A320Neo

Training Manual AIRBUS A 320 ATA 71--80 POWER PLANT P&W 1100 NEO

ATA Spec. 104 Level 3

For training purpose and internal use only. Copyright by Technical Training LATAM S.A. All rights reserved. No parts of this training manual may be sold or reproduced in any form without permission of:

Technical Training LATAM S.A. Aeropuerto Int. C.A.M.B., Clasificador 74 Av. Américo Vespucio 901, Renca Santiago -- Chile Tel. +56 (0)2 601 99 11 Fax +56 (0)2 601 99 24

Technical Training LATAM S.A.

POWER PLANT GENERAL

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

POWER PLANT TABLE OF CONTENTS 1.-- Power Plant 2.-- Engine Construction 3.-- Fuel and Engine Control 4.-- Airflows 5.-- Ignition 6.-- Indicating 7.-- Lubrication 8.-- Starting 9.-- Exhaust

For Training Purposes Only

10.-- Abbreviations and Glossary

SCL /JGB / REV.00 / Jun--2016

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POWER PLANT GENERAL POWER PLANT Overview The PW1100G--JM turbofan engine powers the Airbus A320 NEO ( New Engine Option) aircraft. It is axial-- flow, twin spool turbofan engine with a ultra--high bypass ratio, low speed gear--driven fan. The engine includes core-mounted Angle and Main gearboxes, and is mounted on a pylon that extends below and forward of the wing leading edge. Also know as the power plant, the engine supplies propulsive energy to the aircraft and provides electrical power and hydraulic pressure for aircraft systems. In adition it supplies pressurized air for the aircraft Environmental Control System (ECS) that includescabin pressurization, heating and cooling. The power plant includes the basic engine with its control components, the nacelle, engine mountys, and engine buldup units (EBU). The engine is controlled by the Full Authority Digital Electronic Control (FADEC) system amd is designed for safe and reliable operation. Propper ground run danger zones must be observed for the power plant operation, as well as steps for preservation and replacement. Advantages of the PW1100G--JM Engine Conventional gas turbine engines cannot perform as efficiently as the PW1100G--JM. The low compressor and low turbine are restricted to less-than--optimal operating speeds so that fan speed can be maintained in a range most efficient for fan diameter. The geared technology of the PW1100G--JM allows the fan and low rotor to operate at optimal, independent speeds for peak efficiency. These improvements in performance reduce fuel consumption, air pollution, and noise. At the same time, operating costs and environmental impact are drastically reduced. WARNING:


AIRBUS A-- 320 PW 1100 NEO 71 -- 00

Aircraft and Engine Specifications

PW 1100G Design

Thrust

24.000 to 35.000 Lbs

By Pass Ratio

12 : 1

Engine Inlet Diameter

Overall Nacelle

102 inch (259 cm)

Fan Blade Tip

81 inch (205.7 cm)

Engine weight of demountable power plant

7457 Lbs ( 3382 Kg)

Aircraft Models

A319, A320, A321

Passenger capacity

124 to 235

Sample Namig Convention A320 Family PW

Prat & Whitney

1

Engine Model 1000

1

Airbus Airframe

33

Thrust in Lbs x 1000

G

Geared Turbo Fan (GTF)

J

Japanese Aero Engines Corporation (JAEC)

M

Motored und Turbine Union Aero Engines (MTU)

BE CAREFUL WHEN YOU WORK ON THE ENGINE AFTER SHUTDOWN. THE ENGINE AND ENGINE OIL CAN STAY HOT FOR A LONG TIME. IF YOU DO NOT OBEY THIS WARNING, INJURY CAN OCCUR. REFER TO THE MSDS FOR ALL MATERIAL USED AND THE MANUFACTURER S SAFETY INSTRUCTIONS FOR ALL EQUIPMENT USED. IF YOU DO NOT OBEY THIS WARNING, INJURY CAN OCCUR.


