

Fault Detection The EEC monitors the thermocouple circuits for faults. Any faults are transmitted to the Aircraft.

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NACELLE TEMPERATURE MONITORING Introduction There are two air temperature sensors used to monitor the temperature of the air. One sensor is in engine zone 1 and one in engine zone 3. Zone 1 The engine Zone 1 fan compartment air temperature is continuously monitored by the EEC, to make sure that it stays in limits. If the air temperature (T Zone 1) becomes higher than the specified limit, the nacelle anti ice and starter duct valves are closed and the EEC transmits an ADVISORY indication to the flight crew.


Zone 3 The engine Zone 3 air temperature is continuously monitored by the EEC, to make sure that it stays in limits. If the air temperature (T Zone 3) becomes higher than the specified limit, the EEC transmits an ADVISORY indication to the flight crew.

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FAN ZONE TEMPERATURE SENSOR Function The primary function of the Fan compartment Temperature Sensor is to detect a burst duct event. If the T zone 1 input signal to the EEC becomes higher than a temperature of 160 deg C (360 deg F), it transmits an ADVISORY indication to the flight crew. The Nacelle anti−ice and starter duct valves supplying hot air are closed by the A/C systems. Location The fan compartment temperature sensor is installed in engine zone 1 and is attached to the engine oil breather pipe located in the lower region of the engine zone. The sensor has a stainless steel housing which contains two temperature−sensing elements.


Description The fan compartment temperature sensor has two 100−ohm platinum resistance temperature detectors (RTDs). The RTD elements are connected to each EEC channel via simplex 2−wire cables, providing separate sensor outputs for each channel. As the temperature in zone 1 increases, the resistance of the sensors will change sending a signal via the EEC. The T zone 1 signal will be made available to the Engine Monitoring Unit (EMU) through the EEC for engine health monitoring.

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ZONE 3 TEMPERATURE THERMOCOUPLE Purpose The Zone 3 Temperature Thermocouple (T Zone 3) is used to sense an increasing air temperature resulting from leaking hot air ducts in the zone 3 area. Location The thermocouple is located on the left side of the engine and secured via a bracket to the Turbine Case Cooling valve assembly.


Description The unit is a dual element insulated junction type thermocouple. The output from each element is connected together to provide a single average output to Channel A of the EEC. Input to EEC channel B is provided by cross wiring from channel A. The temperature measurement is used to generate an indication in the cockpit on the lower ECAM screen Flight deck notification (no crew action), is given when the temperature limit is exceeded to prompt maintenance action to determine the cause of the temperature increase. Should the signal fail a range check then the nacelle temperature indication in the cockpit turns amber and is replaced by XX.

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ATA 79 ENIGINE OIL SYSTEM ENGINE OIL SYSTEM ARCHITECTURE System Architecture The engine oil system serves to lubricate and cool the engine internal drives, gears and bearings. The system is composed of: S an oil tank used for the storage of the oil, S an oil pump unit, which supplies pressure to move the oil to or from the drives, gears and bearings, S a fuel oil heat exchanger, which decreases the oil temperature and increases the fuel temperature, S a scavenge filter, which avoids unwanted particles in the re−circulated oil to enter into the oil tank. The oil filter has a differential pressure transducer, which compares the difference between upstream and downstream pressures to determine if the filter is clogged. In this case, the difference will increase, and transducer will send a signal to the cockpit FADEC (Full Authority Digital Engine Control). After data possessing, the FADEC will send a signal to the FWS (Flight Warning System) for cockpit indications. Each oil tank has an electric magnetic chip detector to attract magnetic debris in the oil. This metallic oil contamination is shown on the ECAM and on the OMS (Onboard Maintenance System) devices for maintenance.



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OIL SYSTEM OVERVIEW Description The engine oil system is a full flow recirculatory system. It must give adequate lubrication and cooling for all engine bearings, gears and driving splines during all operating conditions. The complete system is divided into three primary areas: S The Feed oil and the cooling S The Return oil S The Vent, De−aeration and the Breather System A self−contained oil tank is installed on the right side of the fan case. It incorporates a quantity sight glass and provision is made for gravity oil filling. The system is vented through a centrifugal breather, installed on the rear face of the external gearbox. Oil Cooling The cooling of the feed oil is achieved by a Fuel/Oil Heat Exchanger (FOHE), which controls the oil temperature in the limits.

Indications S The following indications are provided in the cockpit: S Oil tank quantity S Oil temperature S Oil pressure S Oil filter clog


Oil Filtration & Inspection A pressure filter, scavenge filter and line filters (last chance) provide the necessary filtration. Location for magnetic chip detectors (MCDs) are provided in the scavenge lines.

Pump Assembly The pump assembly consists of a pressure pump element to move the oil around the system, and nine scavenge pumps elements, as follows: S LP Turbine Bearing Chamber Scavenge Element S IP Turbine Bearing Chamber Scavenge Element S HP Turbine Bearing Chamber Scavenge Element S Internal Gearbox Front Scavenge Element S Internal Gearbox Rear Scavenge Element S Front Bearing Chamber Scavenge Element S The intermediate gearbox assembly (step aside gearbox) and gearbox input drive (lower bevel box). S External Gearbox Scavenge Element S Centrifugal Breather Scavenge Element

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FEED OIL, LUBRICATION & COOLING Feed oil is circulated by a single pressure pump, which draws oil from the oil tank through a gauze strainer. Additionally the pump has a pressure relief valve, which acts as a by−pass for cold starting and system blockage protection. A 125−micron filter cleans feed oil. A differential pressure transducer monitors filter condition and provides a cockpit indication that the filter is becoming clogged, this switch is set to operate at a differential pressure of 23 psid. The FOHE will keep the oil temperature within limits. The FOHE has two functions: S To decrease the temperature of the oil S To increase the temperature of the fuel An oil pressure relief valve protects the cooler core when the engine oil is very cold or if the core is blocked. An anti−syphon tube prevents oil suction from the FOHE during engine shut down. From the FOHE the feed oil is supplied through external tubes to the main engine bearings, gears and drives.

De−aeration, Breather and Vent System Labyrinth oil seals and the sealing airflows from the engine compressors, prevents oil loss from the bearing chambers. The oversized scavenge pumps and the vent pipes remove the sealing air, which flows continuously through the seals and into the bearing chambers. The return flow is an oil/air mixture. All scavenge oil is de−aerated when it enters the oil tank by a cyclone type separator. The air, which still contains a small amount of oil, is transferred to the inlet of the centrifugal breather. The centrifugal breather separates the air and oil before discarding the air to atmosphere, the oil is scavenged from the breather housing back into the combined scavenge line back to the oil tank.

Return Oil (Scavenge) The return oil/air is scavenged by nine pump elements in the pump module from each of the eight primary lubricated locations of the engine and the breather (air/oil separator). There are positions for installing nine (9) magnetic chip detectors (MCDs), to sample return oil from the engine main bearings and the gearboxes. The oil outlets from the scavenge pumps join to form a combined scavenge return flow which is sampled by the electric master chip detector before passing through a 15−micron fine scavenge filter. The filter has a by−pass valve (20 psid) and a pressure differential transducer (13 psid) to give cockpit indication of impending by−pass. Temperature sensors in the return line between the scavenge pumps and scavenge filter provide cockpit indication of oil temperature.

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OIL TANK Purpose To store the engine oil. Location The oil tank is attached to the A3 & A4 flanges of the LP compressor case on the right side. Capacity Tank contents at the full mark: − 28 US Quarts


Features The tank is a magnesium casting to which other components attach to make up the oil tank assembly. These components are as follows: S Oil quantity transmitter S Sight glass S Oil filler assembly S Scavenge filter assembly S Scavenge Filter Differential Pressure switch S Outlet tube S Vent tube S Electric Magnetic Chip Detector (EMCD) S Oil Temperature Sensors (2)

Description The oil tank provides the reservoir for the engine oil system. The feed line from the oil tank supplies the pressure pump, which feeds the oil system. There is also a coarse filter in the tank to prevent contamination of the oil feed system. The scavenge pumps returns the oil from the various bearing chambers and gearboxes back to the oil tank, along with large quantities of air. A de−aerator in the tank separates the oil and the air. The released air passes from the air space at the top of the tank, via a vent tube to the centrifugal breather mounted on the gearbox. An anti−syphon line carries a small flow of oil from the main feed line back to the oil tank, which is used to clean and cool the sight glass. The oil filler assembly has a quick release cap. Internally the filler has a flap valve, which closes under engine running pressure maintaining sealing if the filler cap is not fitted.

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ENGINE OIL SERVICING Servicing The engine oil servicing is done by the filling of the engine oil system with approved oil and by inspecting the system in order to find and correct the engine oil contamination.

HUMAN FACTOR POINTS:

Inspection of scavenge oil filter To examine the scavenge filter element, it must be removed first. The presence of particles has to be detected and these particles have to be removed. The filter and the filter housing have to be cleaned if magnetic particles are found in the scavenge filter, the electric magnetic chip detector also has to be examined. After examination of the particles, the filter and its seal have to be replaced by new ones. The fuel and oil leak check on the scavenge filter housing is required before putting the aircraft back into operation.

WARNING:

YOU MUST WAIT FOR A MINIMUM OF 10 MINUTES AFTER THE ENGINE HAS STOPPED BEFORE YOU DO A CHECK OF THE OIL LEVEL. THIS WILL LET THE OIL LEVEL BECOME STABLE AND THIS WILL PREVENT OIL SPLASH DUE TO RESIDUAL PRESSURE.

CAUTION:

AVOID SPILLAGE WHEN SERVICING OIL.


Inspection of Electric Magnetic Chip Detector To examine the electric magnetic chip detector, it must be removed first. The presence of particles has to be detected and these particles have to be removed. The detector and the particles are examined. After examination of the particles, the detector can be re−installed but the two seal rings have to be replaced by new ones. The oil leak check on the detector housing is required before putting the aircraft back into operation. Refill of engine oil tank Before replenishing the oil tank, a visual check of the engine oil level in the sight glass of the oil tank must be done. If the engine has been stopped for more than six hours; the engine has to be operated at IDLE before refilling the oil tank by respecting the duration between engine shutdown and the oil tank refilling. To refill the tank, the oil filler cap is removed and the engine lubricating oil also known as material No. OMat 1011 is added. The check for fuel fumes in the tank is required before putting the aircraft back into operation.

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WARNING:

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BE CAREFUL, THE ENGINE PARTS (BLEED DUCT, OIL TANK) CAN STAY HOT FOR ALMOST 1 HOUR AFTER SHUTDOWN.



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OIL QUANTITY TRANSMITTER Purpose The oil quantity transmitter measures the quantity of engine oil in the oil tank to provide a cockpit indication. Location Installed into the top of the oil tank and secured by bolts in the mounting boss. Access by opening the right fan cowl.


Description The oil quantity transmitter is a potentiometer style device with changes in resistance indicating different oil levels. The transmitter consists of a series of reed switches and resistors that form a ladder activated by a float containing a permanent magnet. As the float moves along the stack different reed switches are activated thereby changing resistance. The EEC provides a constant current to the transmitter and as the resistance changes, this results in a change in the output voltage across the resistance stack. The output voltage is measured by channel B of the EEC, which conditions the signal and transmits the oil quantity level to the Electronic Instrument System (EIS) for cockpit display on the lower ECAM (System Display) screen. (Also displayed on ECAM CRUISE page). The needle and the digital indication are normally green. If the oil quantity drops below 4 quarts the digital indication pulses.

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OIL PUMP AND PRESSURE FILTER ASSEMBLY Purpose The oil pump and pressure filter housing supplies the pressurised oil to lubricate the engine bearings and gears. The pump assembly also scavenges oil back to the oil tank. The filter housing contains the pressure oil filter that cleans the feed oil. Location The oil pump and pressure filter assembly is installed on the rear face of the external gearbox between the centrifugal breather and the lower bevel gearbox.


Oil Pump Description The vane type oil pump assembly consists of a pressure pump and nine scavenge pumps to scavenge oil from the various areas of the engine back to the oil tank. The pressure pump has a pressure relief valve, this protects the system for cold starting and blockage protection and is set to 600 psi. The relief valve opens and oil is fed back to the pump inlet, which reduces the system pressure. Pressure Filter Description The pressure filter is installed inside the filter housing, access can be gained by removal of the filter cover. The filter is a 125 micron cleanable type filter and has a life of 3 cleans. The filter is a non−bypass type. A check valve in the housing prevents the loss of oil when the filter is changed. The housing also contains an anti−leak valve to prevent oil draining back to the pump when the engine is shut down.

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MAGNETIC CHIP DETECTORS (MCDS) Purpose A Vickers Electric Master Chip Detector (EMCD) and nine Muirhead Vatric screw−in magnetic chip detectors are installed in the return oil system to allow monitoring of the following: S Front Bearing Housing S Internal Gearbox (Front) S Internal Gearbox (Rear) S HP Turbine Bearing Chamber S IP Turbine Bearing Chamber S LP Turbine Bearing Chamber S Intermediate and Lower Bevel Gearboxes S External Gearbox S Centrifugal Breather

Screw−in MCD The screw−in MCD assembly consists of a housing and Magnetic Chip Detector, which has a magnetic end. When the MCD is installed the magnetic end is located in the return (scavenge) oilways. The MCD housing contains a self−closing check valve to prevent oil leakage when the MCD is removed for inspection. If metallic particles are found on the EMCD during inspection, MCDs can be installed in the ports on the oil pump assembly. This allows the problem to be isolated by checking each scavenge oil line.

Location The EMCD is installed in the combined scavenge return line on the forward side of the oil tank. There are nine ports on the bottom of the oil pump which can be used to install additional MCDs..


NOTE:

During normal engine operation, only the Master EMCD is installed. If metallic particles are found on the EMCD during inspection, diagnostic MCD‘s can then be installed to isolate the source of the debris.

Electric MCD The Electric MCD is positioned at the inlet to the scavenge filter and collects ferrous metal particles from the engine oil. The head of the EMCD has two electrically isolated magnetic poles. A circuit is made when debris bridges the two poles. The EMU continuously monitors the EMCD during flight and generates a EMCD debris maintenance message 10 secs after landing which is sent to the Aircraft via the EEC.

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SCAVENGE FILTER ASSEMBLY Purpose To remove contamination from the scavenge oil that is returning to the oil tank. Location Installed on the rear of the oil tank on the right side of the fan case. Description The assembly consists of the following items: S Scavenge filter assembly (15 micron) S Scavenge filter differential pressure switch S Scavenge filter by−pass valve Scavenge Filter The scavenge filter element cleans the combined scavenge oil returning to the oil tank. The filter element is a non−cleanable, throw away type filter. If the filter element becomes clogged, the by−pass valve opens and allows the oil to flow directly back to the tank. Scavenge Filter Differential Pressure Switch The scavenge filter differential pressure switch monitors the pressure at the inlet & outlet of the filter and provides an indication when the filter becomes partially clogged. The switch is set at 13 psid.


NOTE:

The EEC will inhibit the filter clog message when oil temperature is low, to prevent nuisance messages.

Scavenge Filter By−pass Valve The scavenge filter is fitted with a by−pass valve which operates independently of the differential pressure switch. The by−pass valve operates at 20 psid to maintain oil flow in the event of scavenge filter blockage.

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CENTRIFUGAL BREATHER Purpose To remove the oil from the vent air, before discarding the air overboard. Location Installed in the external gearbox, and located on the rear face between the oil pump and the fuel pump.


Description The centrifugal breather has a rotor that contains retimet segments and is driven by the external gearbox. Aerated oil from the bearing chamber vent system and the oil tank is delivered to the centrifugal breather. The aerated oil tries to pass through the retimet segments but is centrifuged out. The air can pass through the retimet segments into the hollow rotor and is vented overboard. The centrifuged oil is scavenged back to the oil tank by the breather scavenge pump element.

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FUEL/OIL HEAT EXCHANGER (FOHE) Purpose The FOHE has two functions as follows: S To reduce the temperature of the engine oil S To prevent the water content of the fuel from turning to ice. Location The FOHE is mounted horizontally below the oil tank on the right side of the fan case. Description The FOHE consists of the following units: S FOHE matrix core S By−pass Valve S Oil pressure transmitters (2) S Low oil pressure switch S LP Fuel Filter S Fuel Filter Differential Pressure Switch


FOHE Matrix Core Heat is transferred from the oil to the fuel in the core of the FOHE. The oil flow is made slower by many baffle plates around the steel tubes through which the fuel is flowing. The slower oil flow enhances the exchange of heat. An anti−syphon hole connects the inlet to outlet to prevent oil suction from the FOHE during engine shut down. By−pass Valve If the oil pressure in the FOHE becomes more than a specified limit a by−pass valve will open allowing the oil to by−pass approximately two−thirds of the core. This normally occurs when the engine oil is extremely cold. Oil Pressure Transmitters & Low Oil Pressure Switch Description on page 6−19 LP Fuel Filter, Differential Pressure Switch & By−pass Valve Description in Fuel System

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OIL PRESSURE INDICATION

LOW OIL PRESSURE SWITCH

Purpose The oil pressure transducers (2) assembly measures the differential oil pressure between the output of the oil pump and the internal gearbox chamber scavenge line.

Purpose The low oil pressure switch measures the same differential oil pressure as the oil pressure transducers to provide an indication when the oil pressure drops to a pre−set level.

Location The oil pressure transducers assembly is installed on the FOHE assembly. Access by opening the right fan cowl door.

Location The switch is installed on the FOHE.

Description The oil pressure transducer assembly consists of two transducers supplied by the same oil pressures. They are ratiometric devices supplied with a constant reference voltage, as the pressure changes the internal resistance of the transducers change, resulting in variable voltage outputs, which are read by the EEC. One pressure transducer provides a signal to EEC channel A, the other transducer provides a signal to channel B. The EEC conditions and filters the signal to remove pump ripple effects and sends the smooth oil pressure signal to the EIS for cockpit display on the lower ECAM screen. The needle and the digital indication are normally green. The needle and the digital indication will turn red if the oil pressure drops below 25 psi.

Low Oil Pressure Indication Oil pressure warnings are provided for two conditions: S Oil pressure less than an N3 related value (Amber) S Oil pressure less than the minimum threshold (Red) The amber minimum oil pressure is proportional to N3 and is output to the cockpit display to help format the oil pressure display. When the oil pressure falls below the amber value but remains above the low pressure switch limit, the cockpit display will become amber and a maintenance message will be set. When the oil pressure falls below the threshold value a Red Alert is set. This is derived from the 3 inputs, the oil pressure transducers and low oil pressure switch. With the engine running 2 out of the 3 inputs are required to enunciate the low oil pressure condition.


Oil Pressure Limits S Minimum oil pressure 25 psi.

Description When the differential oil pressure falls below the switch pre−set value, the internal contacts close indicating a loss of oil pressure. The output of the switch is fed directly to the aircraft system (AFDX) and from there to the EEC.

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OIL PRESSURE FILTER DELTA P TRANSDUCER Purpose The oil pressure filter differential pressure (dP) transducer monitors the condition of the filter element in the oil pressure filter. It provides a cockpit indication of impending filter blockage. Location The pressure filter dP transducer is located on the oil pump & pressure filter assembly. Access to the unit is via the right fan cowl. Description The oil pressure filter differential pressure transducer measure the pressure drop across the filter element. The transducer is a ratiometric device supplied with a constant reference voltage. As the pressure changes, the internal resistance of the device changes resulting in a variable voltage output that is read by the EEC. As the filter becomes clogged with debris, the pressure drop increases and when it reaches a pre−determined level, the filter is considered to be blocked. The pressure filter dP transducer sends the output signal to channel B of the EEC. Note: When the oil temperature is too low, the EEC will inhibit the filter block indications, to prevent nuisance indications.


Indications Impending blockage of either the pressure filter or scavenge filter will give the following indications: Engine Warning Display (EWD) S OIL FILTER CLOG message S Aural Warning S Master Caution System Display (SD) S Amber CLOG message

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OIL SCAVENGE FILTER DELTA P TRANSDUCER Purpose The oil scavenge filter differential pressure (dP) transducer monitors the condition of the filter element in the oil scavenge filter. It provides a cockpit indication of impending filter blockage. Location The scavenge filter dP transducer is located on the oil tank. Access to the units is via the right fan cowl. Description The Scavenge oil filter differential pressure transducer measures the pressure drop across the filter element. The transducer is a ratiometric device supplied with a constant reference voltage. As the pressure changes, the internal resistance of the device changes resulting in a variable voltage output that is read by the EEC. As the filter becomes clogged with debris, the pressure drop increases and when it reaches a pre−determined level, the filter is deemed to be blocked. The scavenge filter dP transducer send the output signal to channel B of the EEC. NOTE:

When the oil temperature is too low, the EEC will inhibit the filter block indications, to prevent nuisance indications.


Indications Impending blockage of either the pressure filter or scavenge filter will give the following indications: Engine/Warning Display (E/WD) S OIL FILTER CLOG message S Aural Warning S Master Caution System Display (SD) S Amber CLOG message

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OIL TEMPERATURE SENSOR Purpose The oil temperature sensors within the engine oil system are used to sense the scavenge oil. Location The two temperature sensors are installed on the oil tank in the scavenge return line.. Description The sensors are resistance temperature devices (RTD‘s) which have a variable resistance with temperature. Each temperature sensor sends a signal to one channel of the EEC. The EEC processes the output of the sensors and provides an output for cockpit display. Oil temperature is also used for starting and accel fuel scheduling. The EEC continuously monitors the outputs of the RTD‘s. In the event of a failure of one of the sensors, the EEC shall select the remaining valid sensor. When there is a disagreement between the two measurements, the higher of the two values will be selected.


Indications The oil temperature displays in degrees C on the ECAM SD screen and is normally green. The indication turns Amber if the temperature exceeds TBD °C . Engine/Warning Display (E/WD) S OIL LO TEMP S OIL HI TEMP S Aural Warning S Master Caution System Display (SD) S Engine STATUS page

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OIL SYSTEM SERVICING (AMM 79−00−00−610−801) WARNING:

YOU MUST BE CAREFUL WHEN YOU DO WORK ON THE ENGINE PARTS AFTER THE ENGINE IS SHUT DOWN. THE ENGINE PARTS CAN STAY HOT FOR ALMOST 1 HOUR.

WARNING:

YOU MUST NOT TOUCH HOT PARTS WITHOUT APPLICABLE GLOVES. IF YOU GET AN INJURY PUT IT IN COLD WATER FOR 10 MINUTES AND GET MEDICAL AID.

WARNING:

YOU MUST NOT LET ENGINE OIL STAY ON YOUR SKIN. FLUSH THE OIL FROM YOUR SKIN WITH WATER. YOU MUST NOT BREATHE THE FUMES. YOU MUST NOT GET OIL IN YOUR EYES OR MOUTH. PUT ON GOGGLES OR A FACE MASK. IF YOU GET OIL IN YOUR MOUTH, YOU MUST NOT CAUSE VOMITING BUT GET MEDICAL AID IMMEDIATELY.

Procedure: The procedure in the AMM is briefly described below: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers S Open the oil tank access panel in the right fan cowl door S Do a visual check of the oil level in the oil tank sight glass NOTE:

You must wait at least 10 minutes after engine shutdown for the oil level to become stable.

If the engine rpm was not stabilised at idle before engine shutdown, the oil system will not have become stable. In this condition the oil quantity indication can apparently be low, this is normal and the engine oil system must not be filled. S If the engine has been stopped for less than 6 hours and the oil level is low, fill the engine oil tank.


NOTE:

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S If the engine has been stopped for more than 6 hours and the oil level is low, but not below 4.73 litres (5 US quarts) from the required level, then: − Do not fill the engine oil system − Start the engine and operate at idle for 5 minutes − Stop the engine − Do a check of the engine oil level again (after waiting at least 10 minutes for the oil level to become stable) − If the oil level is low, fill the oil tank S If the engine has been stopped for more than 6 hours and the oil level is below 4.73 litres (5 US quarts) from the required level, then: − Drain the external gearbox (AMM 79−00−00−680−801) − Fill the engine oil tank − Start the engine and operate at idle for 5 minutes − Do a check of the engine oil level again (after waiting at least 10 minutes for the oil level to become stable) − If the oil level is low, fill the oil tank Notes: 1. Add clean approved engine oil (Omat 1011) to the tank. 2. Check for fuel fumes when you remove the oil tank cap (fuel fumes are easier to find when the oil is hot) AMM task 79−00−00−280−801 3. Check condition of seal ring in the groove of the oil filler cap before installing. Replace if loose or damaged.

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OIL SCAVENGE FILTER REMOVAL/INSTALLATION (AMM 79−22−45) ATTENTION: Warnings and Cautions: Observe all Warnings and Cautions given in the AMM


Removal Procedure: The procedure in the AMM is briefly described below: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers S Open the right fan cowl door S Put a clean 10 L container into position to catch the oil S Remove the drain plug from the filter housing and drain the oil into the container (Do not discard the oil at this step) S Remove seal from the drain plug, install a new seal and refit the drain plug in the housing and torque S Hold the housing and remove the bolts and washers S Carefully remove the housing and filter from the scavenge filter cover S Examine the element and the drained oil for contamination (AMM Task 79−00−00−280−801) S Discard the element S Remove and discard the seal ring

Installation Procedure S Examine the inner area of the housing and make sure it is clean and clear of unwanted material S Install a new seal ring to the housing S Carefully install the filter element in the scavenge filter cover. Make sure you hold the filter element in this position. S Put the housing in position on the scavenge filter cover S Attach the housing with the bolts and washers S Torque the bolts to the value given in the AMM S Fill the engine oil system S Do a fuel and oil leak check of the scavenge filter housing S Put the aircraft back to its initial configuration

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Figure 136 FRA US/T

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OIL PRESSURE FILTER REMOVAL/INSTALLATION (AMM 79−22−43) ATTENTION: Warnings and Cautions: Observe all Warnings and Cautions given in the AMM Removal Procedure: The procedure in the AMM is briefly described below: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers S Open the right fan cowl door S Put a clean container into position to catch the oil S Remove the lockwire or safety cable that safeties the housing S Turn the housing counter clockwise to release it (Use a strap wrench if necessary) S Carefully remove the housing and filter element from the oil pump assembly

Installation Procedure S Make sure there is a new seal ring installed on the new element S Install a new seal ring on the housing S Carefully install the filter element in the oil pump assembly. Make sure you hold the filter element in this position. S Put the housing in position on the oil pump assembly S Turn the housing in a clockwise direction with your hand until it is tight You must only tighten the filter housing with your hand only. If you use tools you can cause damage to the screw threads. Safety the housing with lockwire or Safety Cable Fill the engine oil system Do a fuel and oil leak check of the pressure filter housing Put the aircraft back to its initial configuration

NOTE: S S S S

MAKE SURE YOU REMOVE THE ELEMENT WITH THE HOUSING. IF THE FILTER ELEMENT IS NOT REMOVED AT THE SAME TIME IN CAN FALL AND DAMAGE THE PART S Drain the oil from the housing and element into a clean container S Remove and discard the seal ring S Examine the element and the drained oil for contamination (AMM Task 79−00−00−200−804)


CAUTION:

NOTE: The element can be cleaned and used again S Put the element into a clean container for its protection

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EMCD INSPECTION (AMM TASK 79−00−00−200−802) ATTENTION: Warnings and Cautions: Observe all Warnings and Cautions given in the AMM Procedure: The procedure in the AMM is briefly described below: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers S Open the right fan cowl door S Remove the master EMCD S Put the master EMCD probe into clean kerosene and remove the oil (The kerosene should be in a clean non−metallic container) Make sure you do not contaminate the electrical contacts with kerosene Be careful to only remove the oil from the EMCD and not any contamination, which may be present. Examine the master EMCD in good light for contamination using a 05X magnifying glass Refer to the contamination standards in the AMM. (The types of contamination are shown on the next page) Keep all contamination which does not cause you to immediately reject the engine or gearbox as a record. Take a piece of 25mm (1 inch) wide transparent self adhesive tape preferably Scotch Magic Tape (OMat 1269) approximately 50mm long and apply the centre of the gummed side over the recessed insulated debris gap. It may require several attempts to remove all the debris with the same piece of tape

NOTE: S S S


S

NOTE:

The contamination record will help you monitor the type of wear in the engine or gearbox

NOTE:

Laboratory analysis is recommended to help with material identification

NOTE:

If the contamination is outside the permitted standards you must refer the contamination to Rolls−Royce for recommended action.

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EMCD INSPECTION (AMM TASK 79−00−00−200−802) ATTENTION: Warnings and Cautions: Observe all Warnings and Cautions given in the AMM Procedure: The procedure in the AMM is briefly described below: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers S Open the right fan cowl door S Remove the master EMCD S Put the master EMCD probe into clean kerosene and remove the oil (The kerosene should be in a clean non−metallic container) Make sure you do not contaminate the electrical contacts with kerosene Be careful to only remove the oil from the EMCD and not any contamination, which may be present. Examine the master EMCD in good light for contamination using a 05X magnifying glass Refer to the contamination standards in the AMM. (The types of contamination are shown on the next page) Keep all contamination which does not cause you to immediately reject the engine or gearbox as a record. Take a piece of 25mm (1 inch) wide transparent self adhesive tape preferably Scotch Magic Tape (OMat 1269) approximately 50mm long and apply the centre of the gummed side over the recessed insulated debris gap. It may require several attempts to remove all the debris with the same piece of tape

NOTE: S S S


S

NOTE:

The contamination record will help you monitor the type of wear in the engine or gearbox

NOTE:

Laboratory analysis is recommended to help with material identification

NOTE:

If the contamination is outside the permitted standards you must refer the contamination to Rolls−Royce for recommended action.

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EMCD AFTER REMOVAL

MCD WASHING

Figure 139 FRA US/T

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TRANFERING DEBRIS ONTO RECORDING CARD

DEBRIS TRANSFER TO SCOTCH MAGNETIC TAPE

Figure 140 FRA US/T

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Debris Transfer

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MCD PICTURE


MCD wet

HARDWARE FAILURE

MCD dry

EMCD wet

EMCD dry

Figure 141 FRA US/T

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Fines appear on an oily MCD as a black sludge. After being degreased they can, with naked eye, be mistaken for very small metallic flakes.

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A380 79 MCD PICTURE

HARDWARE FAILURE

The gear scuffing shown produces relatively coarse fines. Note: normal „wear“ fines are similar in size to those produced by bearing lapping failures.

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MCD PICTURE

HARDWARE FAILURE

RACE

MCD


BALL

These can be sub-divided into ball bearing , roller bearing, bearing track and gear teeth flakes. • Ball bearing and ball bearing track flakes are usually roughly circular with radial splits • Roller bearing and roller bearing track flakes can be roughly rectangular in shape with criss-cross scratches, but are usually similar to ball bearing flakes • Fatique flakes are typically 0,5 - 1,0 mm (0.020 - 0.040 inch) in diameter, and very thin.

EMCD

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Gear teeth fragments - corner pieces aof gear teeth may be evidence of incorrect geatr alignement or bedding, or handling damage during overhaull.

Figure 144 FRA US/T

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MCD PICTURE

HARDWARE FAILURE


Chip, ground surface

Chip, rough surface

Chips - these are very thick flakes or definite lumps of metal usullay with one ground (smooth) surface. Bearing race spalling can produce chips in addition to flakes.

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Note: there may be criteria for the number of rivets found. Refer to Aircraft Maintenance Manual.

Figure 146 FRA US/T

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If cage tanges are found, refer to the Aircraft Maintenance Manual or a local Rolls-Royce Service representative.

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Explanation While every effort is made to remove manufacturing or build debris (swarf), unfortunately small amounts may be present within the engine on build. This beris will be washed down by oil system to the MCD‘s. Pieces of tuning are easily identifiable but milling debris, ehrn broken up, could possibly be confused with gear or steel rubbings and must be carfully examined.

Hairs of seal lining meterial or seal fin material


In assition there may be some running-in or bedding-in of the engine which may produce a small amount of additional debris. Both will reduce after a short period of time. Seal lining material is sometimes released from the bearing chamber oil seal into the system after engine surges.

Chunks of lining or seal fin material

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(E)MCD debris discovered

A380 79

Compare analysis against Aircraft Maintenance Manual acceptance criteria.

Analyse debris using visual and (optional) SEM processes

Yes

Acceptable? No Fit diagnostic MCD‘s in all positions. Carry out a ground run or two, 90 second dry motor cycles

Remove engine from service immediately for investigation

No

Inspect all MCD‘s fitted, scavenge screens and oil filters. Compare against Aircraft Maintenance Manual acceptance citeria. Acceptable? Yes


Indicate an alertr status an accumulated records and request MCD inspections to be taken at more frequent intervals

No

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Does debris rate reduce?

Yes

Resume normal monitoring procedure

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PRESERVATION OF MAIN LINE BEARINGS Reason for the Job Used oil can cause corrosion if you keep an engine in storage for a long period. The preservation procedure makes it necessary to drain all the oil from the oil system. You must then supply new engine oil to the bearings. It is necessary to motor the engine to move the oil through the system. This can only be done with the engine installed or on a test bed. Procedure The procedure is fully described in AMM Task 71−00−00−600−805 but can be briefly described as follows: S Gain access to the engine S Drain the oil tank S Drain the oil from the fuel/oil heat exchanger S Drain the external gearbox S Drain the oil pressure filter housing S Drain the oil scavenge filter housing S Drain the oil pump assembly S Fill the engine oil tank with clean approved oil S Dry motor the engine until you see an oil pressure indication on the ECAM display screen. S If there is not an oil pressure indication after 30 secs after N3 starts to turn, stop the dry motor and carry out the following step. − Make sure you see oil in the oil tank sight glass − If necessary put oil in the tank until you can see oil in the sight glass − Repeat the dry motor of the engine until you see an oil pressure indication on the ECAM display screen. − Put the aircraft back to its initial configuration.

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POWER PLANT OIL

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A380 80

ATA 80 STARTING ENGINE STARTING SYSTEM Introduction The EEC is able to perform Automatic and Manual engine starts initiated by digital command signals from the aircraft AFDX system. To achieve engine starting, the following sub−systems are combined: S Starting S Fuel S Ignition Each channel of the EEC interfaces with the Start Control Valve (SCV), the Heigh Energy Igniter Unit (HEIU) / systems and the minimum pressure and shut−off valve (MPSOV), in order to control their operation during the starting/cranking phases. The EEC controls the engine starting sequences, engine cranking options and the ignition selection in response to aircraft command signals.



POWER PLANT STARTING

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A380 80


e. g. Engine #2

Figure 151 FRA US/T

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A380 80

ENGINE STARTING COMMAND CONTROLS Cockpit Starting/Shutdown/Ignition Controls Engine starting, cranking and ignition selection are commanded from the aircraft cockpit panels. Engine cranking and starting is initiated by the EEC based on the position of the switches on the aircraft cockpit panels. These switch positions are transmitted to the EEC via the aircraft Avionics Full Duplex Switched Ethernet (AFDX). Engine Control Panel The engine control panel is located on the central pedestal in the cockpit and comprises: 1. Four Master Levers, one per engine each with two positions: A. Engine ON B. Engine OFF 2. One Rotary Selector (which serves all four engines) with three positions: A. Crank B. Normal C. Ignition/Start


Engine Manual Start Push−button The engine manual start push−buttons are located on the overhead panel of the cockpit. There is one guarded push−button for each engine. Auto & Manual Starting For automatic engine starts, only the Master Lever (one per engine) and Rotary Switch are used. For manual (alternate) engine starts, the Manual Start push button is used as well as the Master Lever and Rotary Switch. Hardwired Master Lever position discrete signals are also available to the EEC. If the aircraft is in flight and the Rotary Switch position is invalid the EEC will initiate an auto relight if the Master Lever is toggled from OFF to ON.

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Figure 152 FRA US/T

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Cockpit Panels

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COCKPIT INDICATION During the start sequence, the nacelle temperature indications are replaced by the ignition and starting parameters on the System Display (SD). The start parameters displayed are as follows: S Ignition (A, B or AB). S Start control valve position. S Air pressure to the starter.




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Figure 153 FRA US/T

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STARTER CONTROL VALVE (SCV) Purpose The starter control valve (SCV) controls the air supply to the starter motor. Location The starter control valve is installed in the starter duct at the lower left side of the fan case. Access by opening the left fan cowl door.


Operation On command from the EEC, the pneumatic starter control valve, controls the flow of air to the pneumatic starter. During the starting sequence the pneumatic starter control valve is commanded closed after the starter has reached cut−out speed. A solenoid−operated valve and regulator control the supply of starter air duct pressure to an actuator that moves the butterfly valve. The solenoid contains a double coil assembly that is controlled by the EEC, one coil being connected to EEC channel A, the other to EEC channel B. The regulator limits the pressure of air to the pneumatic starter. Two microswitches give an indication to the EEC of the valve position. One microswitch is connected to EEC channel A, the other to EEC channel B. Manual Operation An extension of the butterfly valve shaft has a visual position indicator and a square socket to permit manual operation of the butterfly valve. Access to the square drive is via a sprung loaded flap in the fan cowl door. Manual operation of the pneumatic start control valve does not require the fan cowl door to be opened. This method can be used to dispatch the aircraft with a fault in the SCV system. Failure of SCV Failure of the pneumatic starter control valve to close, or leakage of the starter air ducting, is determined by commanded valve position and a flow detection system in the start air ducting. When a failure is detected the engine cabin bleed and aircraft cross−flow valves are closed to prevent an overspeed of the pneumatic starter or leakage of air into Zone 1.

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Figure 154 FRA US/T

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STARTER MOTOR Purpose The pneumatic starter motor turns the external gearbox for starting and motoring the engine. The external gearbox turns the high−pressure shaft (N3). Location The starter motor is located on the left side front face of the external gearbox. Access by opening the left fan cowl.


Description The pneumatic starter motor consists of: S Inlet housing with containment baffles S Turbine rotor assembly S Reduction gears S Gear cage S A clutch S Transmission housing S Splined output shaft The inlet housing is designed to lessen the danger of turbine blades exiting the starter in the event of a turbine rotor assembly failure. A quick attach−detach (QAD) clamp attaches the starter motor to a QAD adapter, which is bolted to the front face of the external gearbox. Dowels between the starter and adapter ensure correct alignment of the starter motor.

Starter Oil System The starter motor has a self−contained oil system with the following parts: S Gravity fill and overflow plug S Oil sight glass S Drain plug with an integral magnetic chip detector (MCD) The MCD catches ferrous contamination in the starter oil system. To inspect the MCD, unscrew from the drain plug housing and remove.

Functional Description The air supply from the starter air duct turns the turbine at high speed with low torque. The reduction gears reduce the speed and increase torque to the clutch mechanism and output shaft. After passing through the turbine the air is released to ambient through the exhaust deflector baffles. The clutch mechanism (SEC) disengages the starter from the engine once the engine reaches idle speed.

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Figure 155 FRA US/T

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STARTER OIL SERVICING (AMM 80−11−41) The starter servicing procedure is detailed in the AMM 80−11−41 and briefly described below:

Detailed Inspection of the Starter MCD (AMM 80−11−41) The inspection of the Starter MCD is detailed the AMM 80−11−41 and briefly described below: Note: The magnetic chip detector is installed through the center of the drain plug. Do not remove the drain plug.

WARNING:

DO NOT LET ENGINE OIL STAY ON YOUR SKIN. POISONOUS MATERIALS CAN BE ABSORBED THROUGH YOUR SKIN.

CAUTION:

REMOVE ANY OIL SPILLAGE ON THE ENGINE IMMEDIATELY WITH A LINT FREE CLOTH AS THE OIL MAY CAUSE DAMAGE TO THE SURFACE PROTECTION OR SOME ENGINE PARTS.

Procedure S Open the fan cowl doors. S Remove the drain plug and allow the oil to drain from the starter S Install a new seal ring on the drain plug and install the drain plug, torque tighten and safety. S Remove the oil fill and overflow plugs from the starter and discard the seals. S Add clean oil to the starter oil fill position until oil starts to drip from the oil level overflow. S Wait until oil does not drip from the oil level overflow. S Remove oil from the external surface of the starter with a clean cloth S Put new seal rings on the plugs and install the oil fill and overflow plugs, torque tighten and safety. S Look at the oil level sightglass and make sure the level is above the ADD mark.

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Procedure: S Remove the MCD plug and allow the oil to drain from the starter S Put the MCD in clean Kerosene and remove the oil. The Kerosene must be in a clean non−metallic container. S Examine the MCD in good light for contamination. If the MCD has chips larger than 0.1 in (2.54 mm) in one direction is found, reject the starter. S Keep any contamination which you find as this will help you to keep a record of type of wear in the starter. S If you reject the starter, send the contamination which you found with the starter to the service bay.

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Figure 156 FRA US/T

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POWER PLANT IGNITION

A380 74

ATA 74 IGNITION ENGINE IGNITION SYSTEM Purpose Each engine has two ignition systems, “1“ and “2“, which can operate together or independently to supply electrical sparks in the combustion chamber to ignite the fuel/air mixture or keep combustion going.


Location The ignition units are located on the lower left side of the fan case, rear unit (system “1“) and front unit (system “2“). Access through left fan cowl. The two igniter plugs fit through the outer combustion case. Description Each engine has two ignition systems, which can operate independently or together. Each system has an ignition power supply and an ignition distribution system. The ignition systems are operated when supplied with 115 V AC (variable frequency) aircraft power by the EEC The engine master lever must be in the “ON“ position for supply of ignition electrical power. EEC logic is also used to operate the ignition systems when inclement weather is detected at low power or there is an un−commanded engine run down: S Automatic inclement weather protection (continuous ignition) S Auto relight function Continuous ignition is only used when commanded by the EEC for automatic inclement weather protection, there is no manual selection of continuous ignition.

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Ignition Selection The EEC alternates between the ’1’ and ’2’ ignition systems for Automatic engine ground starts. In the event of a “failure to light“ being detected using one system, the EEC automatically commands both ignition systems to operate and sends an “ignition failed“ message to the aircraft identifying the inoperative components. The EEC commands both ignition systems to operate for: S manual ground starts S in−flight starts S automatic inclement weather protection (continuous ignition) S auto relight. Quick Relight If the master lever is accidentally moved to the OFF position during engine operation and subsequently returned to the ON position within 30 seconds, the EEC will command both ignition systems to operate until 10 seconds after the engine reaches idle providing: S The HP shaft speed is above 10% when the master lever is returned to the ON position when in flight S The HP shaft speed is above 50% when the master lever is returned to the ON position when on the ground.

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Figure 157 FRA US/T

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IGNITION SYSTEM COMPONENTS Ignition Unit The ignition unit is a capacitive discharge circuit. The unit converts the 115 V AC variable frequency aircraft power supply input to provide an output voltage of approximately 3 kV. Energy is stored in the ignition unit at 7.5 − 12.0 Joules and released by the unit to give a spark to the igniter plug at a minimum rate of 60 sparks per minute. There is no BITE within the ignition unit. The EEC monitors supply of power to the ignition unit to enable faults to be annunciated and to allow automatic selection of the other ignition system. Igniter Plug The igniter plug is a surface discharge type and is used to ignite the fuel / air mixture in the combustion chamber. The igniter plugs are installed adjacent to the fuel spray nozzles at position numbers eight (system 1) and twelve (system 2) when the engine is viewed clockwise from the rear.


Ignition Lead The ignition leads connect the ignition units to the igniter plugs and have replaceable contact buttons at both ends. The ignition lead from the igniter plug located adjacent to the number eight fuel spray nozzle is connected to the rear ignition unit (system 1) and the igniter lead from the igniter plug adjacent to the number twelve fuel spray nozzle is connected to the front ignition unit (system 2).

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Figure 158 FRA US/T

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IGNITION SYSTEM DESCRIPTION Ignition Power Supply The ignition system 1 and ignition system 2 use power from the 115VAC Emergency (Ignition) and 115VAC Normal (Ignition) electrical buses respectively. The EEC switches the supply of electrical power to each ignition box. When in control, each channel can command both electrical power switches, even when the other channel processor system is faulty. The EEC incorporates a power monitor in each electrical power supply output to the ignition box. The ignition boxes, when supplied with 115VAC convert and output this power to the high tension supply to their respective surface discharge Ignition Plugs. The engine electrical ground plane provides the high tension power return patch.


Single & Dual Ignition Selection The ignition system may be required by a number of functions, with a primary choice of single or dual ignition circuit operation according to the table opposite. To optimise the life of the ignition system (especially ignition plugs) and minimise the risk of operation with dormant faults, the EEC will normally command single ignition circuit operation for ground Autostart. The cycle of ignition circuit operations, during ground autostart, is shown on the table opposite. This will exercise all software controlled combinations of EEC channel in control, ignition exciter, lead and plug and power source to expose otherwise dormant faults.