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Figure 1 SCL /JGB / REV.00 / Jun--2016

Engine Profile Page: 3


POWER PLANT GENERAL

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

NACELLE SYSTEMS Description Nacelle systems components are mounted to the engine and to the pylon. They provide the engine with these capabilities: * an aerodynamic and protective enclosure for engine mounted components * collection and discharge of oil, fuel, and hydraulic fluid from the engine and its components. Nacelle system types and their components are listed in the table at right.

System Type

Components

E i Mounted Engine M t d

* Inlet Cowl * Exhaust E h t Nozzle N l * Exhaust Plug

Pylon y Mounted

* Front Mount * Rear Mount * Thrust Links


Engine g Mounting g

* Fan Cowl * Thrust Reverser


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Nacelle Components Page: 5


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AIRBUS A-- 320 PW 1100 NEO 71 -- 00

INLET COWL Purpose The inlet cowl’s aerodynamic barrel smooths airflow, providing uniform pressure as air reaches the fan. Location: The inlet cowl is secured to the engine fan case Flange A by the aluminum inlet attach ring. The outer barrel is a two--piece assembly extending from the inlet lip interface to the leading edge of the fan cowl. Two splice joints are located on the outer barrel at 5:00 and 7:00.


Description: The cowl’s outer skin provides even airflow across the engine nacelle. The inner skin forms the engine inlet and acoustic treatment. The forward bulkhead and aft bulkhead provide impact protection and structural support for the inlet assembly. Lightning strike protection is provided by an expanded copper screen layer impregnated into the outer barrel assembly. A panel in the outer barrel provides access for maintenance of the Thermal Anti--Ice (TAI) duct that supplies the inlet lip with hot air. Anti--ice air exits the TAI vent located at 6:00 on the inlet lip. Opening the right fan cowl on the aft bulkhead provides access to the anti--ice supply line. A second panel serves the T2 probe and the wiring harness and sense line that are routed aft across the inner barrel to the aft bulkhead mounted interface. A small flush inlet vent scoop located on the inlet outer barrel provides fan compartment cooling.. The inner barrel has gravity drainage holes embedded within the core.


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Inlet Cowwl Page: 7


POWER PLANT GENERAL

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

FAN COWL Purpose: The fan cowl provides aerodynamic smoothness and a protective enclosure for the engine fan case and accessories. Fan cowl doors provide maintenance access to components and systems shown below. * Anti--ice temperature and pressure sensors * Electronic Engine Control EEC * Prognostics and Health PHMU Management Unit * Pylon disconnects * Ignition exciter box * Thrust reverser torque box *Thrust Reverser Actuation TRAS System

A fan cowl assist side latch secures the right--hand side fan cowl half to the inlet cowl. The side latch facilitates fan cowl closing by one person. The Oil Tank Access Door (OTAD) provides quick access to the oil tank, allowing oil service without opening the cowling. The OTAD is hinged at the front edge and secured with two latches that provide redundancy at the aft edge of the panel. Each of three fan cowl door latches has a fan cowl door proximity sensor to alert the ground crew to the fan cowls position. Each sensor is able to detect whether the relative latch is locked, to avoid any fan cowl door loss in flight.


Location: The fan cowl covers the engine fan cases and is positioned between the inlet and reverser cowls. Description: Fan cowl doors are a one--piece structure. Two fan cowl axial locators per door align the cowls as they close. A copper mesh is embedded in the fan cowl laminate for lightning strike protection. Cowl doors are manually opened, and held open with forward and aft Hold Open Rods (HOR). The fan cowl is secured to the pylon above the engine by pins installed through the cowl door hinges and pylon fittings. The fan cowls are also secured beneath the engine by three tension latches. Visual indicators on the latch handles show when the latches have not been properly secured. An aerodynamic strake deflects airflow as required in certain maneuvers related to aircraft performance. The strake is mounted to the fan cowl outer surface on the inboard and outboard side. A fan cowl vent provides overpressure protection in the event of a burst anti--ice duct. Drain holes at the bottom of the fan cowls provide fluid drainage. Each door has three hoist provisions used for removal and installation.