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Figure 159 FRA US/T

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REPLACEMENT OF THE IGNITION LEAD CONTACTS Removal of the Ignition Lead Contacts This task is covered in AMM Task 74−21−52−000−802 YOU MUST ISOLATE THE ELECTRICAL POWER SUPPLY AT LEAST 3 MINUTES BEFORE YOU WORK ON THE IGNITION SYSTEM. THIS WILL LET THE SYSTEM VOLTAGE DECREASE. THE IGNITION SYSTEM USES VERY HIGH VOLTAGES WHICH ARE DANGEROUS. THE ELECTRICAL POWER IS SUFFICIENTLY STRONG TO CAUSE AN INJURY OR KILL YOU. Make sure you observe all the Warnings in the AMM procedure. S Disconnect the electrical input lead connector from the ignition system you are going to work on. S Put a blanking cap on the disconnected connector. WARNING:

YOU MUST NOT BEND THE IGNITION LEADS TOO MUCH WHEN YOU DISCONNECT/CONNECT THEM. THE IGNITION LEADS CAN BE DAMAGED AND CAUSE ELECTRICAL CIRCUIT DEFECTS. Remove and discard the lockwire and disconnect the ignition lead connector from the applicable ignition unit. Remove and discard the lockwire and disconnect the ignition lead connector from the applicable igniter plug. Using pliers, remove the locating ring. Remove the ceramic insulator Remove the contact from the contact body Put blanking caps on the ignition unit, igniter plug and ignition lead.

CAUTION:

S


S S S S S

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Installation of the Ignition Lead Contacts This task is covered in AMM Task 74−21−52−400−802 Make sure you observe all the Warnings in the AMM procedure. S Remove the blanking caps from the ignition lead S Install the contact in the contact body S Install the ceramic insulator on the contact body S Attach the insulator with the locating ring S Connect the ignition lead connector to the applicable ignition unit S Torque the connector and safety with lockwire or safety cable. S Connect the ignition lead connector to the applicable igniter plug S Torque the connector and safety with lockwire or safety cable. S Connect the electrical input lead connector on the ignition unit S Do a test of the Ignition System.

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Figure 160 FRA US/T

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IGNITER INSPECTION (AMM 74−21−51−200−801) YOU MUST ISOLATE THE ELECTRICAL POWER SUPPLY AT LEAST 3 MINUTES BEFORE YOU WORK ON THE IGNITION SYSTEM. THIS WILL LET THE SYSTEM VOLTAGE DECREASE. THE IGNITION SYSTEM USES VERY HIGH VOLTAGES WHICH ARE DANGEROUS. THE ELECTRICAL POWER IS SUFFICIENTLY STRONG TO CAUSE AN INJURY OR KILL YOU. Observe all AMM Warnings Examine the igniter plug for the following damage: S Examine the igniter plug body, the joints and the igniter plug insulation above the contact button for cracks. If cracked, reject. S Frettage in the outer shell of the igniter plug. See AMM for limits. S Examine the igniter tip for erosion. See AMM for limits. S Examine the contact button for damage. If damaged, reject. WARNING:

NOTE:

If excessive contact pitting can be seen due to arcing impingement, reject.

If excessive contact pitting can be seen, examine the ignition lead contact for excessive pitting. If excessive pitting can be seen, replace the ignition lead contact. S If the igniter is corroded and/or pitted, reject.


NOTE:

NOTE: The igniter body has a rough surface when new. S Examine the plug center electrode. If the center electrode is not there, reject. If rejected, make an inspection of the HP turbine blades. S Examine the insulator (between the center and outer electrode) for cracks or missing pieces. If it is cracked or has missing pieces, reject.

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Figure 161 FRA US/T

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IGNITER PLUG REPLACEMENT IGNITER REMOVAL (AMM 74−21−51−000−801) YOU MUST ISOLATE THE ELECTRICAL POWER SUPPLY AT LEAST 3 MINUTES BEFORE YOU WORK ON THE IGNITION SYSTEM. THIS WILL LET THE SYSTEM VOLTAGE DECREASE. THE IGNITION SYSTEM USES VERY HIGH VOLTAGES WHICH ARE DANGEROUS. THE ELECTRICAL POWER IS SUFFICIENTLY STRONG TO CAUSE AN INJURY OR KILL YOU. Observe all AMM Warnings WARNING:

Procedure S Apply anti−seize compound (Omat 4−62) to the threads of the igniter plug S Use HU43915 (socket) to install the igniter plug and torque to the value given in the AMM It is not necessary to carry out an immersion depth check when installing an igniter. The adjusting shims are located between the adapter and outer case and these are not removed during igniter plug replacement. Clean the contact buttons with emery paper (Omat 5−43). Remove the dust with a lint free cloth Connect the ignition lead to the igniter plug, torque and safety Connect the applicable electrical input lead connector to the ignition unit Reset the applicable circuit breakers Carry out a test of the Ignition system

NOTE:

S S S S S


Procedure: S On the OMT, get access to the Power Distribution Control management pages and open & safety the applicable circuit breakers. S If working on an inboard engine, make sure the thrust reverser is deactivated. S Open the fan cowls and fan exhaust cowls S Disconnect the applicable electrical input lead connector on the ignition unit and cap the connector S Remove the lockwire and disconnect the applicable ignition lead from the igniter plug and cap the connector. S Use HU43915 (socket) to remove the igniter plug and install blanking caps on the plug and opening.

IGNITER INSTALLATION (AMM 74−21−51−400−801) Observe all safety Warning & Cautions

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Figure 162 FRA US/T

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A380 78

ATA 78−30 THRUST REVERSER TRENT 900 NACELLE OVERALL PRESENTATION The nacelle is the aerodynamic structure around the engine. The primary functions of the nacelle : S Ensure smooth airflow both around and into the engine S Protect the engine and the engine accessories S Provide engine noise attenuation S Permit access to the engine & its components for servicing and maintenance S Reverse engine fan flow after landing to brake aircraft S Provide ventilation of the engine fan and core zones S Participate to engine load distribution (load sharing)


The TRENT900 nacelle is composed of : S Air Intake Cowl S Fan Cowl Doors S Thrust Reverser Cowl Doors only for inboard engines. S Electronic Thrust Reverser Actuation System, mounted on the Thrust Reverser forward frame. S Fan Exhaust Cowls for outboard engines S Exhaust Nozzle and Plugs (rear and forward)

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Thrust Reverser or Fan Exhaust Cowl

Pylon

Exhaust Nozzle


Exhaust Plugs

Fan Cowl Door

TRENT TRENT 900 900 Engine Engine Figure 163 Trent 9000 Nacelle Overalll Presentation

Air Intake Cowl FRA US/T

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THRUST REVERSER COWL DOORS Thrust Reverser Cowl Doors Presentation The A380 inboard engines are equipped with Thrust Reverser Cowl Doors. (Outboard engines are equipped with Fan Exhaust Cowls. The main function of the Thrust Reverser is to contribute to the aircraft braking at landing. The Thrust Reverser assembly encloses the engine core with an aerodynamic flow path, and provides a fan exhaust duct and nozzle exit. The Thrust Reverser assembly is located between the Fan Cowl Doors and the Exhaust. It is attached to the wing pylon by four hinges. Two hinges are attached to floating rods. The Thrust Reverser assembly is a cascade type Thrust Reverser with translating cowls and blocker doors. It is made of two halves that make a duct around the engine core. Each half consists of a fixed structure, which provides support for the cascades and actuation system, and a translating cowl. The Thrust Reverser halves open at 6 o’clock and rotates around the 12 o’clock hinge beam to give access to the engine during maintenance operations. The Thrust Reverser system is composed of : S the structure S the Powered Cowl Opening System S the control and indicating : the ETRAS (Electronic Thrust Reverser Actuation Controller)

Thrust Reverser Cowl Doors Structure The Thrust Reverser structure is composed of a fixed structure (outer and inner fixed structure) and the translating cowl. FIXED STRUCTURE Outer Fixed Structure Assembly The outer fixed structure assembly is composed of the following components: S the forward frame S the J−ring S the cascades assembly (12 cascades) Inner Fixed Structure Assembly The inner fixed structure assembly is composed of the following components: S the 12 o’clock hinge beam S the 6 o’clock latch beam S the Inner Fixed Structure (IFS) TRANSLATING STRUCTURE The translating structure is composed of the following components: S the translating cowl (including two translating sleeves), S the blocker doors and links (six for each translating cowl).



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PRECOOLER SCOPE

TCC INLET

FINGER SEALS

CASCADES 12 POSITIONS

BLEED VALVES

BLOCKER DOORS (6 POSITIONS)


J-RING

INNER FIXED STRUCTURE

LINKS (6 POSITIONS) TRANSLATING COWL

Weight: 650 kg (1 430 lbs) per Cowl Door Figure 164 FRA US/T

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FAN EXHAUST COWL This component is dedicated to be installed on the utboard nacelles in replacement of the Thrust Reversers. The Fan Exhaust Cowls (FEC) therefore share the same external interfaces as the Thrust Reverser. It differs from the Thrust Reverser Cowl Doors by : S No ETRAS S No blocker doors and links S No translating structure S No latch L8 S Carrying a Core Pressure Relief Door If the Core Pressure Relief Door is opened, the red popout goes out and is visible from the ground.



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Weight: 380 kg (840 lbs) per Cowl

View from internal side

No ETRAS

Pressure Relief Doors (left side)


Pop Out

View from external side

No Blocker Doors No Links No Translating Structure

No Latch L8

Figure 165 FRA US/T

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RR Trent 900

78

FAN EXHAUST COWL/THRUST REVERSER COWL Opening The fan exhaust cowl / thrust reverser cowl doors can be opened for maintenance purposes on the engine. The unlatching sequence is carried out from the latch access door and the latches all installed at the bottom of the fan exhaust cowl / thrust reverser cowl. Unlocking of these latches is done in a defined sequence: S the Latch L1 is opened first by pulling the lock trigger and the handle to the down position, S the Latch L6.1 handle has to be fully open, S the Latch L6.2,

CAUTION:

MAKE SURE THAT THE WIND SPEED CONDITIONS ARE NOT MORE THAN 45 KNOTS.

CAUTION:

BEFORE OPENING THE FAN EXHAUST COWLS / THRUST REVERSER COWLS, MAKE SURE THAT SLATS ARE RETRACTED AND THAT THEY CANNOT MOVE, TO PREVENT FROM POSSIBLE INTERFERENCES.

CAUTION:

BEFORE OPENING THE FAN EXHAUST COWLS / THRUST REVERSER COWLS OF INBOARD ENGINES (2 OR 3), MAKE SURE THAT THE THRUST REVERSER SYSTEM HAS BEEN DEACTIVATED FOR MAINTENANCE.

A push/pull cable connects the latch L6.1 to the latch L7. When you unlatch the latch L6.1, you unlatch the latch L7 at the same time. S the Latch L5.2 S the Latch L5.1, S the Latch L4, S the Latch L3, S the Latch L2. Once the fan exhaust cowl / thrust reverser cowl doors are unlocked, the opening is done from the switch box. There is one switch box per side. The maintenance personnel must push the UP switch until the fan exhaust cowl / thrust reverser cowl door is opened and the HOR is locked. When the HOR is locked, the green stripe is visible. Then the maintenance personnel must push the DOWN switch to hold the fan exhaust cowl / thrust reverser cowl on the HOR. The fan exhaust cowl / thrust reverser cowl doors have two open positions: S initial position of 35 degrees, S full open position of 45 degrees.


NOTE:

NOTE:

FRA US/T

To open the cowl up to 45 degrees, it must be open up to 35 degrees before and hold open rod have to be put on its 45 degrees fitting.

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POWER PLANT EXHAUST

Figure 166 FRA US/T

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POWER PLANT EXHAUST

A380

RR Trent 900

78



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POWER PLANT EXHAUST

Figure 167 FRA US/T

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RR Trent 900

78

Specific Latch for Inboard Engines In addition, the two fixed halves of the thrust reverser structures for the inboard engines are connected together by an eighth latch L8. The latch L8 is composed of a telescopic locking system permanently connected to the LH structure at 12 o’clock and a pin latch at 6 o’clock position. A handle controls this latch, and locks/unlocks simultaneously the 6 o’clock pin latch via a command rod and the 12 o’clock latch. The 12 o’clock latch is linked to the 6 o’clock pin latch by the push/pull cable. To open the thrust reverser cowls, you must: S Turn counterclockwise the handle, S Pull the handle until the red strip becomes visible to unlock the latch L8, S Turn clockwise the handle. To close the cowls, you must make sure that the latch L8 returned correctly in its stored position. You must also secure it by re−installing the safety ball pin. L8 is the first latch to be opened for the opening sequence of the thrust reverser cowl on the inboard engines.


NOTE:

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POWER PLANT EXHAUST

Figure 168 FRA US/T

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A380 78

The Hold Open Rod (HOR) Each Thrust Reverser Cowl Door opens and is maintained in the opened position by one HOR. The other mean of retention is the cowl door opening actuator. Two opening positions : 35° and 45°. The HOR keeps the Thrust Reverser Cowl Door in the opened position for ground maintenance. The ends of the HOR are attached to a fitting on the Thrust Reverser Cowl Door and to a bracket on the engine. The 35° fitting is the storage fitting. It is necessary to move the HOR on a second forward frame fitting to open the Thrust Reverser Cowl Door in the 45° position. Coloured flags enable to know the HOR state : S red indicator : unlock S green indicator : lock



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Hold Open Rod in stored position


35° fitting

45° fitting Hold Open Rod in 45° opened position Figure 169 FRA US/T

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Hold Open Rod

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RR Trent 900

78

Fan Exhaust Cowl/Thrust Reverser Cowl Closing At the end of maintenance tasks on the engine, the fan exhaust cowl / thrust reverser cowl doors have to be closed to put the aircraft back into operation. First of all, the maintenance personnel must push the UP switch to unload the HORs. The HOR is then unlocked and in this case, the red stripe is visible. The DOWN switch must be pushed and held until the fan exhaust cowl / thrust reverser cowl door closes completely. You cannot close the cowl from the 45 degrees position to the 35 degrees position directly. At 35 degrees, put the hold open rod on the 35 degrees HOR fitting. The locking of these latches is done in a defined sequence: S the Latch L2 first, S the Latch L3, S the Latch L4, S the Latch L5.1, S the lLatch L5.2, S the Latch L6.1 handle has to be fully closed, NOTE:

A push/pull cable connects the Latch L6.1 to the Latch L7. When you latch the Latch L6.1, you latch the Latch L7 at the same time. S the Latch L6.2, S and at the end, the Latch L1 is closed by pulling the lock trigger and the handle upward. Once the latches are locked, the latch access door has to be closed.


NOTE:

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POWER PLANT EXHAUST

Figure 170 FRA US/T

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Manual Opening/Closing Each fan exhaust cowl / thrust reverser cowl is equipped with an opening actuator, which has a MDU (Manual Drive Unit). This MDU can be used for the opening/closing of the fan exhaust cowl / thrust reverser cowl when the electrical power is off, or in case of failure of the electrical functions in the cowl opening system. The fan exhaust cowl / thrust reverser cowl can be open manually to 35 or 45 degrees. To open the fan exhaust cowl / thrust reverser cowl manually, you must: S Remove the MDU cap S Put the tool in the MDU 3/8” square S Use the tool to open the actuator to 35 degrees position (clockwise) until HOR locking S Use the tool to release the weight of the cowl on the HOR (counterclockwise) S To open the cowl to 45 degrees position, put the HOR on its 45 degrees fitting, and use the tool to open the actuator to 45 degrees position (clockwise) until HOR locking


NOTE: The HOR makes a rattling noise at the locked position. S Use the tool to release the weight of the cowl on the HOR (counterclockwise) S Remove the tool and reinstall the cap at the end. NOTE:

The HOR locking sleeve must slide to show the green stripe (locked position) and hide the red stripe.

CAUTION:

IF POWERED TOOLS ARE USED, ONLY USE POWERED TOOLS WITH A TORQUE LIMITER TO MAXIMUM VALUE 133.5 IN.LBS (15 N.M).

CAUTION:

TURN THE TORQUE WRENCH SLOWLY WHEN THE HOLD OPEN ROD IS NEAR THE LOCKED POSITION TO PREVENT DAMAGE TO THE HOLD OPEN ROD AND TO THE THRUST REVERSER STRUCTURE.

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THRUST REVERSER INHIBITION Power Inhibition for Maintenance To make sure that the thrust reverser system is unserviceable for maintenance, the TRPU (Thrust Reverser Power Unit) has to be deactivated by inhibiting the power supply of the thrust reverser system. The TRPUs are installed on the inboard engines only, under the LH fan cowl door. In normal operation, the TRPU is powered with 115 VAC 3 phase by the EIPM logic. The TRPU then energizes the ETRAC (Electronic Thrust Reverser Actuation Controller), which will supply the PDU (Power Drive Unit) to control the actuators. So the power supply inhibition requires the removal of the ball pin from the TRPU and to turn the TRPU lever to the ”INHIBITED” position. The ball pin must be re−installed. When the TRPU lever is in the ”INHIBITED” position, the ETRAC is no longer supplied. A lock−out pin with the ”REMOVE BEFORE FLIGHT” flag must be installed in the TRPU hole.



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Power Inhibition and Mechanical Inhibition Before Flight To make sure that the thrust reverser system is unserviceable for flight, the TRPU has to be electrically deactivated and the two translating cowls mechanically deactivated. The mechanical inhibition device of the thrust reverser is accessed by loosening the four captive screws on the mechanical inhibition access panel installed on the rear lower part of the thrust reverser cowls. The mechanical inhibition requires the removal of the ball pin from the mechanical device. By using a lever, the mechanical inhibition device is then set to the ”INHIBITED” position. The ball pin has to be re−installed to lock the mechanical inhibition device. When the thrust reverser system is mechanically inhibited a red pop−out is visible on the inhibition access panel. TO INHIBIT THE THRUST REVERSER SYSTEM YOU MUST INHIBIT IT ON ENGINE 2 AND 3, REFER TO MMEL + CDL.


WARNING:

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THRUST REVERSER CONTROL COMPONENT DESCRIPTION Major Component Identification The major components of the ETRAS are installed on the forward frame of the LH (Left Hand) fan exhaust cowl: S the PDU is installed on the upper part, S the ETRAC and TRPU are installed on the middle side, S the TLS Power Unit is installed on the lower part.



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THRUST REVERSER CONTROL FUNCTION OPERATION General The A380−800 Trent 900 Electrical Thrust Reverser Actuation System (ETRAS) is an electro−mechanical system which allows the translating cowls of the engine 2 & 3 to be deployed and stowed in response to electrical commands from the EEC and from the aircraft interfaces. The thrust reverser assembly is installed at the aft part of the nacelle, only on the aircraft inboard engines (No. 2 & 3). The assembly is a conventional fixed cascade translating cowl blocker door type. It is made of two halves that make a duct around the engine. Each halve has a fixed structure, which is used as a support for the cascades, the actuation system and the translating cowl. Both engine−translating cowls are mechanically linked and slid onto the thrust reverser upper and lower tracks. The Thrust Reverser halves open at the 12 o’clock hinge beam to give access to the engine during maintenance operations. The ETRAS carries out the following functions: S deployment of the thrust reverser translating cowls when the deploy command is set, S stowage of the thrust reverser translating cowls when the stow command is set, S avoidance of inadvertent deployment of the thrust reversers, S manual deployment and stowage of the translating cowls for maintenance, S manual inhibition and deactivation of the translating cowls for maintenance. Architecture The Deploy command has three independent electrical command lines upon a reverser thrust selection on the throttle control assembly: S an aircraft 115 VAC power supply commanded by the flight/ground control PRIM to the tertiary lock system, S an aircraft 155 VAC from EIPM to TRPU, S an electrical command from EEC to ETRAC. For ETRAS monitoring, fault reporting and BITE test, the EEC communicates with the OMS (On−board Maintenance System) and CDS (Control and Display

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System) via ADCN. For maintenance equipments, a thrust reverser operational test (deploy/stow) is available on the OMS. The ETRAS is basically composed of: S ETRAC (Electronic Thrust Reverser Actuation Controller), S TRPU (Thrust Reverser Power Unit), S PDU (Power Drive Unit) electrical motor, S 6 ball screw actuators mechanically driven through a synchronizing flexible shaft power train system from the PDU. Actuation The ETRAS (Electrical Thrust Reverser Actuation System) operates in normal mode, when the following initial conditions are met: S aircraft on the ground, S engines are running, S and Throttle Lever in Reverse thrust position. The Electrical Power is supplied from the aircraft to the TRPU. The TRPU supplies electrical power through the ETRAC to all the electrical components. The ETRAC releases all the locks and the PDU brake. Electrical power is transformed into mechanical power by the PDU. The PDU is composed of: S a motor and a resolver assembly, S a brake assembly. The disc brake of the PDU needs to be energized for release. When the brake solenoid is de−energized, the disc brake engages: S to maintain preload of actuation system in fully stowed position, S to lock the T/R in fully deployed position.

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The electrical motor of the PDU gives torque and rotational speed to the flexible shafts, S Mechanical power is then distributed to middle ball−screw actuators by 2 flexible shafts. S Mechanical power is distributed to the other 4 actuators by flexible shafts. Middle actuators have a MDU (Manual Drive Unit) which allows the manual deployment / stowage for maintenance operations. There are two primary locks, one on the top right actuator and one on the top left actuator. These internal locks are part of retention means of the thrust reverser system. Their function is to lock the thrust reverser when stowed. Two resolver sensors mounted on the lower actuators monitor the position of the translating cowls. The EEC detects that: S the upper translating actuators (LHS and RHS) are locked, through the two primary lock system proximity sensors. S the translating cowls are in the stowed position through the lower actuator position cowl resolvers. The ETRAC implements the ETRAS control functions except for tertiary lock. The ETRAC commands the left PLS to unlock for deployment, and through the TRPU: S the right PLS and the disc brake to unlock for deployment, S the disc brake to engage at the end of the deploy sequence, to secure the T/R in fully deployed position, S the disc brake to unlock for stowing, S the electrical motor of the PDU for deployment or stowing, S provides monitoring data to the EEC, including ETRAS BITE results, data of the TRPU internal power switch. The TLS (Tertiary Lock System) is installed on the nacelle structure at the rear bottom of the left translating cowl. The function of the TLS is to lock the thrust reverser when it is stowed, in order to prevent an inadvertent deployment, mainly in flight. The TLS design follows a fail−safe motion in which the TLS engages into a locked position when the electrical power is removed.