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AIRBUS A-- 320 PW 1100 NEO 71 -- 00

HOLD OPEN RODS (HOR) OIL TANK


AXIAL LOCATOR (4)

ACCESS DOOR (OTAD) OUTLET VENT/

FAN COWL LATCH

PRESSURE RELIEF

PROXIMITY SENSOR (3) Figure 4 SCL /JGB / REV.00 / Jun--2016

Fan Cowls Page: 9

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

THRUST REVERSER Description Thrust reverser cowl doors are comprised of two halves that are mechanically independent. The halves hinge at the pylon, latching together along the bottom split line. They can be opened using the Door Opening System (DOS) by means of a hydraulic hand pump. Each door is equipped with a Hold Open Rod (HOR).



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Thrust Reverser Cowls Page: 11


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AIRBUS A-- 320 PW 1100 NEO 71 -- 00

COWL STRAKES Purpose Strakes are mounted to the fan cowl outer surface to improve flight characteristics by controlling airflow. Location: Strakes are mounted to both the left and right fan cowl doors.


Description: Strakes decrease the turbulence of the airflow between the fan cowl door assembly and the wing. The strake is attached to the inboard and outboard fan cowl door by 14 fasteners engaged into floating nut plates, located on the fan cowl interior skin.


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Cowl Strakers Page: 13

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

ENGINE MOUNTS Description The engine mounts transfer engine loads to the pylon. Mount assemblies have four functions: * support the weight of the engine * transmit the thrust of the engine to the pylon * prevent the engine from turning on its axis * hold lateral loads. Two mount assemblies and one sub--assembly are located between the engine and the pylon: the forward and aft mounts, and the thrust links, respectively. The forward mount assembly is connected at the top of the engine’s Compressor Intermediate Case. The aft mount assembly is connected at the top of the engine’s Turbine Exhaust Case. The thrust link sub--assembly is connected to the Compressor Intermediate Case at approximately 9:30 and at 2:30, and to the forward mount through a balance beam.



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Mounts Page: 15

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

FORWARD MOUNT ASSEMBLY Description The forward mount is attached to the Compressor Intermediate Case, and is connected to the pylon. The forward mount supports side and vertical loads using a two--link arrangement with bolts loaded in shear. Main components of the forward mount assembly are below. * Main beam * Side links (2) * Shear pin slots (2) * Fail--safe bolt The side and vertical loads at the forward mount couple with the aft mount loads to support overall engine pitch and yaw. Side links provide the primary load paths from the fan case into the front beam. The forward mount is attached to the pylon with four bolts and two shear pins, which transmit vertical and shear loads into the pylon. The mount bolts use captive barrel nuts to ease removal.



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Forward Mounts Page: 17

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

AFT MOUNT ASSEMBLY AND TORQUE TUBES Description The aft mount is attached to the Turbine Exhaust Case (TEC) and reacts to engine fore--aft, side, vertical, and roll loads. Main components of the aft mount assembly are below. * Main beam * Outer links (2) * Shear pin slots (4) * Fail safe bolt The two outer links are assembled to the beam with four shear bolts. Each outer link is attached to the TEC with one shear bolt. The center link is attached directly to the TEC with another shear bolt. The failure of any one link on the rear mount, including the thrust links, will cause the system to transfer loads to a secondary load path.



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Aft Mounts Page: 19

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

ENGINE DRAIN SYSTEM Description The Engine Drain System collects and discharges oil, fuel, and hydraulic fluid from the engine and pylon, delivering residuals to the lower bifurcation drain mast through dedicated drain tubes. Drain components are located on the left and right sides, and at the bottom of the engine. Drain paths have gaps around the Latch Access Panel (LAP) and eight drain holes in the lower bifurcation fixed/access panels. The LAP has outlets that allow Main Gearbox, engine component and pylon drain tubes to drain fluids overboard from the nacelle. Related components from each of eight outlets are identified on the pylon tube at the drain mast and at placards on the LAP. Engine components that drain through the drain mast are listed below. * 2.5 bleed valve actuator * HPC primary and secondary stator vane actuators * LPC stator vane actuator * Integrated Drive Generator IDG * Integrated Fuel Pump and Control IFPC * Hydraulic Engine Driven Pump EDP * Ecology collector tank



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Engine Drain Page: 21


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AIRBUS A-- 320 PW 1100 NEO 71 -- 00

DRAIN MAST LEAKAGE LIMITS Drain Location

Fluid

Leakage Limit CC/ Hrs

Drops/Min

IFPC Carbon Seal Hydraulic Pump Carbon Seal

Oil

10

3.3

15

5

2

1

IDG Carbon Seal 2.5 Bleed Actuator LPC Stator Vane Actuator HPC Primary Staor Vane Actuator

Fuel

HPC Secondary Staor Vane Actuato


IFPC Seal


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Drain Mast Schematic Page: 23

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

ENGINE GROUND RUN DANGER ZONES General The high velocity, high temperature, and toxicity of discharged exhaust gases can be dangerous. Jet wakes in the exhaust area can be significant. Entry corridors and hazard areas are shown for personnel guidance when approaching an operating engine.



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Danger Areas on Idle Power Page: 25

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

NOTES:



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Danger Areas on Take Power Page: 27

Technical Training LATAM S.A. For Training Purposes Only

POWER PLANT GENERAL

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

PRESERVATION AND REPLACEMENT Preservation of the Engine Engine preservation provides maximum protection to critical engine components such as gears, bearings, and accessory components against damage from excessive moisture, debris, and other environmental conditions. Preservation is used for the protection of out--of--service engines, whether on--wing or off, and for engines to be put into storage. Preservation methods include: * Method 1: Preserve The Engine For 60 Days Or Less * Method 2: Preserve The Engine For More Than 60 Days. Use the preservation method that is necessary for the time that the engine will be in storage. If you are not sure as to the length of storage time, use Method 2. Method 1: Preservation of Engine for 60 Days or Less Preserve the Engine Oil System * Change the engine oil. Drain, replace oil filter and fully service the system. * Motor the engine for five minutes. The fan must be turning. * Drain the oil system and remove the oil filter. Preserve the Gearbox * Inspect the housing for hydraulic fluid, and clean according to standard practices procedures. * Spray oil on the gearbox pads and install pad covers. Seal the Engine for Storage * Install covers over the engine intake area. * Install a cover over the engine exhaust area. Make a Record of Preservation for Each Engine The record should note the following information: * method of preservation * date of preservation * oil system is drained and empty.


If the engine will continue to stay in storage for more than the 60 day limit, you can choose between the two options listed below. * Option 1: Preserve the Engine For More Than 60 Days (Method 2) * Option 2: Operate the engine Option 2 may only be used to repeat Preserve the Engine For 60 Days Or Less a maximum of one time. Option 2 may not be used if the engine stays in storage more than a total of 120 days. If the engine is to remain in storage longer than 120 days, then it is necessary to preserve the engine using Method 2, Preserve the Engine For More Than 60 Days. Preservation Chart 1-- This chart does not specify all the steps in the preservation method. 2-- Sampling option is restricted to one 60 days period. 3-- In option B method I, oil system flush is not necesary if sample test results are ok. 4-- When preserving for more than 60 days, fuel system preservation is also necesary. 5-- If the engine will stay in storage after 60 days, you can preserve the engine again after the initial 60 days period. See text for instructions.

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Preservation Chart Page: 29


POWER PLANT GENERAL Method 2: Preservation of Engine for More Than 60 Days It will be necessary to air motor or run the engine in a Test Cell to preserve the engine for more than 60 days. Preserve the Engine Oil System * Change engine oil, drain, replace the oil filter, and fully service the system. * Motor the engine for five minutes. The fan must be turning. * Drain the oil system after fuel system preservation and remove the oil filter. Preserve the Engine Fuel System * Drain the fuel system by removing the plug from the Fuel/Oil Manifold below the IFPC. Then re--install the plug. * Attach the preservation adapter and drain tube. * Use the preservation cart to pump preservation fluid into the system while wet motoring the engine. * Remove tooling and reconnect the fuel lines. Change Starter Oil Preserve the Gearbox * Inspect the housing for hydraulic fluid, and clean according to standard practices procedures. * Spray oil on the gearbox pads and install pad covers. Preserve The Engine For Relative Humidity * Relative humidity must be at 40 percent or less inside the engine during the preservation time. * Put dehydrating agent in the engine. Put half in the inlet and half in the tail pipe areas. * Put relative humidity indicators inside the inlet and exhaust areas. * Seal the engine for storage. * Install covers over the engine intake area. * Install cover over the engine exhaust area. Protective covers must have windows so that you can see the relative humidity indicators inside the engine.