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The TLS (Tertiary Lock System) is mechanically locked. The Tertiary Lock System must be electrically released to allow deployment. Two proximity sensors are mounted on this Tertiary Lock System. The EEC detects that the TLS is locked or unlocked through the two TLS proximity sensors. First Defense Line When The EEC detects that the aircraft is on the ground (LGERS discrete signal) and TRA (Throttle Reverser Angle) thresholds are reached (−9 for deploy signal and −8 for stow signal), the EEC sends to the ETRAC (Electronic Thrust Reverser Actuation Controller) deploy/stow order for thrust reverser operation. Second Defense Line Actuating as the second line of defense of the ETRAS: S The Engine Interface Power Management (EIPM) will control the switching of low power supply (28 VDC) to the ETRAC for basic control of the thrust reverser system in normal operation and during maintenance operation when the aircraft is on the ground (LGERS discrete signal). S The EIPM controls and monitors the switching of the 115 VAC 3 phases power supply to the TRPU. Third Defense Line The PRIM (PRIMary flight control and guidance computer) installed in the avionic bay will control the switching of the SSPC (Solid State Power Controller) providing the third line of defense of the ETRAS system. The 115 VAC power supply for the Tertiary Lock System will be transformed and rectified into DC voltage through a TLS Power Unit. The thrust reverser tertiary lock is the third line of defense to avoid an inadvertent deployment in flight. It stops the mobile structure in case of failure of the primary locks. The tertiary lock is composed of one electro−mechanical lock, installed on the left 6 o’clock beam. The tertiary lock can be manually deactivated in the unlock position to manually deploy the sleeves to get access to the cascades. Two proximity sensors send the TLS position to the EEC.

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Deploy Sequence The two translating cowls are initially stowed. The EEC detects that: S the upper translating actuators (LHS and RHS) are locked through the proximity sensor signals of the Primary Lock System (PLS). S the translating cowls are in the stowed position through the translating cowl resolver signals. The deploy command is set. The third defense line closes alternative current contactor and energizes the TLS once the TRA (Throttle Resolver Angle) is detected below the 4.5 degrees position. The EIPM (Engine Interface Power Management) will command the 115 VAC (3 phases) power supply at the TRPU (Thrust Reverser Power Unit) input once the TRA is detected below the − 7 degrees position and ETRAC is supplied with 28 VDC. The EEC confirms that the TLS is released through the TLS sensor A & B signals. The deploy command will be sent by the EEC to the ETRAC through ARINC 429 bus. The EEC will apply an hysteresis of 0.9 degrees on the throttle position: the throttle deploy condition will be true when the selected TRA is below − 9.0 degrees and will remain true until the selected TRA goes above − 8.1 degrees. The engine throttle lever moves to a position below − 9.0 degrees. The TRPU is distributing the electrical power to all the electrical components through ETRAC, which commands locks and the brake to be released. The PDU (Power Drive Unit) transforms the electrical power into mechanical power. The mechanical power is distributed to: S the two middle ballscrew actuators by two synchro flex shafts. S the upper and lower actuators by four other flex shafts and allow the translating cowl to move in the deployment position. The EEC detects that: S the PLS are unlocked through the PLS unlock proximity sensor signals. S the translating cowls are no more in the stowed position through left and right translating cowl resolvers signals.

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The time for both translating cowls to deploy is monitored. When both translating cowls reach 80 % of the full stroke, the EEC detects that the thrust reverser is fully deployed through left and right translating cowl resolvers signals. Near full deploy position; the speed is reduced to slow. When 100 % of the stroke is reached, the end actuator hard stop is engaged. The motors stall at low speed and force limit. The ETRAC de−energizes lock drivers, brake drivers and disables inverter. The aircraft opens alternative current contactor and may de−energize the Tertiary Locking System. Stow Sequence When the Aircraft is on the ground and a deploy command has previously been executed, or partially executed, the pilot selecting forward thrust will cause the EEC to initiate a stow command. The EEC will send the stow command to the ETRAC via the data bus which will then release the brake and command the motor to rotate in the opposite direction drawing the sleeves to close. The stow command will be transmitted continuously by the EEC to the ETRAC until the EEC detects the thrust reverser to be fully stowed. The STOW sequence follows different steps: The system is initially in the deployed position. The engine throttle lever moves to the forward position and above − 8 degrees. The EEC detects the TRA position above − 8 degrees. The stow command sent by the EEC to the ETRAC through ARINC 429 bus. The aircraft closes alternative current contactors and energizes the TLS once the TRA is detected up to − 4.5 degrees position. The TRPU supplies the electrical power to all electrical components through the ETRAC, which commands the locks and brake to be released. The electrical power is transformed into mechanical power by the PDU.

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The mechanical power is supplied to: S the two middle ballscrew actuators by the two synchro flex shafts. S the upper and lower actuators by four other flex shafts and let the translating cowl move in the stowage position. S The EEC detects that: S the PLS are unlocked through the PLS unlock proximity sensor signals. S the translating cowls are no more in the deployed position through left and right translating cowl resolver signals. The translating cowls reach the position at which the tertiary lock is mechanically locked. The EEC detects that: S the TLS is locked through the TLS sensor signals. S the PLS are locked through the PLS unlock proximity sensor signals. S the translating cowls are in the stow position through the left and right translating cowl resolvers signal. S the thrust reverser is stowed and locked. The EIPM switches off the 115 VAC (3 phases) power supply at the TRPU input at the end of the stow sequence when the EEC indicates that the thrust reversers are locked with a confirmation of 1 second. The ETRAC 28 VDC will be isolated by the EIPM.



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Summary The following schematic summarizes the deploy and stow sequence.



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THRUST REVERSER MAINTENANCE Manual Deploy/Stow the Thrust Reverser Translating Cowl The procedure to manually deploy the thrust reverser translating cowl is: S Make the thrust reverser unserviceable for maintenance (TRPU (Thrust Reverser Power Unit) inhibition), S Make the TLS (Tertiary Lock System) of the thrust reverser unserviceable, S Unlock the PLS (Primary Lock System) of left thrust reverser S Release the brake of the PDU (Power Drive Unit), S Unlock the PLS of right thrust reverser S Make the right and left MDUs (Manual Drive Units) of the thrust reverser operative, S Manually deploy the thrust reverser translating cowl, S Make the right and left MDUs of the thrust reverser inoperative. The procedure to manually stow the thrust reverser translating cowl is: S Make the thrust reverser unserviceable for maintenance, S Make the TLS of the thrust reverser unserviceable, S Make the right and left MDU of the thrust reverser operative, S Active the PLS of the right thrust reverser S Active the PLS of the left thrust reverser S Manually stow the thrust reverser translating cowl, S Make the right and left MDU of the thrust reverser inoperative, S Active the brake of the PDU.



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Make the Thrust Reverser Unserviceable for Maintenance YOU MUST MAKE THE THRUST REVERSER UNSERVICEABLE (INSTALLATION AND SECURIZATION OF THE INHIBITION DEVICE) BEFORE YOU DO A WORK ON OR AROUND THE THRUST REVERSER. IF YOU DO NOT INSTALL AND SECURE THE INHIBITION DEVICE, YOU CAN CAUSE ACCIDENTAL OPERATION OF THE THRUST REVERSER AND INJURY TO PERSONS AND/OR DAMAGE TO THE EQUIPMENT. The opening of fan cowl doors gives access to the TRPU. To make the TRPU unserviceable, you must: S Remove the ball pin from the TRPU, S Move the TRPU lever to the ”inhibited” position, S Install the ball pin on the TRPU,

WARNING:

CAUTION:

AFTER INSTALLATION OF THE BALL PIN, CHECK THAT THE BALL PIN IS CORRECTLY INSTALLED BY PULLING IT.


S Install the lock out pin with the ”Remove Before Flight” flag in the TRPU hole.

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Make the TLS of the Thrust Reverser Unserviceable WARNING:

IF YOU INHIBIT THE THRUST REVERSER ENGINE 2, YOU MUST INHIBIT THE THRUST REVERSER ENGINE 3 AND OPPOSITE.

YOU MUST MAKE THE THRUST REVERSER UNSERVICEABLE (INSTALLATION AND SECURIZATION OF THE INHIBITION DEVICE) BEFORE YOU DO A WORK ON OR AROUND THE THRUST REVERSER. IF YOU DO NOT INSTALL AND SECURE THE INHIBITION DEVICE, YOU CAN CAUSE ACCIDENTAL OPERATION OF THE THRUST REVERSER AND INJURY TO PERSONS AND/OR DAMAGE TO THE EQUIPMENT. Thrust reverser TLS is located at the lower part of the thrust reverser behind the mechanical inhibition access door To make the thrust reverser TLS unserviceable, you must: S Remove the ball pin from the TLS, S Move the yellow lever on the ”UNLOCKED” position, S Install the ball pin to lock the lever. When the access door is closed, make sure that the visual pop−out is visible. WARNING:



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Unlock/Active the PLS of the Thrust Reverser There are two PLSs per engine. The opening of fan cowl doors gives access to PLSs To unlock the PLS of the thrust reverser, you must: S Remove the inhibition pin from the end of the lever, S Rotate the lever to put the translating axis knob in the ”UNLOCKED” position, S Install the inhibition pin at the end of the lever. To active the PLS of the thrust reverser, you must: S Remove the inhibition pin from the end of the lever, S Rotate the lever to put the translating axis knob in the ”ACTIVE” position, S Install the inhibition pin at the end of the lever.



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Release/Active the Brake of the PDU The PDU (Power Drive Unit) is installed behind the fan cowl doors. To release the brake of the PDU, you must: S Move the yellow lever to the ”UNLOCKED” position. To active the brake of the PDU, you must: S Move the yellow lever to the ”ACTIVE” position.



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Make the MDUs of the Thrust Reverser Operative/Inoperative There are two MDUs (Manual Drive Units) per engine. They are installed behind the fan cowl doors. To Make the MDUs of the thrust reverser operative, you must: S Make sure that the yellow levers of the two MDUs are on the ”ACTIVE” position, S If not, move them on the ”ACTIVE” position. To Make the MDUs of the thrust reverser inoperative, you must: S Make sure that the yellow levers of the two MDUs are on the ”LOCKED” position, S If not, move them to the ”LOCKED” position.



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Manually Deploy/Stow the Thrust Reverser Translating Cowl To manually deploy the translating cowl of the thrust reverser, you must: S Put the 3/8 inch speed wrench at the MDU, NOTE: NOTE:

make sure that the male square drive of the speed wrench is correctly engaged in the MDU, before you turn the MDU. you can use special pneumatic or electrical tool to turn the MDU.

S Push and turn the MDU clockwise to deploy the translating cowl of the thrust reverser until the MDU torque limiter releases, The translating cowl can be deployed either with the left MDU or with the right MDU. S Remove the speed wrench. To manually stow the translating cowl of the thrust reverser, you must: S Put the 3/8 inch speed wrench at the MDU, NOTE:

NOTE: NOTE:

Note: make sure that the male square drive of the speed wrench is correctly engaged in the MDU, before you turn the MDU. you can use special pneumatic or electrical tool to turn the MDU.

S Push and turn the MDU counter−clockwise to stow the translating cowl of the thrust reverser until the MDU torque limiter releases, The translating cowl can be stowed either with the left MDU or with the right MDU. S Remove the speed wrench.


NOTE:

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EXHAUST The Exhaust system is composed of two parts, the Nozzle and the Plugs (Rear and Front). It is an acoustically treated structure that provides flow contour for engine exhaust gas. The upper part of the Nozzle is equipped with finger seals, which are fire seals. There are three spigots on the Nozzle. Spigots are locators and ease the Nozzle removal and installation.



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Weight:

Nozzel Finger Seals

50 kg (110 lbs) for Nozzle 35 kg (75 lbs) for FWD Plug 15 kg (35 lbs) for Rear Plug

3 Spigot


Rear Plug

Forward Plug

Nozzle

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24 VARIABLE FREQUENCY GENERATOR (VFG) Purpose The function of the VFG is to supply 115 V ac, 3− phase power for use in the aircraft electrical systems. Location The VFG is installed on the rear left side face of the external gearbox of each engine. Description The VFG is an AC generator, which provides 115v AC 3−phase variable frequency electrical power to the aircraft. The VFG is attached directly to the gearbox via studs on the gearbox rear face.


VFG Oil System The VFG has an integral oil system to lubricate and keep it cool. The oil is cooled by fan air passing through a heat exchanger, which is then vented overboard through the left fan cowl. There is also a filter to remove contaminants from the oil, which has a pop out indicator to show when it is blocked. VFG Air/Oil Heat Exchanger Location The heat exchanger is installed on the lower left side of the fan case and consists of: A fin−and−plate heat exchanger An outlet duct A seal which abuts the inner surface of the left fan cowl door. Functional Description The air/oil heat exchanger dissipates the VFG heat by exchanging heat between the VFG oil and engine fan air. The fan air inlet is on the inner surface of the fan case and allows air to enter the heat exchanger which is bolted to the fan case outer surface. The heat exchanger has a bypass valve, which allows the oil to bypass the cooler matrix when the oil temperature is low.

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24 VFG − OIL SERVICING Description The VFG (Variable Frequency Generator) is mounted on the engine accessory gearbox located in the engine nacelle. The filling ports and a sight glass are mounted on the LH side of the VFG. The VFG is cooled and lubricated by an internal oil cooling system with an engine mounted, called the ACOC (Air Cooled Oil Cooler). This ACOC uses airflow from the fan case to cool the VFG oil. The cooling system limits the temperature VFG oil inlet to 125_C. A manual disconnection of a faulty VFG could be done through the DRIVE P/B on the overhead ELEC panel (1225 VM). When disconnected, VFG cannot be reconnected and should be removed from the A/C.



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ELECTRICAL POWER

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OIL SERVICING Level Check There are two ways to do a check of the oil level: S on the vertical sight glass, directly installed on the VFG, S with the ROLS (Remote Oil Level Sensor). The ROLS gives indications on the CMS (Central Maintenance System), through the servicing report or on the ECAM in case of a low level. The ROLS has three sensors for three different levels: S an oil level overfill sensor which results in the message ”FULL/OVERFILL OIL LEVEL”, S a low level sensor, which gives the message ”LOW OIL LEVEL” (500 hours before refill), S a very low level sensor which gives the message ”VERY LOW OIL LEVEL”. Those messages are presented on the SD status page. They appear as well as on the servicing report or by using the CMS via the fault flight report. The ROLS operates on ground four minutes after engine shutdown. NOTE:

Note: To access to the VFG, it is necessary to open the fan cowl.


DPI The DPI (Differential Pressure Indicator) extends if the oil filter is clogged. At the same time, it triggers a message through the GGPCU (Generator and Ground Power Control Unit) to the CMS. In that case, applicable procedure for oil filter check should be done.

WARNING:

YOU MUST BE CAREFUL WHEN YOU DO WORK ON THE ENGINE PARTS AFTER THE ENGINE IS SHUTDOWN. THE ENGINE PARTS CAN STAY HOT FOR ALMOST 1 HOUR.

WARNING:

YOU MUST NOT LET ENGINE OIL STAY ON YOUR SKIN. FLUSH THE OIL FROM YOUR SKIN WITH WATER.

WARNING:

YOU MUST NOT BREATHE THE FUMES.

WARNING:

YOU MUST NOT GET ENGINE OIL IN YOUR EYES OR IN YOUR MOUTH. PUT ON GOGGLES AND A FACE MASK.

WARNING:

IF YOU GET ENGINE OIL IN YOUR MOUTH, YOU MUST NOT CAUSE VOMITING BUT GET MEDICAL AID IMMEDIATELY.

HUMAN FACTOR POINTS: CAUTION:

TO PREVENT DAMAGE, DO NOT DO THE SERVICING OF A DISCONNECTED VFG.

CAUTION:

DO NOT USE DEVICES OTHER THAN THE APPROVED OVERFLOW DRAIN−HOSE FITTING. HARD METAL OBJECTS SUCH AS SCREWDRIVERS CAN CAUSE DAMAGE TO THE OVERFLOW DRAIN−VALVE SEAT.

CAUTION:

USE ONLY NEW OIL CANS, WHEN YOU FILL THE VFG WITH OIL OR ADD OIL TO THE VFG. THE CONTAMINATION IN OIL THAT STAYS IN OPEN CANS CAN CAUSE FAST DETERIORATION OF THE OIL AND WILL DECREASE THE LIFETIME OF THE VFG.

CAUTION:

DO NOT USE SOLVENTS THAT CONTAIN CHLORINE TO CLEAN THE EQUIPMENT (PUMP, HOSES, TANK AND FUNNEL) USED TO FILL THE VFG WITH OIL. CHLORINE CONTAMINATION OF THE OIL CAN CAUSE FAST DETERIORATION OF THE OIL AND WILL DECREASE THE LIFETIME OF THE VFG

Refilling For the refilling connect the pressure fill hose of the pump−hand, oil filling−up to the pressure fill port. The oil overflow is collected from the over fill drain port into a container. A visual check could be done on the sight glass. Human factor points:

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ELECTRICAL POWER

Figure 190 FRA US/T

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Oil Servicing − Level Check ... Refilling 03 |24 |L2

Page 373


FIRE PROTECTION ENGINE

A380

RR Trent 900

26

ATA 26 FIRE PROTECTION FIRE/OVERHEAT DETECTORS Purpose The engine fire detection assemblies monitor the temperature in the engine zones. Location Each detector assembly has two elements (Loop A and Loop B) which attach to a support tube. The two elements run parallel to each other along the support tube and monitor the temperature along their length. They provide a continual analog output to the conversion module for the engine. Quick release clamps and bushing support the elements along their length. There are five detector assemblies as follows: Zone

Position

1

1

Fancase - forward side of gearbox

2

1

Fan case - rear side of casebox

3

2

IP Comp/Intermediate case - lower

4

3

Above Combustor attached to pylon

5

3

LP Turbine


Assembly

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FIRE PROTECTION ENGINE

Figure 191 FRA US/T

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Fire Detection Loop Location 26−01 |L3

Page 375


HYDRAULIC POWER

A380

RR Trent 900

29

ATA 29 HYDRAULIC POWER HYDRAULIC SYSTEM Purpose The engine driven hydraulic pumps (2) is the primary pump for the aircraft hydraulic systems. Location The hydraulic pumps are installed on the right side front face of the external gearbox of each engine.


Description The hydraulic pump is a variable displacement pump. The pump is designed to operate at a nominal pressure of 5000 psi (343 bar). The pump is fitted with an electrically operated Hydraulic Pump Offload Solenoid. The EEC controls the solenoid. Case drain hydraulic flow cools and lubricates the engine driven pump. There is a non−bypass case drain filter with visual indication of blockage installed on the right side of the fan case, one for each pump. A ripple damper smoothes the pump pressure output. Functional Description The engine external gearbox drives the hydraulic pump when the engine is operating. Pump pressure goes to the hydraulic system when the offload solenoid valve is not energised. To improve engine inflight restart capability, the EEC commands a relay in either channel to switch Aircraft 28 V dc to operate the Hydraulic Pump Offload Solenoid to de−pressurise the hydraulic system during windmill starts. An indication lamp is illuminated in the cockpit when the Hydraulic Pump Offload Solenoid is energised.

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29



HYDRAULIC POWER

Figure 192 FRA US/T

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Hydraulic System 01 |29 |L2

Page 377

A380

RR Trent 900

29 ENGINE DRIVEN PUMP DESCRIPTION HYDRAULIC PUMP Disengagement/Re−engagement Open the fan cowl to access the EDP (Engine Driven Pump). It is possible to do a mechanical disengagement of the EDP (Engine Driven Pump) from the engine gearbox by means of the declutch system. In flight, the EDP disengagement will be operated through an electrical solenoid, from an electrical switch (28VDC) on the hydraulic panel. In this case, both pumps of the given engine will be de−clutched. The disengagement of the EDP is irreversible in flight until a specific maintenance action is done. On the ground, for maintenance purpose, the EDP can be manually de−clutched by pulling a ring outward. The re−engagement of the EDP can be made manually (engine shut down) by acting on the reset port. WARNING:

WHILE APPROACHING THE ENGINE, THE AIR INTAKE SUCTION OR EXHAUST BLOW COULD INJURE. THEREFORE ACCESSING THE ENGINE FROM ITS SIDE WITHIN THE SAFETY AREA WILL PREVENT THIS.

CAUTION:

WHEN MANUALLY RE−ENGAGING THE EDP, IT IS POSSIBLE NOT TO REARM PROPERLY THE SYSTEM, LEADING TO AN INADVERTENT DISENGAGEMENT WHEN THE ENGINE WILL BE RUNNING (VIBRATION EFFECT). TO PREVENT THIS, TURNING THE RESET SHAFT UNTIL THE RESET MARK WILL FULFILL A PROPER RE−ENGAGEMENT OF THE PUMP.



HYDRAULIC POWER

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29



HYDRAULIC POWER

Figure 193 FRA US/T

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Hyd. Pump Dis-/Re−Engagement 1 02 |29 |L2

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RR Trent 900

29



HYDRAULIC POWER

Figure 194 FRA US/T

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RR Trent 900

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HYDRAULIC POWER

Figure 195 FRA US/T

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Page 381


PNEUMATIC

A380

RR Trent 900

36 AIRCRAFT PNEUMATIC SYSTEM Description Air for the pneumatic system can come from the following: S Engine compressors S Auxiliary power unit (APU) S Ground air source

Temperature Regulation A precooler and Fan Air Valve (FAV) installed in the Aircraft pylon achieves bleed air temperature regulation using LP compressor (fan) air.

APU Bleed The Auxiliary Power Unit (APU) is the primary source of compressed air on the ground. The APU can also be used to supply compressed air in flight through the APU Bleed Valve. With APU air bleed supply, the Crossbleed Valve is automatically opened and the Pressure Regulating Valves (PRV) are automatically closed.


Ground Supply A ground air source is an alternative to the APU for the supply of compressed air on the ground. There are three High−Pressure (HP) ground connectors installed on the Aircraft Engine Bleed The engines are the primary source of compressed air in flight. The air is bled from the 8th stage of the IP compressor and 6th stage of the HP compressor. Depending on engine speed, air is tapped off either the IP compressor (IP8) or the HP compressor (HP6). The IP8 Bleed Check Valve protects the HP compressor from reverse flow. The HPV is fitted with a dual solenoid and is opened and closed by the Engine Electronic Control System (EEC). This will ensure maximum efficiency from the engine. Pressure Regulation Bleed air from the IP8 and HP6 is ducted to the Pressure Regulating Valve (PRV), which regulates downstream pressure. An Overpressure Valve (OPV) protects the precooler and downstream user systems against potential overpressure.

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PNEUMATIC

Figure 196 FRA US/T

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Aircraft Bleed System 01 |36 |L3

Page 383

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RR Trent 900

36 ENGINE BLEED AIR SUPPLY Component Location The engine bleed air supply system consists of the following: S IP8 Bleed Check Valve S HP6 Bleed Valve (HPV) S Pressure Regulating Valve (PRV) S Bleed Air Ducts The engine bleed air supply system valves and ducts are installed on the left side of the core engine.