AIRBUS A-- 320 PW 1100 NEO 71 -- 00 Make a Record for Each Engine Make a record with the following information: * method of preservation * date of preservation * confirmation that oil system is drained and empty. Storage Examine the engine during preservation (every 15 days or less) and refer to General Instructions For The Preservation of Engines for Storage, SPOP 428.

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Engine Preservation Page: 31

AIRBUS A-- 320 PW 1100 NEO 71 -- 00

NOTES:



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AIRBUS A-- 320 PW1100 NEO 72 -- 00

ENGINE CONSTRUCTION



POWER PLANT ENGINE


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POWER PLANT ENGINE

AIRBUS A-- 320 PW1100 NEO 72 -- 00

PW1100 ENGINE General Description The PW1130G engine is an axial--flow turbofan with a twin--spool low and high pressure rotor system. The engine is electronically controlled through Full Authority Digital Engine Control (FADEC) with an axial flow turbofan and has a low pressure ratio fan supplied torque by a driveshaft, but turning at a lower speed through a Fan Drive Gearbox (FDG). The FDG is a planetary gear system in a star configuration. Also, the fan case includes the forward mount location for the nacelle cowling hardware. Downstream from the fan there are two isolated airstreams. The first is the primary airflow, it moves through the engine core to make high temperature, pressurized gasses that supply a propulsive force to the aircraft. Secondary (or outer) airstream is mechanically compressed by the fan on entry to the engine and is exhausted through the Fan Exit Guide Vanes (FEGVs) to make thrust. The primary airstream goes into the compressor section which includes a three stage Low Pressure Compressor (LPC) and an eight stage High Pressure Compressor (HPC). The LPC and HPC compress the air before it moves into the diffuser and annular combustor. Expanded gas pressure from the combustor is sent through vanes to turbine section to turn the two stage High Pressure Turbine (HPT). The torque made by the HPT turns the HPC which share a high pressure rotor. The remaining pressurized gas moves to the three stage Low Pressure Turbine (LPT). The torque made by the LPT turns the LPC and FDG which share a low pressure rotor. The low pressure rotor connects the fan rotor through the FDG. The primary engine structure is made from four frames, Fan Intermediate Case (FIC), Compressor Intermediate Case (CIC), Turbine Intermediate Case (TIC) and the Turbine Exhaust Case (TEC). The engine cases form the primary structure of the engine when bolted together and function as a support for all of the inner parts through struts and bearings. The engine has five bearing compartments and seven bearings. The engine includes a main gearbox installed to core that supplies torque to turn the aircraft accessories. There is also an angle gearbox which transmits torque from the high pressure rotor to the main gearbox and accessories. During engine start the torque is sent the opposite direction.