PNEUMATIC

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PNEUMATIC

Figure 197 FRA US/T

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Engien Bleed Air Supply 02 |36 |L3

Page 385


ICE & RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

A380 30−20

ATA 30−20 ENGINE AIR INTAKE ICE PROTECTION ENGINE ICE PROTECTION AREAS Introduction Ice may form on the leading edge of the Inlet Cowl, Spinner and P20/T20 probe when the engine is operating in conditions of low temperature and high humidity. Ice build up could affect engine performance and could cause damage to the compressor from ice ingestion. To prevent ice formation, anti−icing protection is provided to the following areas: S The Inlet Cowl leading edge (Thermal) S The P20/T20 Probe (Thermal) S The Spinner (Dynamic) Inlet Cowl Leading Edge The area inside the “D“ chamber on the inlet cowl leading edge is heated by hot air from the HP compressor stage 3 when the ENG anti−ice system is selected on.


Spinner A solid rubber tip that vibrates naturally to break up and dislodge the ice immediately it starts to form, protecting the spinner from ice build up. P20/T20 Probe The P20/T20 probe is heated by a single electrical heating element during engine operation. The electrical power for heating the probe is provided by the aircraft 115VAC supply via the Engine Interface Power Management (EIPM) unit and controlled by relays in the EEC which either channel can control. In the event of EIPM failure, airframe 115VAC is permanently available to the probe heater whenever the airframe electrical network is powered. After engine shutdown on the ground, the probe heater is powered for a period of 15 minutes maximum In the case of an engine fire, in flight or on the ground, the probe heater 115VAC power supply is removed immediately following operation of the fire handle.

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A380 30−20




Figure 198 FRA US/T

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Engine Ice Protection Areas 01 |30−20 |L3

Page 387



A380 30−20

AIR INTAKE COWL Air Intake Cowl Presentation The Air Intake Cowl is an interchangeable component attached to the engine fan case. The function of the Air Intake Cowl is to provide: S Smooth and sufficient air flow to the engine S Smooth air flow over the nacelle outer surfaces S Engine noise attenuation by using acoustic materials Full acoustic attenuation treatment by using zero splice contact. The Air Intake Cowl contains the ice protection system. The Air Intake Cowl incorporates the interphone jacks and the engine P20/T20 probe. Panels are provided on the Air Intake Cowl to give quick access to the internal components. The Air Intake Cowl main components : S Lip (aluminium alloy) S Forward Bulkhead S Anti−Ice and P20/T20 access panels S Outer Barrel (composite) S Inner Barrel (acoustic composite) S Aft Bulkhead (titanium) S Anti−ice ducting S Fan compartment ventilation inlet scoop S Interphone jack (only on Left Hand Side) S P20/T20 probe, pipe and harness S Fan Cowl Opening Switch Boxes (both sides)

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A380 30−20

Outer Barrel

Ventilation Scoop

P20/T20 Access Panel

P20/T20 Harness and Pipe Access Panel

Aft Bulkhead

Inner Barrel


Lip

Anti-Ice Access Panel

Left Fan Cowl Opening Switch Box

Interphone Jack

Nacelle Anti Ice Interface

Right Fan Cowl Opening Switch Box

Weight: 350 kg (770 lbs)

Figure 199 FRA US/T

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Air Intake Cowl

02 |30−20 |L3

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A380 30−20

Nacelle Anti−Ice System Hot engine air is provided through the anti−ice duct and through a tube, based on cyclone concept, inside the lip. The hot air flows inside the lip and is released overboard through the anti−ice access panel that is connected to the forward bulkhead. The anti−icing system can operate in all flight and ground conditions.


NAI Components The NAI components installed on the inlet cowl are as follows: S Nose cowl lip skin − defines the inlet cowl protected area S Forward bulkhead - ensures confinement of hot air in the forward part of the nose cowl S Supply duct assembly - Ensures routing of hot air from the EBU/Nose cowl interface to the nozzle S Cyclone Nozzle - delivers primary hot air massflow for anti−icing S Cyclone Mixer - ensures mixing of primary mass flow with recirculating air and avoids direct impingement of primary flow on nose cowl lip S Exhaust panel - discharges hot air overboard and ensures mixing with aerodynamic flow S Protection pipe - ensures containment of hot air that may leak from the supply duct and directs it to the intake lip where it is discharged overboard through the exhaust grid.

General Description The hot air from HP compressor via the NAI valves, is ducted through the feed pipe into the air intake lip. The feed pipe is located between the aft and the forward bulkheads in the space between the inner and outer barrel. A protection pipe is located around the feed pipe and gives protection in the event of duct rupture, so that it does not have an effect on the nose cowl structure. The cyclone system is located inside the intake lip and gives a swirling movement to the airflow inside the intake lip. The cyclone system consists of the following: S An injector with one centre nozzle and two lateral nozzles S A mixer S Spring brackets The mixer prevents direct impingement of the hot air onto the inlet, thus preventing overheating. It also causes a jet pump effect, which gives good hot air recirculation around the intake lip. The hot air circulates around the intake lip a few times before being discharged overboard through the exhaust grid.

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A380 30−20

Cyclone Nozzle Forward Bulkhead

Anti Ice Air Exit Grid (shown transparent) Anti Ice Access Panel (shown transparent)


Cyclone Tube

Aft Bulkhead Lip (shown transparent)

Figure 200 FRA US/T

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Anti Ice Duct

Nacelle Anti-Ice System 03 |30−20 |L3

Page 391

A380 30−20

NAI SYSTEM System Description The NAI system is controlled by one of the two Anti−Ice Control Units (AICU) on the aircraft, with inputs from the Anti−Ice switches on the overhead panel. When the system is operated in AUTO mode, the AICU uses inputs from aircraft mounted Ice detectors. In manual mode the AICU using the switch position inputs from the overhead panel switches. There are two nacelle anti−ice valves on each engine: S One shut−off valve (SOV) S One Pressure Regulating Anti−Ice Valve (RAIV) The shut−off valve is controlled by the AICU, via one of the solenoids in the right bleed valve controller in Zone 2. The valve is located in Zone 3 and is a pneumatically operated valve using HP3 muscle pressure from the bleed valve controller solenoid. Downstream of the SOV is a venturi restrictor which operates as a flow restrictor in case of a burst duct in the fan compartment (Zone 1), This will limit flow in these conditions, but will not affect normal operation. The pressure regulating anti−ice valve (RAIV) is located in Zone 1 and is pneumatically operated using downstream pressure. The valve regulates downstream pressure to avoid too high pressures on the lip skin. The valve incorporates two switches which both monitor the valve downstream pressure. The LP switch outputs to the AICU and is used to detect failures of the On/Off function. The HP switch outputs to the EEC, which sets a fault message if the downstream pressure is too high.

AICU Control Unit (AICU) Control The two aircraft mounted AICU s perform activation and deactivation of Wing Anti−Ice (WAI) and Nacelle Anti−Ice (NAI) systems based on inputs from the ice detectors and cockpit pushbutton switches. Each AICU is a dual channel controller, which controls the WAI & NAI systems. They also perform monitoring of the WAI & NAI system valves. The AICU s performs the following NAI functions: S AICU1 Channel A performs: − Engine 2 NAI control − Engine 4 NAI monitoring S AICU1 Channel B performs: − Engine 2 NAI monitoring − Engine 4 NAI control S AICU2 Channel A performs: − Engine 1 NAI control − Engine 3 NAI monitoring S AICU2 Channel B performs: − Engine 1 NAI monitoring − Engine 3 NAI control




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A380 30−20




Figure 201 FRA US/T

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NAI System Schematic 04 |30−20 |L3

Page 393



A380 30−20

NAI SHUT−OFF VALVE (SOV) Purpose To control the flow of HP3 air to the NAI system. The valve acts as an ON/OFF system Location The NAI SOV is located on the lower right side of the HP combustion outer case in Zone 3. Description The Shut−Off Valve is a solenoid operated and pneumatically actuated valve. The solenoid is contained in the right bleed valve controller in Zone 2 (covered in 75 − Air Systems). When the solenoid is energised, HP3 servo pressure from the solenoid valve acts on the piston in the SOV, and the valve is closed against the spring pressure. When the solenoid is de−energised, the HP3 servo pressure line to the valve is vented and the spring and HP3 air pressure from the compressor case, act on the piston and valve respectively to open the valve. NOTE:

The valve is sprung−loaded in the open position.


Manual Override If there are faults in the NAI system, the SOV may be locked in the open position. The manual override is unscrewed and locked in the fully open position. The manual override has a different colour to allow people to know if the valve is open or closed.

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A380 30−20




Figure 202 FRA US/T

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NAI Shut Off Valve

05 |30−20 |L3

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A380 30−20

ANTI−ICE PRESSURE REGULATING VALVE Purpose To regulate the pressure of the HP3 air to the inlet cowl. Location The NAI Pressure Regulating Valve is located on the lower left side of the fan case in Zone 1.


Description The Pressure Regulating Valve is a spring−loaded open, pneumatically actuated valve. When inlet pressure is applied to the valve, air flows past the open butterfly valve and provides a source of pressure at the downstream sensing port. This downstream pressure is used for: S Pressure regulation S Actuator positioning S Pressure switch operation The downstream pressure is regulated to provide an outlet pressure from the valve in the range of 65 - 83 psig. The LP switch operates in the range of 14 - 18 psig and is used to detect failures of the On/Off function. The LP switch outputs to the AICU The HP switch operates in the range of 85 - 98 psig and is used to detect failures of the valve regulation system. The HP switch outputs to the EEC, which sets a fault message if the downstream pressure is too high. Manual Override The valve includes a visual indicator and a manual lock arm that can lock the NAI pressure regulating valve in either the fully open or fully closed position. The override uses a manual lock bolt that serves a dual purpose. When seated in the storage position, the manual lock bolt holds a manual lock valve open in the downstream sense line, which allows the valve to function normally. On the removal of the manual lock bolt, the manual lock valve closes in the downstream sense line, blocking off downstream pressure and venting the regulator assembly and actuator to ambient. The manual lock bolt is retained by a lanyard and threads into the manual override to lock it in position.

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A380 30−20




Figure 203 FRA US/T

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NAI Pressure Regulating Valve 06 |30−20 |L3

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A380 30−20

MODE OF OPERATION AND COCKPIT INDICATIONS Manual Mode When the engine anti−ice is selected on the ENG Anti−Ice P/BSW‘s the ON legends on the switches illuminate and the MAN ENG A ICE message comes on as a MEMO item on the EWD. When manual mode is selected and icing conditions are detected by the system, a warning message will be displayed if the engine anti−ice is not selected ON. This will consist of a MASTER CAUT, single chime and message advising selection of ENG anti−ice ON. When icing conditions are no longer detected for more than 190 seconds and the ice protection systems are selected ON, the ICE NOT DET message illuminates as a MEMO on the EWD.


Indications ENG P/BSW‘s − ON (blue light) or FAULT (amber light). The fault light indicates a failure of one nacelle anti−ice system.

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Page 398

A380 30−20




Figure 204 FRA US/T

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NAI Operating Mode and Indication 07 |30−20 |L3

Page 399


POWER PLANT ENGINE GROUND OPERATION

A380

RR Trent 900

71

DANGER AREAS OF THE ENGINE WORKING AREA Engine Not Running Even if the engine is not running, the area is still dangerous and the personnel has to obey the precautions, which are given to operate an engine safely.

WARNING:

Engine Running To enable personnel safety when he has to act exceptionally on a running engine, the power level must be kept to the minimum necessary by setting throttle control levers to the IDLE position. The restricted areas are: S the intake suction area: in a radius of 4.5 m (15 ft), S the exhaust danger area: a corridor of 30_ from the exhaust nozzles to 70 m (230 ft) afterwards. To work on the engine safely, you must use the entry corridors located at the engine outboard side 1.3 m (4 ft) aft of the air intake cowl.

KEEP ALL PERSONS OUT OF THE DANGER AREAS DURING ENGINE OPERATION. CLEAN THE RAMP IF THERE IS SNOW, ICE, WATER, OIL OR OTHER CONTAMINATION OR MOVE THE AIRCRAFT TO A LOCATION THAT IS CLEAN. MAKE SURE THAT ALL PERSONS ARE SAFE BEFORE YOU START THE ENGINE. MAKE SURE THE PERSONS IN THE COCKPIT CAN SPEAK TO ALL PERSONS NEAR THE DANGER AREA DURING ENGINE OPERATION. OBEY ALL OF THE GROUND SAFETY PRECAUTIONS FOR THE ENGINES. THE ENGINES CAN PULL PERSONS OR UNWANTED MATERIALS INTO THEM AND CAUSE SERIOUS INJURIES OR DAMAGE TO EQUIPMENT

To work on the inboard engines, the outboard engines must be shut off first. Human factor points:


NOTE:

WARNING:

BE CAREFUL WHEN YOU DO WORK ON THE ENGINE PARTS AFTER THE ENGINE IS SHUTDOWN. THE ENGINE PARTS CAN STAY HOT FOR ALMOST 1 HOUR.

WARNING:

UNDER NORMAL CONDITIONS, EXCEPT IN THE ASSISTED MANUAL START SEQUENCE, THERE IS NO NEED AND IT IS NOT ALLOWED TO PERFORM MAINTENANCE TASKS ON A RUNNING ENGINE.

WARNING:

DO NOT GO NEAR AN ENGINE THAT IS IN OPERATION ABOVE LOW IDLE. IF YOU DO, IT CAN CAUSE AN INJURY. GO NEAR AN ENGINE IN OPERATION THROUGH THE ENTRY CORRIDORS ONLY.

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A380

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71

8.9 m (29 ft)

4,5 m (15 ft) 1,3 m (4 ft 3 in)


70 m (230 ft)

30 TO 548.6 m (1800 ft) AFT OF EXHAUST NOZZLES

30 ° INTAKE SUCTION DANGER AREA MINIMUM IDLEWPOWER EXHAUST DANGER AREA

INTAKE SUCTION DANGER AREA MAX TAKE−OFF POWER

ENTRY CORRIDOR

EXHAUST DANGER AREA

Figure 205 FRA US/T

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Engine Danger Areas 01|71|L3

Page 401



A380 RR Trent 900

71

POWER PLANT GROUND OPERATION PREPARE THE AIRCRAFT / ENGINE FOR GROUND OPERATION Task 71−00−00−860−803−A Obey the instructions that follow: WARNING:

Make sure that the aircraft brakes are on. Make sure that the CHOCK − WHEEL are in position in front of each forward wheel of the WLG (Wing Landing Gear).

YOU MUST OBEY THE PRECAUTIONS THAT ARE GIVEN FOR PERSONS TO OPERATE AN ENGINE SAFELY. IF YOU DO NOT, AN INJURY AND/OR DAMAGE CAN OCCUR.

The ground operation time for the engine. Do not start the engine, unless it is necessary. You must keep the ground operation time of the engine to a minimum. When you operate the engine, change the engine speed slowly. NOTE:

Movements of the throttle lever (to increase or decrease the engine speed) must be smooth and slow. Unless the instruction is different, the controlled movement must take at least 30 seconds.When you operate the engine on the ground, keep the power level and the time of the operation to the minimum that is necessary.

When you must do more than one task that makes it necessary to operate the engine, try to do the tasks at the same time. The safety precautions for the ground operation of an engine. Make sure that the aircraft is pointed into the wind. Make sure that the ground surface in the engine ground operations area is not broken or loose and is clear of unwanted materials. FOR TRAINING PURPOSES ONLY!

NOTE:

The ground operation area is 12.19 m (40 ft.) each side of the engine center line and 18.28 m (60 ft.)forward from the rear of the engine. Make sure that the aircraft is clear of all structures and other aircraft. Make sure the engine exhaust danger area for all engines is clear. NOTE:

NOTE:

FRA US/T

If a blast fence is necessary, it is recommended that the engine nozzle is positioned at least 60.96 m (200 ft.) from it. If the engine is operated less than 60.96 m (200 ft.) from the blast fence, it is possible that the engine will not be stable.

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Sep 10, 2008

If engine start is followed by a high power run, use CHOCK−WHEEL, ENGINE RUN UP (98L10001005000). Make sure that there are no COVER − PROTECTION on the engine. Make sure that there are no unwanted objects in the engine inlet and exhaust. Make sure that the engine inlet and exhaust danger areas are clear of persons and ground support equipment. Make sure that unwanted persons and unwanted vehicles cannot easily enter the danger areas. Make sure that persons with loose clothing do not go near the engine. Make sure that the ground fire extinguisher equipment is in its position with the applicable persons. Make sure that the fan cowl panels are closed before you operate the engine. NOTE:

The fan cowl panels can be open for specified tests, for example leak tests. The applicable procedure will tell you when to keep the fan cowl panels open. The entry corridors. NOTE:

YOU MUST NOT GO NEAR AN ENGINE THAT IS IN OPERATION ABOVE LOW IDLE. IF YOU DO, IT CAN CAUSE AN INJURY. WHEN AN ENGINE IS IN OPERATION AT LOW IDLE, YOU CAN ONLY GO NEAR IT THROUGH THE ENTRY CORRIDORS. Make sure that you use the entry corridors to go near an engine that is in operation at low IDLE (forward thrust only). Make sure that you do not operate the engine above low IDLE with persons in the entry corridor. WARNING:

NOTE:

02|71/L3

If the engine is in operation at more than low IDLE, refer to the take−off power danger areas or to the breakaway power danger areas.

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A380 RR Trent 900

71

Ear protection. WARNING: WARNING: YOU MUST USE EAR PROTECTION WHEN YOU ARE NEAR AN ENGINE THAT IS IN OPERATION. THE NOISE MADE BY AN ENGINE IN OPERATION CAN CAUSE PERMANENT DAMAGE TO YOUR EARS. Make sure that you use the correct ear protection when you are near an engine that is in operation. The engine operation. Do not operate the engine on the ground, at more than the engine limits Engine power control. Keep the power level and the time of the operation at the minimum that is necessary. Make sure that you move the throttle levers slowly unless, because of the procedure, it is necessary to move them differently.


NOTE:

Fast movement of the throttle levers can cause the engine temperature to change quickly. This will decrease the life of the engine.

Crosswind conditions and engine surges. Make sure that you obey the wind direction and velocity limits for engine operation. Bad wind conditions (turbulence, gusty, crosswind) while you operate the engine at middle power and above can cause: S The engine parameters (TPR (Turbine Power Ratio), EGT (Exhaust Gas Temperature), RPM (Revolution Per Minute)) to increase or decrease and not stay constant S A roar from the air intake to be heard. If you hear the intake roar you must immediately decrease the engine power. It is not permitted that you run the engine in these conditions. You can identify an engine surge by an increase in EGT and a sudden increase in noise from the engine.

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If an engine surge occurs you must: S Immediately decrease the engine power until the engine surge stops. S Make sure that the EGT decreases. S Let the engine become stable at low idle. S Slowly increase the engine power. S Look at the engine parameters to see if the engine has a surge again. S If the surge does not occur again, continue with the engine test procedure. S If the engine has a surge again, because of bad wind conditions, stop the engine procedure. It is not necessary to examine the engine for this type of surge. S If the engine has a surge again, not caused by bad wind conditions, immediately decrease power to low idle. Let the engine cool for 5 minutes, then stop the engine. Find the cause of the surge. Static electricity discharge from the LP compressor spinner. If you operate the engine on the ground in a low−humidity condition, you can see sparks around the LP compressor spinner. This will not cause damage to the engine.

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71

The engine anti−ice. You must use the engine anti−ice if the conditions that follow occur: If the OAT (Outside Air Temperature) is less than 6 deg.C (43 deg.F) and moisture (fog, rain, snow, sleet or hail) is seen. Fog is specified as visibility lower than 300 m (984 ft.) due to moisture. If it is necessary to use the engine anti−ice, set the applicable ENG ANTI ICE switch to ON immediately after the engine gets to low idle. Do not do the performance test procedures when the ENG ANTI ICE switch is set to ON. NOTE:

The performance limits are given with the air off−takes set to OFF. Do the steps that follow if you must do the performance test when conditions make the use of the anti−ice system necessary. Make sure that there is no ice on the air intake cowl before you set the ENG ANTI ICE switch to OFF. When you must make a record of the engine indications, set the ENG ANTI ICE switch to OFF for a maximum of 60 seconds. Immediately after you make a record of the engine indications, set the ENG ANTI ICE switch to ON. Make sure that there is no ice on the air intake cowl before you set the ENG ANTI ICE switch to OFF again. NOTE:


CAUTION:

FRA US/T

IF YOU OPERATE THE ENGINE IN FREEZING FOG, YOU MUST DE−ICE THE ENGINE CORE INLET REGULARLY. THE MAXIMUM PERMITTED TOTAL TIME OF OPERATION IN THESE CONDITIONS IS 60 MINUTES (UNLESS THE ENGINE CORE INLET IS DE−ICED DURING THAT PERIOD). IF YOU OPERATE THE ENGINE FOR LONGER, TOO MUCH ICE CAN BUILD−UP ON THE CORE INLET COMPONENTS. THE SUBSEQUENT RELEASE OF THIS ICE, AT HIGHER POWER, CAN CAUSE DAMAGE TO THE COMPRESSOR.

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The engine start. The engine start is satisfactory if these conditions occur: On the panel 1125VU, the engine starts in 30 seconds or less after the ENG MASTER control switch is set to ON. The engine speed increases smoothly and continuously to low idle.The EGT stays in the limits. The engine start is unsatisfactory if these conditions occur: Hot start or impending hot start S A start when the EGT goes near or higher than the start limit. Hung start S The engine light−up is satisfactory but it does not accelerate correctly (speed increases slowly or decreases) and the EGT goes near to its limit. Aborted start S The start procedure is stopped before the start is completed. A cold weather condition start. CAUTION:

At ambient temperatures below −40 deg.C (−40 deg.F), do a check of the vibration indication on the ECAM screen. If the vibration indication is not seen on the ECAM screen, do the steps that follow: Warm the EMU (Engine Monitoring Unit). To warm the EMU, on the panel 1215VM, select the applicable ENG START selector switch to IGN START.