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General Description Page: 3


POWER PLANT ENGINE

AIRBUS A-- 320 PW1100 NEO 72 -- 00

ENGINE MODULES Fan Rotor and Fan Case The fan rotor module has a low pressure ratio fan rotor located at the forward end of the engine. The rotor includes 20 fan blades with fan blade fairings located between each fan blade root platform to keep a smooth airflow between the blades. The fan rotor is supplied torque, but at a decreased speed, through the FDG connected to the fan rotor by a gearbox front shaft. The function of the fan rotor is to supply the primary source of engine thrust through the initial compression of ambient air. Compressed air that leaves the fan rotor is divided by the FIC and flows through both the fan nozzle (secondary flow) and the LPC to the core of the engine (primary flow). The fan rotor is housed in the fan case assembly. The function of the fan case assembly is to contain and make the fan airflow straight, to give protection and containment for a fan blade failure, and to supply structural attachment points between the inlet cowl and the core engine. The inner case has acoustic liners to decrease noise at the front and rear of the case, a rubstrip liner for the fan blades, and a liner which supplies protection from ice. At the rear of the fan case assembly is an aluminum support ring which has a V groove that supplies alignment and functions as a support for the thrust reverser doors. The fan case assembly also includes 48 FEGVs. They are located at the rear of the fan case and extend diagonally forward to the FIC to function as a support for the fan case and make the bypass air straight. Fan Drive Gearbox (FDG) The FDG is a planetary gear unit in a star arrangement that decreases the fan shaft speed from the LPT shaft. The FDG gets the torque from the LPT and uses it to turn a sun gear. This sun gear then turns five planet gears against an outer ring gear set which is connected to the gearbox front shaft. The gearbox front shaft is connected to the fan rotor. There are journal bearings at each planet gear position. With these bearings, the cylindrical surfaces of the gear inner diameters turn against their mating pivots. The gears are housed in a carrier which is attached at the rear to the FIC by inner and outer assemblies. Pins through the inner carrier function as a radial support. The journal bearings, FDG gears, and inner pins are lubricated with pressure oil from an oil manifold at the rear of the FDG. Oil from the journal oil shuttle valve goes through the oil manifold, journal bearing oil supply tubes, and journal bearing oil filters to the journal bearings. The Variable Oil Reduction Valve SCL /JGB / REV.00 / Jun--2016

(VORV) controls the flow of oil through the VORV oil manifold to the gear oil filters and nozzles and then to the FDG gears and the inner pins. A windmill pump, which turns at fan speed, is located in the compartment. The gravity valve in the front of the assembly changes the flow of oil during flight conditions. During normal engine operation (when main oil pump pressure supplies the FDG bearings ), the gravity valve lets oil from the compartment auxiliary reservoir be sent by the windmill pump to the main oil tank through the journal oil shuttle valve. During windmill conditions, the gravity valve lets oil from the sump go through the windmill pump to the journal bearings through the journal oil shuttle valve. During zero and negative G conditions, the gravity valve lets oil from the auxiliary reservoir go through the windmill pump to the journal bearings through the journal oil shuttle valve. Fan Intermediate Case (FIC) The FIC is made from titanium and located between the fan case and the LPC. It gives an aerodynamic flow path between the fan rotor and the LPC and also functions as a support between the fan case and engine core through its connection to the FEGVs. The FIC houses the FDG and No. 2 bearing. The No. 2 bearing functions as a support axially and radially for the LPC and fan input coupling which supply torque to the FDG. There are eight hollow struts that are supports and supply a passage through the airflow for oil, air, and electrical components. Low Pressure Compressor (LPC) The LPC is located aft of the FIC. The three stage LPC initially compresses the fan air on the path to the core of the engine (primary flow). The LPC rotor is supplied torque by the LPT. The LPC rotor hub is attached to the LPT shaft which connects the LPC to the fan rotor through the fan input coupling and FDG. The FDG lets the LPC turn at a higher speed than the fan rotor for increased efficiency. At the inlet to the LPC, there is one stage of Variable Inlet Guide Vanes (VIGVs). The remaining part of the LPC has three Integrally Bladed Rotors (IBRs) and two stator stages. To increase the range of operation and to reject dirt from the LPC airstream, there is an annular LPC (Stage 2.5) bleed valve at the rear of the LPC.

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Engine Modules Page: 5


POWER PLANT ENGINE Compressor Intermediate Case (CIC) The CIC functions as the support between the LPC and the HPC. The case includes nine aerodynamic struts, two primary engine mounts, and a redundant engine mount. The CIC also transmits the thrust loads from the engine at two thrust mounts. The module includes a fire containment ring that both supplies protection for the LPC structure from the core nacelle fire zone and functions as a support for the thrust reverser inner cowl. The CIC also houses the No. 3 bearing compartment. The No. 3 bearing functions as a radial and axial support at the front of the HPC. A pressurized layer of oil supplied to the circumference of the No. 3 bearing assembly support absorbs rotor radial movement and decreases vibration. The compartment includes carbon face seals at the front and rear and labyrinth seals to control the flow of oil and air in the compartment. Forward of the No. 3 bearing is the No. 3 bearing bevel gear which engages a gearbox drive bevel gear to supply torque between the HPC and the angle gearbox. High Pressure Compressor (HPC) The HPC increases the speed, pressure and temperature of the primary gaspath air and supplies this air to the diffuser and combustor. There are eight stages to the HPC. The VIGVs (part of the CIC) and the first three stages of vanes in the HPC are variable. The 4th to 7th stage stators are fixed (cantilevered) and seal against the adjacent rotor abrasive. Knife edge seals on the rotors seal against the stator vanes. The rotor stages are held together with a rotor shaft. This shaft connects the front hub with the rear hub and extends rearward to the HPT. The HPC is held radially and axially at the front by the No. 3 bearing. It is held radially at the rear by the No. 4 bearing. The outer case wall is made of the front case set between the CIC and the diffuser case. The outer air seals make the inner case wall. The HPT supplies torque to the HPC. A rotor nut at the rear hub holds the HPC rotor assembly together. The compressor supplies air to the customer bleed system and other engine systems. The front six stages of the rotor are decreased in temperature by bleed air. The rear two stages of the compressor are compartmentalized from the front by a ring on the rotor shaft. The rear rotor compartment temperature is controlled by airflow slots in the spacer and the rotor tube assembly.


AIRBUS A-- 320 PW1100 NEO 72 -- 00 Diffuser and Combustor The diffuser case, combustion chamber, and turbine nozzle supply the hot exhaust gasses of internal combustion to the turbine modules. The diffuser also functions as a structural support for the HPC and HPT cases. Compressed air flows from the exit of the HPC into the diffuser and combustor section. During entry, the airflow is made straight while the contour of the inner diffuser case lets the air expand before it moves to the combustion chamber. Metered fuel is supplied to the combustion chamber through 18 fuel nozzles and the fuel and air mixture is burned inside the combustion chamber. The HPT 1st stage vanes point the high temperature, high velocity gasses, out of the combustion chamber to supply the turbines and make thrust at the exhaust nozzle. High Pressure Turbine (HPT) The HPT module changes some of the high temperature, high velocity gas energy into torque. The 1st stage turbine vanes point the gasses out of the combustion chamber at an efficient angle which turns the HPT. The HPT module is a rotor system with two stages and 44 blades in each of the rotors. The rotor system is housed by the HPT case which contains the 26 Blade Outer Air Seals (BOAS) for each stage. Cool air flows through holes in the BOAS and into the gaspath. The HPT case is the containment structure for the turbine rotors and functions as structural support to the diffuser case and also houses the TIC. Turbine Intermediate Case (TIC) The TIC is housed in the HPT case and is made of an inner case, a turbine stator assembly, and the No. 4 bearing assembly. The TIC is between the HPT and the LPT. The TIC assembly uses eight TIC rods as supports to connect the inner case to the HPT case through the hollow vanes in the TIC stator assembly. The TIC rods function as supports for the No. 4 bearing compartment. The TIC stator assembly has 16 hollow airfoils that turns the gaspath airflow to align with the LPT rotor. Pressure, scavenge, and drain oil tubes from the No. 4 bearing compartment go through the hollow vanes and also a buffer air tube goes through into the No. 4 bearing compartment.

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Engine Modular Page: 7


POWER PLANT ENGINE Low Pressure Turbine (LPT) The LPT is made of a three stage rotor, a turbine case assembly, two stator vane stages, and an LPT driveshaft. The LPT provides the rotational force for the LPC and FDG. The energy from the hot combustion gasses is changed into torque by the LPT blade and rotor assemblies. There are 84 1st stage, 86 2nd stage, and 78 3rd stage solid nickel alloy shrouded blades that have a fir--tree root to attach them to the hub.


Turbine Exhaust Case (TEC) The TEC assembly is a main structural part of the engine and attaches to the rear flange of the LPT. The TEC assembly functions as a support for the No.5 and No. 6 bearing assemblies and has connection points for the aft engine mounts. There are bosses in the outer wall of the TEC for four Exhaust Gas Temperature (EGT) probes that extend into the gaspath. The TEC is a one piece case welded structure with ten airfoil--contour vanes (struts) which hold two bearing housings at its inner flanges. The No. 6 bearing is an oil--damped type in which a layer of pressurized oil around the bearing outer race absorbs engine vibration. The scavenge oil tube and bearing pressure oil tube go through the bottom struts of the TEC to supply oil to the No.5 and No. 6 bearings.