IF THE ENGINE IS IN A COLD ENVIRONMENT, THE ENGINE OIL CAN BECOME TOO COLD. IF THE ENGINE IS NOT OPERATED, AND IS IN THIS ENVIRONMENT, YOU MUST DO A CHECK OF THE ENGINE OIL TEMPERATURE REGULARLY. IF NECESSARY, REGULARLY START THE ENGINE AND OPERATE IT AT IDLE TO KEEP THE OIL TEMPERATURE ABOVE MINUS 10 DEG.C (14 DEG.F). Make sure that the oil temperature is more than −10 deg.C (14 deg.F) before you start the engine. To find the engine oil temperature: S Do the steps necessary to set the ECAM (Electronic Centralized Aircraft Monitoring) screen to the engine page S Look at the ECAM display screen and get the oil temperature value from the applicable SD (System Display).

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When the EMU is warm, a vibration indication will be seen on the ECAM screen. It can be 15 minutes before the EMU is warm. When the vibration indication is seen on the ECAM screen, on the ENG START section of the panel 1215VM, select the applicable ENG START selector switch to NORM. Do the applicable autostart or manual start, immediately. When you start a cold soaked engine these conditions can occur: S The engine oil pressure can be more than 100 psi (6.89 bar). S The indication for oil quantity can decrease. When the engine becomes stable at the low idle condition: S The oil temperature will rise. S The oil pressure will decrease. S The indication for the oil quantity will become normal. Do not operate the engine above idle until the oil quantity indication is satisfactory. NOTE:

YOU MUST NOT START, DRY MOTOR OR WET MOTOR THE ENGINE IF THE OIL TEMPERATURE IS LESS THAN MINUS 10 DEG.C (14 DEG.F). LOW OIL TEMPERATURES CAN CAUSE DAMAGE TO THE ENGINE BEARINGS.

CAUTION:


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NOTE:

If the start procedure takes a long time to complete, the indication for the oil pressure can decrease. You can also have a warning for low oil pressure. These conditions are permitted if the oil system parameters go to their normal values when the engine becomes stable at low idle.

NOTE:

If you stop the start because of an indication of low oil quantity and LP warning, you can start the engine again. But do not add oil to the oil tank.

NOTE:

If the start is satisfactory, make sure that the oil parameters return to the normal limits when the engine is at low idle.

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ENGINE OPERATION LIMITS Exhaust Gas Temperature (EGT) limits. Ground start: S 700 deg.C (1292 deg.F) at less than 50 percent N3. S In−flight relight: 850 deg.C (1562 deg.F). Maximum continuous: 850 deg.C (1562 deg.F). Takeoff: 900 deg.C (1652 deg.F) for a maximum of 5 minutes, or 10 minutes in the event of an engine failure.

Rotor operation speed limits. The maximum N1 is 96.1 percent rpm (100 percent = 2900 rpm). The maximum N2 is 97.8 percent rpm (100 percent = 8300 rpm). The maximum N3 is 97.8 percent rpm (100 percent = 12200 rpm). NOTE:

You must tell Rolls−Royce plc if an overspeed condition occurs. Give Rolls−Royce the exceedance data from the OMT and the engine standard from the data plate.

Oil pressure limits. The minimum oil pressure with N3 at or above idle is 25 psi (1.7 bar) differential pressure.

NOTE:

Engine speed can be reduced by the EEC (Engine Electronic Controller) to less than the specified limits above.

NOTE:

An engine oil pressure indication is shown in the cockpit. This is an adjusted indication if the engine RPM is above 70% N3. This is so that the in−flight LP advisory message and the low oil pressure warning are given at the same time. The low oil pressure warning is given if engine oil pressure is 25 psi (1.724 bar) or less.

Oil temperature limits. YOU MUST NOT START, DRY MOTOR OR WET MOTOR THE ENGINE IF THE OIL TEMPERATURE IS LESS THAN MINUS 10 DEG.C (14 DEG.F). LOW OIL TEMPERATURES CAN CAUSE DAMAGE TO THE ENGINE BEARINGS. At start, the oil temperature must be more than a minimum of −10 deg.C (14 deg.F). Before acceleration to take−off, the oil temperature must be 60 deg.C (140 deg.F) or more. The maximum oil temperature is 196 deg.C (384.8 deg.F) when the operation condition is stable.


CAUTION:

Oil consumption limits. The maximum oil consumption is 0.45 l/hr (0.48 USQT/hr). The usual engine oil consumption is 0.095 l/hr (0.1 USQT/hr). If the engine has an increase in oil consumption above 0.14 l/hr (0.15 USQT/hr), do the high oil consumption troubleshooting FRA US/T

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Static engine operation. Stable operation in the speed range 64 to 72 percent N1 or above 78 percent N1 is not permitted during static ground operations. But temporary operation through the speed range 64 to 72 percent N1 is permitted while the thrust increases or decreases. The EEC automatically prevents operation in these speed ranges in primary and rated reversionary thrust control modes. Starter motoring limits. Continuous motoring: S The maximum time is five minutes. S After continuous motoring of five minutes the starter must be cooled for 30 minutes, before the starter is motored again. S Intermittent motoring. The total motoring time permitted is five minutes in any 35 minutes time period. Vibration Advisory Limits. Cockpit alert level: 1 LP band 5 units (1.00 in./second peak velocity). 2 IP band 5 units (0.9 in./second peak velocity). 3 HP band 5 units (0.7 in./second peak velocity).

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71 60°

60°


60°

The wind velocity is for stable wind conditions. Decrease the maximum wind limit 5 knots for gusty wind conditions.

RELATIVE WIND GROUND OPERATIONS UP TO LOW IDLE NO LIMIT PERMITTED 35 KNOTS MAXIMUM WIND VELOCITY

PERMITTED 10 KNOTS MAXIMUM WIND VELOCITY

PERMITTED 20 KNOTS MAXIMUM WIND VELOCITY

PERMITTED 5 KNOTS MAXIMUM WIND VELOCITY (limited up to high idle)

Figure 206 FRA US/T

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GROUND OPERATIONS UP TO THE MAXIMUM N1 GROUND LIMIT (ref TASK 71−00−00−860−810) RECOMMENDED 30 KNOTS MAXIMUM WIND VELOCITY

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ENGINE START ASSISTANCE OPS/CTL & IND General This is a general view of the A380 cockpit.




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Figure 207 FRA US/T

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Engine Start Controls & Panels Location On the overhead panel, the engine controls are: S ENGine FADEC GrouND PoWeR panel, S ENG START mode selector, S ENG MANual START panel, On the center pedestal, the engine start controls are: S ENG MASTER levers, S ENG Throttle Control Levers.




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Figure 208 FRA US/T

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ENGINE START ASSISTANCE DESCRIPTION Ignition System Architecture Each engine has two independent ignition systems, which give an electrical spark used to start ignition of the fuel/air mixture in the engine. The engine starting, ignition controls are found on the following cockpit panels: S ENGine START control panel, on the overhead control panel, S ENG MASTER control panel, on the center pedestal, S ENG MANual START panel, on the overhead control panel. All these controls are linked to the Input/Output Modules (IOMs) type A. The IOMs are themselves linked to the EEC (Engine Electronic Controller) via the ADCN (Avionics Data Communication Network). This enables the EEC to control the engine starting sequences, engine cranking options and the ignition selection in response to aircraft command signals. The ENG MASTER levers are hardwired to the EEC for reset and back−up purposes if the ADCN fails. The EEC also interfaces with the ignition units and the Starter Control Valves, in order to control and monitor their operation during the starting or cranking phases. The EIPM (Engine Interface Power Management) maintains the power supply to the EEC and supplies the ignition system with 115 VAC. The ignition leads transmit the electrical power from the ignition units to the igniter plugs. Throttle control lever sends an analogic signal to the EEC to enable it to compute the correct thrust to be applied. Automatic Start The EEC does the selection of an automatic engine start after reception of the appropriate cockpit commands. The EEC will automatically shut down the engine if the start procedure is not satisfactory. For automatic start of the engine, the controls have to be selected as follows: S the MAN START P/BSW on the ENG panel is off (ON legend is off), S the thrust lever is set to the IDLE position, S the rotary selector on the ENG START panel is set to the IGNition/START position, S the lever on the ENG MASTER panel is set to the ON position. S once the engine is running the ENG START rotary selector is set to NORM position. FRA US/T

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Manual Start Alternatively the engine can be started manually with the flight crew or maintenance personnel in control of the start sequence. In this mode the engine starting control is under limited authority of the EEC. After reception of the appropriate cockpit commands the EEC system has a limited interaction to control the starter control valve, fuel and igniters. For manual ground start of the engine, the controls have to be selected as follows: S the thrust lever is set to the IDLE position, S the rotary selector on the ENG START panel is set to the IGN/START position, S the MAN START P/BSW on the ENG panel is set to ON, S the applicable N2 (intermediate pressure shaft) and N3 (HP shaft) rotation speeds are monitored on the ENGINE page of the ECAM SD and the procedure continues when they are reached, S the lever on the ENG MASTER panel is set to the ON position at 25% N3 and EGT (Exhaust Gas Temperature) less than 150_C (302_F), S the normal EGT rise is checked on the E/WD by means of an indication light−up, S the applicable N1 (LP compressor rotation speed) is monitored on the EWD and the procedure continues when it is reached, S once the engine is running MAN START P/BSW goes off, S the ENG START rotary selector is set to NORM position. WARNING:

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Figure 209 FRA US/T

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Engine Start After the preliminary cockpit preparation has been done, the engine start can be initiated. Normal Start Procedure In normal start procedure, you start engine in automatic mode. To start the engine in normal procedure: S Turn the ENG START selector on ”IGNition START” position, the ECAM ENG page appears, S Check if all indications are normal and all parameters for logical indications, S Ask the ground clearance to start the engine, S If you obtain this clearance, announce the start engine, S Check the air pressure is above 30 psi on the ECAM, S Set the ENG MASTER lever to ON, S On the ECAM, check that the start valve is in line and the start valve air pressure is upper 30 psi. The N3 and the oil pressure increase, When N3 is equal to 25%, the active igniter (A or B) is indicated and the fuel flows. S Start the chronometer and within 30 seconds, the EGT increases, S At N3 above 30%, N1 increases and the oil pressure is green, S At N3 above 50%, the start valve is cross line and the igniter indication disappears, S When the engine is at idle, check parameters for logical indication. Engine Start Valve Fault If a start valve fault is detected during engine start in automatic mode: S a single chime sounds, S the MASTER CAUTion lights come on, S the ENG 2 START VALVE FAULT message is displayed on E/WD, S the ENG START VALVE FAILED CLOSED caution is displayed on ECAM ENG page. With the failure of the engine start in automatic mode, you can perform the engine start in manual mode with ground mechanic assistance.

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Figure 210 FRA US/T

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Engine Start with Assistance In case of the engine start failure in automatic mode, you will have to perform a specific procedure to start the engine in manual mode with ground mechanic assistance: Start Procedure Before you start the engine, check the cockpit configuration: S the ENG 1 MASTER LEVER is OFF, S the ENG 2 MASTER LEVER is OFF, S the ENG 3 MASTER LEVER is OFF, S the ENG 4 MASTER LEVER is OFF, S the ENG START selector switch is on NORMAL position, S the ENG 1 MAN START is NORMAL, S the ENG 2 MAN START is NORMAL, S the ENG 3 MAN START is NORMAL, S the ENG 4 MAN START is NORMAL, S the ALTerNate MODE is NORMAL, S the ENG 1 FIRE is NORMAL, S the ENG 2 FIRE is NORMAL, S the ENG 3 FIRE is NORMAL, S the ENG 4 FIRE is NORMAL, S the ENG 1 THROTTLE CONTROL LEVER is on FWD IDLE position, S the ENG 2 THROTTLE CONTROL LEVER is on FWD IDLE position, S the ENG 3 THROTTLE CONTROL LEVER is on FWD IDLE position, S the ENG 4 THROTTLE CONTROL LEVER is on FWD IDLE position, S the ENG 2 THRUST REVERSER LEVER is on STOW POSITION, S the ENG 3 THRUST REVERSER LEVER is on STOW POSITION, S ECAM−E/WD & SD are powered, S the FUEL PUMPS are ON, S the FLAPS LEVER POSITION is at 0 position, S the PARKING BRAKE is ON, Inform the GROUND CREW before you start the engine.

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Radio Management Panel Use flight or service interphone to establish the contact with the ground personnel. The RMPs (Radio Management Panels) located in the center pedestal panel give the controls to manage this communication. With the CALLS/MECH P/BSW on the overhead panel you can make a call to the ground mechanic and you can validate a call from the ground mechanic by pressing the MECH P/BSW on the RMP (cf ATA23).

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Figure 211 FRA US/T

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Starter Control Valve Override The starter control valve is installed in the starter duct at the lower left side of the fan case. Normally controlled by the EEC, the starter control valve controls the flow of air to the pneumatic starter. When the starter control valve fails to open, the START VLV FAULT message is shown on the EWD. The starter control valve can be manually open for dispatch reasons without opening the fan cowl door. Access is available through a spring−loaded flap in the fan cowl door. The manual override of the starter control valve shall be possible by applying a DRIVE 3/8−inch − SQUARE to the square socket installed on the valve. During this operation the maintenance personnel has to stay in contact with the cockpit through the service interphone whose connection is located at the air intake cowl. Only the cockpit crew orders to open the valve, the tool has to be rotated counter clockwise until the end stop is reached. The valve is kept in this position until the cockpit crew order to close it. This order is given after 50% of N3 is reached. Rotating the tool clockwise until the end stop is reached closes the starter control valve. WARNING:

BE CAREFUL WHEN WORKING ON A HOT RUNNING ENGINE,

WARNING:

USE SPECIFIC EQUIPMENTS (ACCESS PLATFORM, GLOVES, HEADSET...).




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MAINTENANCE INTERPHONE JACK

ECAM/EWD MEMO ZONE MANUAL OVERRIDE LEFT FAN COWL UP/DOWN SWITCH FAN COWL DOOR

WRENCH

DRIVE 3/8 IN SQUARE

ACTUATOR ASSEMBLY



SQUARE SOCKET (MANUAL OVERRIDE) SPRING LOADED FLAP

Figure 212 FRA US/T

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STARTER CONTROL VALVE

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Engine Manual Start The procedure to start the engine in manual mode is to: S Set the ENG START selector to the ”IGN/START” position. The ECAM ENGINE page is displayed. S Check if all indications are normal and all parameters for logical indications, S Ask the ground clearance to start the engine, S If the GROUND CLEARANCE is OBTAIN, announce the engine 2 start, S Check the air pressure is above 30 psi on the ECAM ENGINE page, S Set the ENG MAN START 2 to the ”ON” position. S Contact the ground crew to turn the start valve in OPEN position. S On the ECAM, check that the start valve is in line and the start valve air pressure is upper 30 psi. The N3 and the oil pressure increase, S When N3 reaches 20%, set the ENG MASTER lever to the ”ON” position. On ECAM ENGINE page the active igniter (A or B) and the fuel flow are indicated. S Start the chronometer and within 30 seconds, the EGT increases, S At N3 above 30%, N1 increases and the oil pressure is green. S At N3 above 50%, contact the ground mechanic crew to close the start valve, the start valve is in cross line and the igniter indication disappears. S Set the ENG MAN START to the ”OFF” position S When the engine is at idle, check parameters for logical indication.




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Figure 213 FRA US/T

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Figure 214 FRA US/T

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IGNITION AND STARTING SYSTEM DESCRIPTION The ignition and starting system has three subsystems: S Starting, S Fuel command, S Ignition. Starting Engines can be started using the APU air bleed, a ground air supply or crossbleed air from an operating engine. The EEC (Engine Electronic Controller) controls the opening and closing of the SCV (Starter Control Valve) in all start modes. The SCV controls the air flow to the pneumatic starter. The Pneumatic starter drives N3 through the accessory gearbox. The starter has three different cycles: S Normal cycle runs: − Up to 2 minutes continuous operation then runs down to zero N3, − Up to 2 minutes continuous operation then runs down to zero N3, − Up to 1 minute continuous operation then runs down to zero N3 and wait 30 minutes for the cooling. S Extended start cycle: − Up to 5 minutes continuous operation followed by 30 minutes wait for the cooling. S Extended crank cycle: − Up to 5 minutes continuous operation followed by 30 minutes wait for the cooling.

Ignition units power supply The EIPM (Engine Interface Power Management) supplies ignition units (A and B) through the EEC (channel A and B) control. A/C EMERgency BUS BAR 115 VAC supplies EIPM CHAN A function. EIPM CHAN A function could supply Ignition unit A or B depending on the EEC switching. A/C BUS 2 BAR 115 VAC supplies EIPM CHAN B function. EIPM CHAN B function could supply Ignition unit A or B depending on the EEC switching. EEC ignition units switching function: Each EEC channel is able to control the switching of the power supply of the two ignition units A and B. During an engine auto start on ground, the EEC controls automatically the switching of the ignition units A or B. During an engine manual start, the EEC controls both ignition units A and B, for ignition efficiency. NOTE:

During Engine auto start in flight, both ignition units are energized, for redundancy.

Fuel command The EEC controls, through the MV (Metering Valve) servovalve, the FMV (Fuel Metering Valve) which regulates the fuel flow to the manifolds. The ENGine MASTER lever controls, through the airframe shutdown solenoid, the closing of the HP SOV (High Pressure Shut−OFF Valve). The HPSOV is also called the MP SOV (Minimum Pressure Shut−Off Valve). EEC controls through the protection motor the closing of the HPSOV. The Fuel Flow Transmitter (FF XMTR) sends his data to the EEC.

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Figure 215 FRA US/T

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Controls from the cockpit The engine start/crank is controlled from the cockpit by: S ENGine rotary selector, S ENGine MASTER levers, S ENGine MANual START P/B SW, S TCA (Throttle Control Assembly).


System functions The engine can be started in two manners: S Automatic start (normal procedure), S Manual start (back−up procedure). The engine can be cranked in two manners: S Dry crank, S Wet crank. Continuous relight function: S Manually selected with ENG START rotary selector to the ”IGN/START” position. Auto−relight function: S When a flame−out is detected the system energizes the two igniters. Quick relight function (in flight only): S When the ENG MASTER lever is inadvertently selected ”OFF”, the ENG MASTER lever can be selected ”ON” again within 30 seconds to cancel the shutdown sequence.

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Figure 216 FRA US/T

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ENGINE START / CRANK CONTROL DESCRIPTION Instructions and Precautions for Engine Ground Operation WARNING:

YOU MUST NOT GO NEAR AN ENGINE THAT IS IN OPERATION ABOVE MINIMUM IDLE. IF YOU DO, IT CAN CAUSE AN INJURY. WHEN AN ENGINE IS IN OPERATION AT MINIMUM IDLE, YOU CAN ONLY GO NEAR IT THROUGH THE ENTRY CORRIDORS.

WARNING:

YOU MUST MAKE SURE THAT ALL AREAS WHERE YOU OPERATE THE ENGINE ARE AS CLEAN AS POSSIBLE. ALL AREAS MUST BE VERY CLEAN TO PREVENT INJURY AND SERIOUS DAMAGE TO THE ENGINE AND AIRCRAFT.

WARNING:


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BEFORE YOU OPERATE THE ENGINES AT POWER SETTINGS ABOVE IDLE, MAKE SURE THAT THERE IS NO RISK OF PRE−PRESSURIZATION OR RESIDUAL PRESSURE IN THE AIRCRAFT AFTER SUBSEQUENT ENGINE SHUTDOWN. TO DO THIS, MAKE SURE THAT THE AIR CONDITIONING OUTFLOW VALVES ARE OPEN DURING THE ENTIRE TEST.

WARNING:

IF PERSONS TRY TO OPEN A DOOR WHEN THERE IS RESIDUAL PRESSURE IN THE AIRCRAFT: −THE DOOR CAN OPEN WITH DANGEROUS SUDDEN FORCE, −THERE IS A RISK OF BAD INJURY OR DEATH, AND −THERE CAN BE DAMAGE TO THE AIRCRAFT.

WARNING:

MAKE SURE THAT THE TRAVEL RANGES OF THE FLIGHT CONTROL SURFACES ARE CLEAR BEFORE YOU MOTOR THE ENGINE. MOVEMENT OF THE FLIGHT CONTROL SURFACES CAN BE DANGEROUS AND/OR CAUSE DAMAGE.

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WARNING:

TO ABORT THE ENGINE START SEQUENCE, YOU MUST PUT THE ENG/MASTER SWITCH BACK TO THE OFF POSITION. IF YOU ONLY CHANGE THE POSITION OF THE ENGINE MODE ROTARY SELECTOR SWITCH (FROM IGN/START TO NORM), THE FADEC SYSTEM WILL NOT ABORT THE START SEQUENCE. THIS CAUSES A RISK THAT THE ENGINE WILL CONTINUE TO START. THIS, IN TURN, CAN CAUSE INJURIES TO PERSONNEL.

CAUTION:

YOU MUST NO OPERATE THE ENGINE IF THE FAN EXHAUST COWLS ARE OPEN. IF THE ENGINE IS OPERATED WHEN THE FAN EXHAUST COWLS ARE OPEN, DAMAGE TO THE POWER PLANT CAN OCCUR.

CAUTION:

YOU MUST NOT START, DRY MOTOR OR WET MOTOR THE ENGINE IF THE OIL TEMPERATURE IS TOO COLD (REFER TO AMM). LOW OIL TEMPERATURES CAN CAUSE DAMAGE TO THE ENGINE BEARINGS.

CAUTION:

IF THE ENGINE IS IN COLD ENVIRONMENT, THE ENGINE OIL CAN BECOME TOO COLD. IF THE ENGINE IS NOT OPERATED, AND IS IN THIS ENVIRONMENT, YOU MUST DO A CHECK OF THE ENGINE OIL TEMPERATURE REGULARY. IF NECESSARY, DO AN ENGINE START AND OPERATE THE ENGINE AT IDLE UNTIL THE ENGINE OIL TEMPERATURE IS SATISFACTORY.

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Figure 217 FRA US/T

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AUTOMATIC START General The automatic start sequence could be automatically or manually aborted. Automatic Start on Ground The procedure to start the engine in automatic mode Initial configuration of controls (engine not running) is: The ENG MASTER lever is set to the ”OFF” position and the ENG START rotary selector is set to the ”NORM” position. Set the ENG START rotary selector to the ”IGN/START” position. The ECAM ENGINE page is displayed and the AGU (Air Generation Unit) flow control valves close. Set the ENG MASTER LEVER to the ”ON” position. The LP (Low Pressure) fuel valve opens and the SCV opens. When N3 reaches 25% and the EGT (Exhaust Gas Temperature) is below 150_C, the following events occur: S The ignition is in function on the igniter A or B, S The HP SOV opens, S The FF increases, S The EGT increases.


NOTE: The maximum EGT during a start is 700_C. When N3 reaches 50%, the SCV closes, the igniters cuts off and the AGU flow control valves reopen if there is no other engine in starting sequence.

Automatic Start Abort Automatic start abort is initiated when the following troubles occur: S Hot start / stall S Hung start, S No light up, S Locked N1 rotor, S SCV failed closed, S High N3. If there is a default, the HP SOV automatically closes and the ignition stops. If there is hot start /stall, hung start or no light up, the EEC automatically initiates a shutdown followed by a dry cranking period (to reduce EGT below 150_C, only in hot start configuration) and then the EEC tries a new start. If there is a N1 rotor locked, a SCV failed closed or too high N3, the engine start is automatically aborted. Automatic Start Manual Abort If a default occurs you can at anytime SET the ENG MASTER lever to the ”OFF” position. ENG MASTER lever set to the ”OFF” position has priority over the automatic mode. At this time the HP and LP SOVs closes, the ignition stops, the SCV closes and the EEC is reset. To restart the engine, proceed to another automatic start.

NOTE: The maximum of EGT during a start is 700_C. Set the ENG START rotary selector to the ”NORM” position. The ECAM ENGINE page disappears. If after engine start, the rotary selector is set to NORM and back to IGN/START, continuous relight is activated on the running engine(s).

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Figure 218 FRA US/T

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MANUAL START General The manual start sequence could be aborted only manually Manual Start on Ground The procedure to start the engine in manual mode. Initial configuration of controls (engine not running) is: The ENG MASTER lever is set to the ”OFF” position, the ENG START rotary selector is set to the ”NORM” position and the ENG MAN START P/B SW ”ON” legend lights off. Set the ENG START rotary selector to the ”IGN/START” position. The ECAM ENGINE page is displayed and AGU flow control valve closes. Set the ENG MAN START to the ”ON” position. The SCV opens. When N3 reaches 25% and the EGT (Exhaust Gas Temperature) is below 150_C, set the ENG MASTER lever to the ”ON” position. The following events occur: S The ignition starts on the igniter A and B, S The LP fuel valve and the HP SOV open, S The FF increases, Start the chronometer: S Within 30 seconds, the EGT increases.

Manual Start Interruption Before to set the ENG MASTER lever to the ”ON” position, you can interrupt the start sequence by setting the ENG MAN START P/B SW to the OFF position. This action causes the closing of SCV. If, a SCV failed closed, a locked N1 or a too high N3 is detected you must set the ENG MAN START P/B SW to the ”OFF” position. After you set the ENG MASTER lever to the ’ON” position, you can interrupt the engine starting by setting the ENG MASTER lever to the ”OFF” position. Following to this action, the LP fuel valve and the HP SOV close, the ignition stops, the SCV closes, and the EEC is reset. The ENG MASTER lever must be set to the ”OFF” position, when the following troubles occurs: S Hot start / stall S Hung start, S No light up, S Locked N1 rotor, S SCV failed closed, S High N3. Before attempting another start, dry crank the engine for 30 seconds at least (2 minutes max) to reduce EGT below 150_C.

NOTE: The maximum EGT during a start is 700_C. When N3 reaches 48%: S The SCV closes. The ENG MAN START pushbutton has to be set to the ”OFF” position only for confirmation. S Ignition stops. When N3 reaches 50%, set the ENG START rotary selector to the ”NORM” position, this actions occurs: S AGU flow control valves reopen if there is no other engine in starting sequence, S The ECAM ENGINE page disappears. If after engine start, the rotary selector is set to NORM and back to IGN/START, continuous relight is activated (on the running engine)

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Figure 219 FRA US/T

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CRANKING General Dry crank or wet crank can be done. Dry Crank Dry crank is used to remove any residual fuel from the combustion chamber and to check if there is not oil leak. During Initial configuration of controls (engine not running): S Open the following C/Bs: − e.g. for Engeine 1 HP FUEL SOV ENG (1KC1) S The ENG MASTER lever is set to the ”OFF” position, S the ENG START rotary selector is set to the ”NORM” position S and the ENG MAN START P/B SW ”ON” legend lights off. S Set the ENG START rotary selector to the ”CRANK” position. − The ECAM ENGINE page is displayed. S Set the ENG MAN START to the ”ON” position. − The SCV opens. S Start the chronometer. NOTE: Dry crank the engine from 30 seconds untill 2 minutes maximum. S Set the ENG MAN START to the ”OFF” position, − SCV closes, S Stop the chronometer. S Set the ENG START rotary selector to the ”NORM” position, − the ECAM ENGINE page disappears.

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Wet Crank Wet crank is used to check if there is not fuel leaks. The Initial configuration of controls (engine is not running) is: S The ENG MASTER LEVER set to the ”OFF” position, S the ENG START rotary selector set to the ”NORM” position S and the ENG MAN START P/B SW ”ON” legend lights off. NOTE: obey the starter limitation (normal cycle : 2 minutes maximum) S Set the ENG START rotary selector to the ”CRANK” position, − the ECAM ENGINE page is displayed. S Set the ENG MAN START to the ”ON” position, − the SCV opens. When N3 reaches 33%, S set the ENG MASTER lever to the ”ON” position, − LP fuel valve and the HP SOV open, S then start the chronometer. After 30 seconds set the ENG MASTER lever to the ”OFF” position, and the following events occur: S LP and the HP fuel valves close (Make sure that the FF is at zero). S After 30 seconds set the ENG MAN START to the ”OFF” position, − this causes the closing of the SCV. S Set the ENG START rotary selector to the ”NORM” position, − the ECAM ENGINE page disappears.

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DRY CRANK

WET CRANK

OPEN C/Bs: e.g. for Engine 1 HP FUEL SOV ENG (1KC1)

CLOSE OPENED C/Bs

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LP and MP Shut−Off Valves Commands and Power Supply The ENGine MASTER Lever controls: S The reset of the EEC (Engine Electronic Controller) A and B channels, S The excitation of the ENG MASTER SW SLAVE relay, which controls the LP fuel valve actuator, S The excitation of the airframe shut down solenoid, which controls the HP fuel valve actuator. When the ENG MASTER lever is set to the ”OFF” position, the 28 VDC ESSentiel bus supplies the slave relay. The slave relay switches the power supply from 28 VDC ESS BUS and 28 VDC NORMal bus to the ”SHUT” position of the motor driver of the LP fuel valve With ENG MASTER lever on the ”OFF” position, the LP fuel valve can be open by pulling the breaker 1KC1, if the ENG FIRE P/B is not released. To operate a dry crank the ENG MASTER lever must be set to the ”OFF” position. When a dry crank is initiated, the fuel must lubricate the LP fuel pump, so the breaker 1KC1 (1, 2, 3 or 4) must be pulled. The breaker 1KC1 (1, 2, 3 or 4) is located on the emergency power center (2500VU) in the emergency avionics compartment




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Figure 221 FRA US/T

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ENGINE OPERATION Engine Auto Start To perform an automatic start of the engine: S Set the ENG START rotary selector to IGN START position, S Set the ENG MASTER lever ON, The start valve opens and the N3 rate will increase. At 25% N3 IGN (ignition) and FF (fuel flow) indications come automatically into view on the ENGINE page. Then the light up is automatically initiated by the FADEC, the EGT increases normally. At 50% N3, the ignition stops and the start air valve closes automatically. Before starting, make sure that the EGT is less than 150 ºC (302 ºF). During the engine start sequence; you have to monitor the status of the packs valve. The packs valves have to close automatically when the ENG START rotary selector leaves the NORMal position. The packs valves will reopen once the start sequence has been completed.


NOTE:

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Figure 222 FRA US/T

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Figure 223 FRA US/T

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Engine Manual Start To perform a manual start of the engine: S Set the ENG START rotary selector to IGN START position, S Directly push the MAN START P/BSW to ON, The start valve opens and the N3 rate will increase. S At 25% N3 set the ENG MASTER lever ON, IGN (ignition) and FF (fuel flow) indications come automatically into view on the ENGINE page. Then the light up is automatically initiated by the FADEC, the EGT increases normally. S At 50% N3, deselect the MAN START P/BSW, the start air valve closes automatically and the ignition stops. NOTE:

You must not start the engine if the EGT is more than 150 ºC (302 ºF). If you do so, the EGT will exceed its limit during the engine start. You can dry motor the engine to decrease the EGT.


CAUTION: MAKE SURE THAT THE EGT IS LESS THAN 150 ºC (302 ºF). Be careful and observe all the engine parameters, there is no automatic protection during the engine manual start.

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Figure 224 FRA US/T

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Figure 225 FRA US/T

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Engine Start Faults In some cases the engine start sequence can be aborted due to the detected faults. Among them: S ENG START FAULT, EGT OVERLIMIT: the EGT has reached the maximum allowed EGT. When the fault occurs during an automatic engine start, the EEC proceeds to an automatic abort sequence, the auto dry cranking period decreases to 150_C (EGT). The cranking re−engages approximately at 10%N3, a new start sequence engages with both igniters A and B. If the fault is still present, the start sequence aborts again. When the fault occurs during a manual engine start, there is no automatic abort sequence and the EGT rises. The operator has to initiate an immediate shutdown. The operator has to select the ENG MASTER lever to OFF before the EGT reaches the maximum allowed start temperature (700_C). S ENG IGN A+B FAULT: there is a fault on the ignition exciters or on the igniter plugs. The EGT does not rise; no light up is done and recognized by the EEC. In the engine automatic start sequence the sequence is aborted and a second attempt is initiated after 15 seconds with a dry cranking cycle. If the fault is still present, the start sequence is automatically aborted. S HP FUEL NOT OPEN/NOT CLOSED: the HP fuel valve is stuck in the closed/open position. When the fault occurs during an automatic engine start, the SAV (Start Air Valve) opens, igniter A or/and B are displayed. N3 is cranked but even the N3 is above 25% the fuel flow remains at 0. The operator has to identify the related warnings and the integrated ”FAULT” light on the ENG MASTER lever comes on if the related ENG MASTER lever is kept in the ON position.

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Figure 226 FRA US/T

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FWD Thrust and Mode Settings The forward thrust is adjusted by moving the thrust lever into the different detent points. In function of the selected detent the corresponding thrust mode and indications are displayed on the EWD. Each engine thrust can be adjusted individually. NOTE:

The EEC (Engine Electronic Controller) software does not let the engine operate in the 64% to 72% N1 speed range. Thus, speed increase will stop at 64% N1, until the throttle lever is in a position for engine operation at 72% N1. The EEC software will then let the engine accelerate through the 64% to 72% N1 speed range. This function is called ”KOZ” (Keep Out of Zone) The full engine thrust will be available only when the aircraft speed is above 45 knots This function is called ”METOTS” (Modified Engine Take Off Thrust Settings)




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Figure 227 FRA US/T

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Thrust Control Faults When the two resolvers and the three potentiometers of the same TCA (Throttle Control Assembly) are failed, the ENG THR LEVER FAULT is displayed on the EWD. As a consequence, the EEC does not get the thrust lever position signal any more and the throttle reference, cyan circle related to the thrust lever position does not follow the actual thrust lever demand. When the aircraft is on ground with thrust levers position between IDLE and TOGA, the FDEC automatically select the IDLE thrust. The TPR (Turbofan Power Ratio) is the primary thrust control parameter used on the engine to calculate the engine thrust. It is used in the EEC for engine control. In case of TPR loss, the thrust indications are no longer available. A ”THR XX” message, underlined by an amber arc is displayed on the EWD. The TPR mode automatically reverts to N1 rated mode. The ENG THRUST LOSS warning and actions to be performed are displayed on the EWD. Pressing the ALTN MODE P/BSW ON forces the FADEC to be in N1 mode.




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Figure 228 FRA US/T

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Figure 229 FRA US/T

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Reverse Thrust The aircraft is on ground and the engine 2 and 3 are running. The engines 2 and 3 Throttle Control Levers are equipped with reverser levers to control the deployment or the stowing of the reversers and adjust the reverse thrust. When they are moved (with the Throttle Control Levers into IDLE detent stop only) the corresponding indications are displayed on the EWD. The thrust reverser activation logic activates the deployment of the reversers.




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Figure 230 FRA US/T

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Reverser Fault When the thrust reversers are fully deployed, the REV green indication is displayed on the N1 dial on EWD. This indication is displayed in amber when the reverse is selected and the thrust reverser is stowing or deploying on ground or the reverse is unlocked. In this case the ENG REV UNLOCKED message is displayed and the operator has to push the Throttle Control Lever to FWD IDLE in order to stow the reversers.



POWER PLANT ENGINE GROUND OPEARTION

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POWER PLANT ENGINE GROUND OPEARTION

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Engine Parameters Faults Engine secondary parameters also indicate some faults when they occur. An amber CLOGGED flag under the fuel flow indicates that the fuel filter is clogged. The amber ENG FUEL FILTER CLOGGED message is displayed on the EWD. The rotor vibration levels are also to be monitored VIB N1, N2, N3, when these VIB indications are above a specific threshold they pulses green. In case of fan unbalance, there can be a sound effect. YOU MUST NOT OPERATE THE ENGINE IF THE FUEL FILTER IS CLOGGED. CONTAMINATED FUEL CAN BYPASS THE FILTER AND CAUSE DAMAGE TO THE ENGINE. REPLACE THE LP FUEL FILTER BEFORE YOU OPERATE THE ENGINE AGAIN. The engine turbine overheat is detected by two thermocouples, one located in front of the IP turbine disk and the other located at the rear of the IP turbine disk. When detected, the red ENG TURBINE OVHT message is displayed on the EWD, the EGT increases abnormally but remains under the maximum EGT MCT limit (850_C). The MASTER WARNING lights flash and the Continuous Repetitive Chime sounds. The engine oil low pressure can drop when the engine is running. In this case the green oil pressure value reverts to amber and becomes red if it drops below 25_PSI. The red ENG OIL PRESS LO message is displayed on the EWD, the MASTER WARNING lights flash, the CRC (Continuous Repetitive Chime) sounds and the ENGINE SD page is automatically displayed.


CAUTION:

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Figure 232 FRA US/T

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Figure 233 FRA US/T

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Engine Operation Limits Summary Here are listed the red lines limits of the engine parameters. The maximum rotor operation speeds are: 96.1% N1, 97.8% N2 and 97.8% N3. The MAX EGT is limited to 700_C for ground start, to 850_C for in flight relight and 850_C for maximum continuous operation. The EGT is limited to 900_C for take−off for a maximum of 5 minutes. The residual EGT before start must be less than 150_C, in contrary case the EEC will initiate an automatic crank during automatic start. The minimum oil low pressure with N3 at or above IDLE is limited to 25 psi differential pressure. The maximum oil consumption rate is approximately 0.45 l per hour (0.48 US QT per hour). At start, the oil temperature must be than a minimum of −10_C and the oil temperature must be 60_C or more before acceleration to Take−Off. The maximum allowed oil temperature is 196_C when the operation condition is stable. The engine has an oil quantity of 15 to 17 quarts; the minimum required corresponds to a decrease of 2 quarts below the nominal value. The engine is started by cycles; the maximum continuous motoring is limited to 5 minutes. After a continuous motoring of 5 minutes, the starter must be cooled for 30 minutes before the starter is motored again. The total motoring time permitted is 5 minutes in any 35 minutes period. The starter is activated at 10% N3 on ground, and will require 30% N3 as re−engagement speeds in flight. The rotor vibration levels are also monitored on the ECAM ENGINE page, they are given in cockpit units. An advisory is displayed when the rotor vibration level overpasses the 2.8 cockpit units for N1, 3.6 cockpit units for N2, 3.6 cockpit units for N3. An advisory is also displayed when the nacelle temperature overpasses 300_C.




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Figure 234 FRA US/T

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ENGINE GROUND OPERATION-OMS PAGES Access Procedure The OMS (Onboard Maintenance System) gives access to the historical data necessary for aircraft maintenance. It gives access to various system BITE tests. With the FADEC ground power established by pushing the FADEC GND PWR P/BSW to ON, you ca access to the different OMT pages: Home page, System Report, Test and then ATA 73. In this ATA chapter you can access to the main menus of the EEC and EIPM computers. The EEC will be permanently powered if there is a test in progress via the OMT. The EEC main menu gives access to its both channels ”EEC 2A” and ”EEC 2B”. In these sub−menu are the different functions: Tests, Reports, Engine procedure and Specific Functions. By selecting one of these sub−menus, you will have access to a more detailed level. The EIPM main menu gives access to the related engine systems, ”ENG 2” and ”ENG 4” for EIPM 1 or ”ENG 1” and ”ENG 3” for EIPM 2. In these sub−menu are the different functions: Tests, Reports, and Specific Functions. By selecting one of these sub−menus, you will have access to a more detailed level.

The EEC sub−menus are: Tests, Reports, Engine procedure and Specific Functions. For each EEC channel you will be able to access the Test sub−menu and perform: S An Audible test of the igniters, S A variable−stator−vanes system test, S A test of the P20T20 probe heater, S A hydraulic pump offload test, S And a harness test. Additional tests are provided for each channel of EEC 2 and EEC 3: S Thrust reverser cycling test, S Thrust reverser monitoring test. For each EEC channel you will be able to access the Reports sub−menu and see: S The EEC configuration, S The EGT exceedance report, S The shaft speed exceedance report, S The Inhibition of the thrust reverser.

EEC Menus

CAUTION:


CAUTION:

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WHEN YOU SET THE CONTROLS AS SPECIFIED IN THE PROCEDURE DISPLAYED ON THE OMS, THE DRY CRANK WILL START IMMEDIATELY. THE P20T20 PROBE WILL BE ENERGIZED FOR 5 SECONDS AND GETS HOT DURING THIS TEST. MAKE SURE THAT NOT COVER, CAP OR PLUG IS INSTALLED ON THE P20T20 PROBE. WHEN YOU SET THE CONTROLS AS SPECIFIED IN THE PROCEDURE DISPLAYED ON THE OMS, THE DRY CRANK WILL START IMMEDIATELY. IN THIS TEST YOU MUST LOOK TO SEE IF THE HYDRAULIC PRESSURE INCREASES AND DECREASES AT THE APPLICABLE TIMES.

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THE ENGINE MUST BE STARTED TO PROVIDE THE AIR PRESSURE TO OPERATE THE BLEED VALVES WHEN COMMANDED BY THE EEC. For each EEC channel you will be able to access the engine procedures sub−menu to get: S The fan trim balance procedure, S The engine core washing procedure, S The bleed valve test scheduling. Notice that it is the engine run discrete signal simulation. The engine is not started for this test. For each EEC channel you will be able to access the Specific functions sub−menu to perform: S An engine running simulation, S Reset of the fuel used. NOTE:

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Figure 235 FRA US/T

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EIPM Menus The EIPM sub−menus are: Tests, Reports and Specific Functions. For each engine you will be able to access the Test sub−menu and perform: S A FADEC ground power light test, S An engine light fault, S The system test (full test of the system). For each engine you will be able to access the Reports sub−menu and see: S The discrete inputs reports, S The discrete outputs reports, S The pin programming report. For each engine you will be able to access the Specific functions sub−menu and perform: S An oil low press and ground, S A thrust reverser 3*115V / 25KW power supply, S An ETRAC manual power supply. WARNING:

REVERSE SECOND LINE OF DEFENSE WILL BE DEACTIVATED, BE CAREFUL TO POSSIBLE REVERSE DOORS ACTIVATION.

WARNING:

ETRAC WILL BE POWER SUPPLIED, BE CAREFUL TO POSSIBLE REVERSE DOORS ACTIVATION.




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Figure 236 FRA US/T

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TABLE OF CONTENTS ATA 71−80 ENGI NE RR TRENT 900 . . . . . . . . . . . . . . . . . . 3 ATA 71 POWER PLANT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRENT 900 FOR THE AIRBUS A380−840 . . . . . . . . . . . POWERPLANT EXTERNAL DIMENSIONS . . . . . . . . . . DANGER AREAS OF THE ENGINE . . . . . . . . . . . . . . . . . MAJOR UNITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ACCESS DOORS AND PANELS . . . . . . . . . . . . . . . . . . . . ENGINE COWLING DESCRIPTION . . . . . . . . . . . . . . . . . MAINTENANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE ATTACHMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MOUNTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE DRAINS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DRAINS MAST AND BREATHER OUTLET . . . . . . . . . . . DRAINS TANK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DRAINS TANK OPERATION . . . . . . . . . . . . . . . . . . . . . . . PYLON ELECTRICAL DISCONNECTS . . . . . . . . . . . . . . PYLON ELECTRICAL RECEPTACLES & CONNECTORS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4 4 6 8 10 12 14 16 18 20 22 26 28 30 32

ATA 72 ENGINE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MAIN ROTATING ASSEMBLIES . . . . . . . . . . . . . . . . . . . . ENGINE MAIN BEARING ARRANGEMENT . . . . . . . . . . TRENT MODULAR BREAKDOWN . . . . . . . . . . . . . . . . . . LP COMPRESSOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPINNER ASSEMBLY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN BLADE ASSEMBLY . . . . . . . . . . . . . . . . . . . . . . . . . . . IP COMPRESSOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTERMEDIATE CASE . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP TURBINE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP TURBINE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXTERNAL GEARBOX . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP COMPRESSOR CASE . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE CORE FAIRINGS . . . . . . . . . . . . . . . . . . . . . . . . .

36 36 38 40 42 44 46 48 50 52 54 56 58 60 62

34

FAN BLADE CLEANING . . . . . . . . . . . . . . . . . . . . . . . . . . . INSPECTION OF LPC BLADE & ANNULUS FILLERS . REMOVAL /INSTALLATION OF THE SPINNER & FAIRING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . REMOVAL /INSTALLATION OF THE REAR SPINNER . REMOVAL / INSTALLATION OF THE ANNULUS FILLER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . REMOVAL/INSTALLATION OF THE FAN BLADE . . . . . FAN TRIM BALANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BORESCOPE ACCESS PORTS . . . . . . . . . . . . . . . . . . . . IP COMPRESSOR BORESCOPE ACCESS . . . . . . . . . . HP COMPRESSOR BORESCOPE ACCESS . . . . . . . . . COMBUSTION CHAMBER BORESCOPE ACCESS . . . HP TURBINE BORESCOPE ACCESS . . . . . . . . . . . . . . . TURNING THE LOW PRESSURE (L.P.) SYSTEM . . . . TURNING THE INTERMEDIATE PRESSURE (IP) SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TURNING THE HP SHAFT . . . . . . . . . . . . . . . . . . . . . . . . . RADIAL DRIVE SHAFT REMOVAL/INSTALLATION . . .

64 66 68 70 72 74 76 78 80 82 84 86 88 90 92 94

ATA 73 ENGINE FUEL & CONTROL . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC FUNCTIONS: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC POWER SUPPLY ON GROUND . . . . . . . . . . . . . ELECTRONIC ENGINE CONTROLLER (EEC) . . . . . . . DATA ENTRY PLUG (DEP) . . . . . . . . . . . . . . . . . . . . . . . . DEDICATED ALTERNATOR . . . . . . . . . . . . . . . . . . . . . . . .

96 96 96 98 100 102 104

ATA 73 ENGINE FUEL & CONTROL; ATA 77 ENGINE INDICATING SHAFT SPEED MEASUREMENT . . . . . . . . . . . . . . . . . . . ENGINE PROTECTION SYSTEMS . . . . . . . . . . . . . . . . . P20/T20 PROBE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXHAUST GAS TEMPERATURE (EGT) . . . . . . . . . . . . . EXHAUST GAS TEMPERATURE (EGT) THERMOCOUPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MONITORING UNIT . . . . . . . . . . . . . . . . . . . . . . ENGINE MONITORING UNIT (EMU) INTERFACE . . . .

106 106 108 110 112 114 116 118

Page i

A380 RR 71−80 B12X1

TABLE OF CONTENTS VIBRATION TRANSDUCER . . . . . . . . . . . . . . . . . . . . . . . . T25 THERMOCOUPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . T30 THERMOCOUPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MASTER CONTROL OPERATION . . . . . . . . . .

120 120 122 124

ATA 76 ENGINE CONTROLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THROTTLE CONTROL ASSEMBLY COMPONENT DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . THROTTLE CONTROL ASSEMBLY INTERFACES . . . .

126

ATA 77 ENGINE INDCIATING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE POWER PHILOSOPHY . . . . . . . . . . . . . . . . . . . THRUST CONTROL FUNCTION OPERATION . . . . . . . FADEC ARCHITECTURE & INTERFACE DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

134 134 136

ATA 73 ENGINE FUEL & CONTROL . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC ANALOG AND DISCRETE INPUTS/OUTPUTS . . EEC COMMAND AND SENSOR INTERFACES . . . . . . . EIPM ARCHITECTURE & INTERFACE DESCRIPTION EIPM INTERFACES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIPM & FADEC POWER SUPPLY DESCRIPTION . . . . FADEC TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL SYSTEM INTRODUCTION . . . . . . . . . . . . . . . . . . . FUEL SYSTEM SCHEMATIC & CONTROL . . . . . . . . . . FUEL PUMP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL OIL HEAT EXCHANGER (FOHE) . . . . . . . . . . . . . LP FUEL FILTER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDROMECHANICAL UNIT (HMU) . . . . . . . . . . . . . . . . . HYDROMECHANICAL UNIT (HMU) . . . . . . . . . . . . . . . . . HMU REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . HMU SHUTDOWN SEQUENCES . . . . . . . . . . . . . . . . . . . FUEL FLOW TRANSMITTER . . . . . . . . . . . . . . . . . . . . . . HP FUEL FILTER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL MANIFOLD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL MANIFOLD INSPECTION . . . . . . . . . . . . . . . . . . . . FUEL SPRAY NOZZLES (FSN) . . . . . . . . . . . . . . . . . . . . . LP FUEL FILTER REMOVAL/INSTALLATION . . . . . . . . .

148 148 150 154 156 158 162 168 170 172 174 174 176 178 180 182 184 186 188 190 192 194

126 132

142

INHIBIT THE ENGINE FUEL SYSTEM . . . . . . . . . . . . . .

196

ATA 77 ENGINE INDICATING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE & FADEC SYSTEMS OPS/CTL & IND (RR) . .

198 198

ATA 75 ENGINE AIR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE AIRFLOW CONTROL INTRODUCTION . . . . . VIGV/VSV CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . VIGV/VSV ACTUATORS . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPRESSOR BLEED VALVE SYSTEM . . . . . . . . . . . . BLEED VALVE SOLENOIDS . . . . . . . . . . . . . . . . . . . . . . . IP AND HP BLEED VALVES . . . . . . . . . . . . . . . . . . . . . . . . COOLING & SEALING INTRODUCTION . . . . . . . . . . . . VARIABLE STATOR VANES SYSTEM TEST . . . . . . . . . BLEED VALVE TESTS SCHEDULING . . . . . . . . . . . . . . . TURBINE CASE COOLING SYSTEM (TCC) . . . . . . . . . TURBINE CASE COOLING SYSTEM (TCC) . . . . . . . . . TCC MANIFOLD AND COOLING DUCT . . . . . . . . . . . . . TURBINE OVERHEAT DETECTION SYSTEM . . . . . . . . NACELLE TEMPERATURE MONITORING . . . . . . . . . . . FAN ZONE TEMPERATURE SENSOR . . . . . . . . . . . . . . ZONE 3 TEMPERATURE THERMOCOUPLE . . . . . . . . .

208 208 210 212 214 216 218 220 222 224 226 228 230 232 234 236 238

ATA 79 ENIGINE OIL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE OIL SYSTEM ARCHITECTURE . . . . . . . . . . . . OIL SYSTEM OVERVIEW . . . . . . . . . . . . . . . . . . . . . . . . . FEED OIL, LUBRICATION & COOLING . . . . . . . . . . . . . OIL TANK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE OIL SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . OIL QUANTITY TRANSMITTER . . . . . . . . . . . . . . . . . . . . OIL PUMP AND PRESSURE FILTER ASSEMBLY . . . . MAGNETIC CHIP DETECTORS (MCDS) . . . . . . . . . . . . SCAVENGE FILTER ASSEMBLY . . . . . . . . . . . . . . . . . . . CENTRIFUGAL BREATHER . . . . . . . . . . . . . . . . . . . . . . . FUEL/OIL HEAT EXCHANGER (FOHE) . . . . . . . . . . . . . OIL PRESSURE INDICATION . . . . . . . . . . . . . . . . . . . . . . LOW OIL PRESSURE SWITCH . . . . . . . . . . . . . . . . . . . .

240 240 242 244 246 248 252 254 256 258 260 262 264 264 Page ii

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TABLE OF CONTENTS OIL PRESSURE FILTER DELTA P TRANSDUCER . . . . OIL SCAVENGE FILTER DELTA P TRANSDUCER . . . . OIL TEMPERATURE SENSOR . . . . . . . . . . . . . . . . . . . . . OIL SYSTEM SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . OIL SCAVENGE FILTER REMOVAL/INSTALLATION . . OIL PRESSURE FILTER REMOVAL/INSTALLATION . . EMCD INSPECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EMCD INSPECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESERVATION OF MAIN LINE BEARINGS . . . . . . . . .

266 268 270 272 274 276 278 280 292

ATA 80 STARTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE STARTING SYSTEM . . . . . . . . . . . . . . . . . . . . . . ENGINE STARTING COMMAND CONTROLS . . . . . . . . COCKPIT INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTER CONTROL VALVE (SCV) . . . . . . . . . . . . . . . . . STARTER MOTOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTER OIL SERVICING . . . . . . . . . . . . . . . . . . . . . . . .

294 294 296 298 300 302 304

ATA 74 IGNITION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE IGNITION SYSTEM . . . . . . . . . . . . . . . . . . . . . . . IGNITION SYSTEM COMPONENTS . . . . . . . . . . . . . . . . IGNITION SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . REPLACEMENT OF THE IGNITION LEAD CONTACTS IGNITER INSPECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITER PLUG REPLACEMENT . . . . . . . . . . . . . . . . . . .

306 306 308 310 312 314 316

ATA 78−30 THRUST REVERSER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRENT 900 NACELLE OVERALL PRESENTATION . . . THRUST REVERSER COWL DOORS . . . . . . . . . . . . . . . FAN EXHAUST COWL . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN EXHAUST COWL/THRUST REVERSER COWL . . THRUST REVERSER INHIBITION . . . . . . . . . . . . . . . . . . THRUST REVERSER CONTROL COMPONENT DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER CONTROL FUNCTION OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER MAINTENANCE . . . . . . . . . . . . . . EXHAUST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

318 318 320 322 324 336

VARIABLE FREQUENCY GENERATOR (VFG) . . . . . . . VFG − OIL SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . .

368 370

ATA 26 FIRE PROTECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FIRE/OVERHEAT DETECTORS . . . . . . . . . . . . . . . . . . . .

374 374

ATA 29 HYDRAULIC POWER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE DRIVEN PUMP DESCRIPTION . . . . . . . . . . . . AIRCRAFT PNEUMATIC SYSTEM . . . . . . . . . . . . . . . . . . ENGINE BLEED AIR SUPPLY . . . . . . . . . . . . . . . . . . . . . .

376 376 378 382 384

ATA 30−20 ENGINE AIR INTAKE ICE PROTECTION . . . . . . . . . . . . ENGINE ICE PROTECTION AREAS . . . . . . . . . . . . . . . . AIR INTAKE COWL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NAI SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NAI SHUT−OFF VALVE (SOV) . . . . . . . . . . . . . . . . . . . . . ANTI−ICE PRESSURE REGULATING VALVE . . . . . . . . MODE OF OPERATION AND COCKPIT INDICATIONS DANGER AREAS OF THE ENGINE . . . . . . . . . . . . . . . . . POWER PLANT GROUND OPERATION . . . . . . . . . . . . . ENGINE OPERATION LIMITS . . . . . . . . . . . . . . . . . . . . . . ENGINE START ASSISTANCE OPS/CTL & IND . . . . . . ENGINE START ASSISTANCE DESCRIPTION . . . . . . . ENGINE START / CRANK CONTROL DESCRIPTION . AUTOMATIC START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MANUAL START . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CRANKING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE GROUND OPERATION-OMS PAGES . . . . . . .

386 386 388 392 394 396 398 400 402 406 408 412 428 430 432 434 438 464

340 342 352 366 Page iii

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TABLE OF FIGURES Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure

29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63

The RB211 Family . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Dimension . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Danger Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propulsion System Components . . . . . . . . . . . . . . . . . . . . . . . . Access Doors & Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Cowl − Opening/Closing . . . . . . . . . . . . . . . . . . . . . . . . . . . Preservation of the Powerplant . . . . . . . . . . . . . . . . . . . . . . . . . Engine Attachment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Mounts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drains System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drain System − Leakage Rates . . . . . . . . . . . . . . . . . . . . . . . . . Drains Mast . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drains Tank Location . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drains Tank Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pylon / Powerplant Electrical Disconnects . . . . . . . . . . . . . . . . Electrical Connectors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Rotating Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Bearing Arrangenment . . . . . . . . . . . . . . . . . . . . . . . . . . Modular Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Compressor Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spinner / Fairing Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Blade . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP Compressor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Intermediate Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . External Gearbox . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Compressor Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Core Fairings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Blade Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LPC Blade Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spinner Fairing Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rear Spinner Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Compresssor Blade Removal . . . . . . . . . . . . . . . . . . . . . . .

5 7 9 11 13 15 17 19 21 23 25 27 29 31 33 35 37 39 41 43 45 47 49 51 53 55 57 59 61 63 65 67 69 71 73

Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure

64 65 66 67 68 69 70 71 72 73 74 75 76 77 78 79 80 81 82 83 84 85 86 87 88 89 90 91 92 93 94 95 96 97 98

LP Compressor Blade Removal . . . . . . . . . . . . . . . . . . . . . . . . Fan Trim Balance Weights Position . . . . . . . . . . . . . . . . . . . . . Borescope Access Ports . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP Compressor Borescope Plugs . . . . . . . . . . . . . . . . . . . . . . . HP Compressor Borescope Plugs . . . . . . . . . . . . . . . . . . . . . . Combustion Chamber Borescope Plugs . . . . . . . . . . . . . . . . . Turbine Borescope Plugs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbine the Low Pressure System . . . . . . . . . . . . . . . . . . . . . . Turning the IP System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turning the High Presure System . . . . . . . . . . . . . . . . . . . . . . . Radial Drive Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC − General Architecture & Supply on Ground . . . . . . . Engine Electronic Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Data Entry Plug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Permanent Magnetic Alternator . . . . . . . . . . . . . . . . . . . . . . . . . Shaft Speed Component Location . . . . . . . . . . . . . . . . . . . . . . Engine Protection System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P20/T20 Probe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGT Thermocouple System . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGT Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Monitoring Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EMU Inputs/Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vibration Transducer & T25 Thermocouple . . . . . . . . . . . . . . . T30 Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Master Control Operation . . . . . . . . . . . . . . . . . . . . . . . Throttle Control Assembly Component Description . . . . . . . . Inboard Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outboard Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Throttle Control Assembly Interfaces . . . . . . . . . . . . . . . . . . . . Engine Power Philosophy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIRBUS Cockpit Universal Thraust Emulator (ACUTE) . . . . AIRBUS Cockpit Universal Thraust Emulator (ACUTE) . . . . AIRBUS Cockpit Universal Thrust Emulator (ACUTE) . . . . . FADEC Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

75 77 79 81 83 85 87 89 91 93 95 97 99 101 103 105 107 109 111 113 115 117 119 121 123 125 127 129 131 133 135 137 139 141 143

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99 EEC Digital Interfaces 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 EEC Digital Interfaces 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 EEC Analog and Discrete Inputs/Outputs . . . . . . . . . . . . . . . 102 EEC Command and Sensor Interfaces 1 . . . . . . . . . . . . . . . . 103 EEC Command and Sensor Interfaces 2 . . . . . . . . . . . . . . . . 104 EIPM Architecture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 EIPM Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106 EIPM & FADEC Power Supply 1 . . . . . . . . . . . . . . . . . . . . . . . 107 EIPM & FADEC Power Supply 2 . . . . . . . . . . . . . . . . . . . . . . . 108 Tests − EEC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109 EEC Tests and Specific Functions . . . . . . . . . . . . . . . . . . . . . 110 EIPM Tests and Specific Function . . . . . . . . . . . . . . . . . . . . . . 111 Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112 Fuel System Schematic and Control . . . . . . . . . . . . . . . . . . . 113 Fuel Pump Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114 Fuel / Oil Heat Exchanger (FOHE) . . . . . . . . . . . . . . . . . . . . . 115 Hydromechanical Unit (HMU) . . . . . . . . . . . . . . . . . . . . . . . . . 116 HMU Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 117 HMU Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 118 System Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119 Fuel Flow Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120 HP Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121 Fuel Manifold & Fuel Spray Nozzles . . . . . . . . . . . . . . . . . . . 122 Fuel Manifold Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123 Fuel Spray Nozzle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124 LP Fuel Filter Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125 Inhibiting the Engine Fuel System . . . . . . . . . . . . . . . . . . . . . 126 FADEC System Ops/Ctl & ind . . . . . . . . . . . . . . . . . . . . . . . . . 127 Engine Control Panels Location . . . . . . . . . . . . . . . . . . . . . . . 128 Indication Presentation − EEC Powering . . . . . . . . . . . . . . . . 129 Indication Presentation − Engine Parameters Display . . . . . 130 Throttle Control Levers & A/THR P/B . . . . . . . . . . . . . . . . . . . 131 Airflow Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132 VIGV/VSV Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133 VSV Actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

145 147 149 151 153 155 157 159 161 163 165 167 169 171 173 175 177 179 181 183 185 187 189 191 193 195 197 199 201 203 205 207 209 211 213


134 135 136 137 138 139 140 141 142 143 144 145 146 147 148 149 150 151 152 153 154 155 156 157 158 159 160 161 162 163 164 165 166 167 168

Bleed Valve System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Valve Solenoids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP/HP Handling Bleed Valves . . . . . . . . . . . . . . . . . . . . . . . . . . Cooling & Sealing Airflows . . . . . . . . . . . . . . . . . . . . . . . . . . . . Variable Stator Vane System Test . . . . . . . . . . . . . . . . . . . . . . Bleed Valve Testing Schedule . . . . . . . . . . . . . . . . . . . . . . . . . TCC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCC Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCC Duct Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbine Overheat Detectors . . . . . . . . . . . . . . . . . . . . . . . . . . . Zone 1 and 3 Temperature Monitoring . . . . . . . . . . . . . . . . . . Zone 1 Temperature Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . Zone 3 NAC Temperature Thermocouple . . . . . . . . . . . . . . . Engine Oil System Architecture . . . . . . . . . . . . . . . . . . . . . . . . Oil System Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil − Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil − Servicing Cautions . . . . . . . . . . . . . . . . . . . . . . . Oil Quantity Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pump Assembly & Pressure Filter . . . . . . . . . . . . . . . . . . Magnetic Chip Detector Locations . . . . . . . . . . . . . . . . . . . . . Oil Scavenge Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Centrifugal Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel/Oil Heat Exchanger (FOHE) . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Filter dP Transducer . . . . . . . . . . . . . . . . . . . . . . Oil Scavenge Filter dP Transducer . . . . . . . . . . . . . . . . . . . . . Oil Temperature Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Scavenge Filter Removal . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Filter Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . EMCD Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EMCD Inspection & Washing . . . . . . . . . . . . . . . . . . . . . . . . . . Debris Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

215 217 219 221 223 225 227 229 231 233 235 237 239 241 243 245 247 249 251 253 255 257 259 261 263 265 267 269 271 273 275 277 279 281 282

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169 170 171 172 173 174 175 176 177 178 179 180 181 182 183 184 185 186 187 188 189 190 191 192 193 194 195 196 197 198 199 200 201 202 203

Bearing Lapping Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gear Wear - Fines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearing Failure - Flakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gear Tooth Fragments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chips . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cage Rivet Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Roller Bearing Cage Tang Failure . . . . . . . . . . . . . . . . . . . . . . Build Debris or Swarft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Action to take when debris is discovereed . . . . . . . . . . . . . . . Preservation of Main Line Bearings . . . . . . . . . . . . . . . . . . . . Starting System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . Cockpit Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Staring Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . Start Control Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter Oil Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition System Components . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition System Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition Lead Contact Replacement . . . . . . . . . . . . . . . . . . . . Igniter Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Igniter Removal Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Trent 9000 Nacelle Overalll Presentation . . . . . . . . . . . . . . . Thrust Reverser Cowl Doors . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Exhaust Cowl . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Exhaust Cowl/Thrust Reverser Cowl − Opening . . . . . Fan Exhaust Cowl/Thrust Reverser Cowl − CAUTION . . . . Specific Latch for Inboard Engines . . . . . . . . . . . . . . . . . . . . . Hold Open Rod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Exhaust Cowl/Thrust Reverser Cowl − Closing . . . . . . Manual Opening/Closing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser − Inhibition for Maintenance . . . . . . . . . . . . Thrust Reverser − Inhibition Before Flight . . . . . . . . . . . . . . . Major Component Identification . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Operation 1 . . . . . . . . . . . . . . . . . . . . . . . . . .

283 284 285 286 287 288 289 290 291 293 295 297 299 301 303 305 307 309 311 313 315 317 319 321 323 325 327 329 331 333 335 337 339 341 343


204 205 206 207 208 209 210 211 212 213 214 215 216 217 218 219 220 221 222 223 224 225 226 227 228 229 230 231 232 233 234 235 236 237 238

Thrust Reverser Operation 2 . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Operation 3 . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Operation 4 . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Movement Summary . . . . . . . . . . . . . . . . . . Deploy/Stow the Thrust Reverser Translating Cowl . . . . . . . Thrust Reverser TRPU Deactivation . . . . . . . . . . . . . . . . . . . TLS Deactivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Unlock/Active of PLS at the Thrust Reverser . . . . . . . . . . . . Release/Active the Brake of the PDU . . . . . . . . . . . . . . . . . . Deactivating the MDUs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Manual Ops. of Thrust Reverser Translating Cowl . . . . . . . Exhaust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VFG Air/Oil Heat Exchanger . . . . . . . . . . . . . . . . . . . . . . . . . . VFG Oil Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Servicing − Level Check ... Refilling . . . . . . . . . . . . . . . . . Fire Detection Loop Location . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hyd. Pump Dis-/Re−Engagement 1 . . . . . . . . . . . . . . . . . . . . Hyd. Pump Dis-/Re−Engagement 2 . . . . . . . . . . . . . . . . . . . . Hyd. Pump Dis-/Re−Engagement 3 . . . . . . . . . . . . . . . . . . . . Aircraft Bleed System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engien Bleed Air Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Ice Protection Areas . . . . . . . . . . . . . . . . . . . . . . . . . . Air Intake Cowl . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nacelle Anti-Ice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NAI System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NAI Shut Off Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NAI Pressure Regulating Valve . . . . . . . . . . . . . . . . . . . . . . . . NAI Operating Mode and Indication . . . . . . . . . . . . . . . . . . . . Engine Danger Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ground Operations-Crosseind Condtions . . . . . . . . . . . . . . . Engine Start Assistance Ops/Ctl & Ind. . . . . . . . . . . . . . . . . . Engine Start Controls & Panels Location . . . . . . . . . . . . . . . . Engine Start Assistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

345 347 349 351 353 355 357 359 361 363 365 367 369 371 373 375 377 379 380 381 383 385 387 389 391 393 395 397 399 401 407 409 411 413 415

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A380 RR 71−80 B12X1

TABLE OF FIGURES Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure

239 240 241 242 243 244 245 246 247 248 249 250 251 252 253 254 255 256 257 258 259 260 261 262 263 264

Engine Start Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter Control Valve Override . . . . . . . . . . . . . . . . . . . . . . . . . Engine Manual Start - Starter Valve Open . . . . . . . . . . . . . . Engine Manual Start - Starter Valve Closed . . . . . . . . . . . . . Engine Starting System Description 1 . . . . . . . . . . . . . . . . . . Engine Starting System Description 2 . . . . . . . . . . . . . . . . . . Precautions for Engine Ground Operation . . . . . . . . . . . . . . . Automatic Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Manual Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Cranking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Valve and MPSOV Commands . . . . . . . . . . . . . . . . . . . . . Engine Operation − Engine Auto Start . . . . . . . . . . . . . . . . . . Engine Auto Start Indication . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Operation − Engine Manual Start . . . . . . . . . . . . . . . Engine Manual Start Indication . . . . . . . . . . . . . . . . . . . . . . . . Engine Start Faults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FWD Thrust And Mode Settings . . . . . . . . . . . . . . . . . . . . . . . Thrust Control Faults 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Control Faults 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reverse Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reverser Fault . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Parameters Faults - Fuel, Vibration . . . . . . . . . . . . . . Engine Parameters Faults - EGT, Oil . . . . . . . . . . . . . . . . . . . Engine Operation Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OMS Pages − Access Procedure & EEC Menus . . . . . . . . . OMS Pages − EIPM Menus . . . . . . . . . . . . . . . . . . . . . . . . . . .

417 419 421 423 425 427 429 431 433 435 437 439 441 443 445 447 449 451 453 455 457 459 461 463 465 467

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TABLE OF FIGURES

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TABLE OF FIGURES

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