AIRBUS A-- 320 PW1100 NEO 72 -- 00 Angle Gearbox (AGB) The AGB connects the CIC to the MGB through the layshaft of the AGB. The AGB housing is a support for the ball and roller bearings and gear shafts and contains a borescope opening and oil drain plugs. The AGB is connected to the CIC with four bolts to the CIC flange. During engine operation the torque from the HPC shaft is transmitted to the gearbox drive shaft and into the AGB. The AGB then transmits the torque with gear shafts through the layshaft to the MGB. Roller and ball bearings hold each gear shaft in the axial and radial positions. In the AGB, oil lubricates the ball and roller bearings and gearshafts. The oil is supplied from external oil supply tubes. Scavenge oil that exits the AGB housing is sent to the oil tank.

Main Gearbox (MGB) The function of the MGB is to supply torque to oil, fuel, and hydraulic system components connected to the main gearbox. The main gearbox is supplied torque by the HPC and AGB. The MGB housing is installed below the HPC and contains ball and roller bearings, carbon seals, and gear shafts which connect the different system components. The deoiler is an internal component of the main gearbox and removes oil from lubrication breather mixtures. There is an opening to turn the N2 system by hand so the HPC and HPT rotors can be inspected. In the MGB, oil lubricates the ball and roller bearings and the gearshafts. The oil is supplied to the MGB housing from an external oil supply tube. Scavenge oil that exits the MGB is sent to the oil tank by the lubrication and scavenge pump. As much as possible the MGB uses internal casting openings to supply oil to the different bearings and components and to supply fuel to different components. This makes many external tubes not necessary and decreases the risk of leakage.


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Engine Modular Page: 9

AIRBUS A-- 320 PW1100 NEO 72 -- 00

GASPATH CONFIGURATION Description Gaspath configuration is a term describing the engine modules that make up the primary path of airflow through the engine. A module is an assembly of parts that can be installed or removed from the engine as a single unit. Gaspath modules and their stage counts are listed below. Each stage is made up of a single rotor assembly and its complementing stator assembly. Note that in the compressor section, the rotor precedes the stator. In the turbine section, the rotor follows the stator.



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Modular Gaspath Page: 11


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

ENGINE FLANGES Description Engine flanges are external features of the engine case that serve various structural purposes, including: * joining module assemblies together * supporting the engine s streamlined enclosure known as the nacelle * supporting the brackets used to mount engine components. Flanges and the modules they support are detailed at right. They are shown in conjunction on the graphic below.

Flange

Modules

A

Inlet Cowl Attachment

D

V Ring Grove

E

High Pressure Compressor

H

Diffuser Case

M

High Pressure Turbine

HPT

N1

Low Pressure Turbine

LPT

P

Turbine Exhaust Case

TEC

To

Exhaust Sleeve Nozzle Attachment

Ti

Exhaust Plug Attachment

HPC


ENGINE STATIONS Description Engine stations are locations in the gaspath. Key stations along the gaspath have pressure probes and temperature sensors. Signals from the probes and sensors are transmitted through the engine’s electronic control component to the flight deck. Stations are illustrated in the graphic below. Each sensor uses the number of the engine station as part of its name. Examples are shown below. Sensor Naming Convention T3 T = Temperature 3 = Station 3 PT2.5 PT = Pressure and Temperature 2.5 = Station 2.5 Stations are illustrated in the graphic below.


Station 2

Location Engine Core Inlet

2.5 3 4 4.5 5 1.4

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Engine Flanges and Stations Page: 13


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

ENGINE MAIN BEARINGS Description Bearings support the weight of engine parts and permit one surface to roll over another with minimal friction and wear. The weight of the parts is transmitted through balls or rollers that are contained by two raceways. Bearings are designed from materials that can withstand extreme pressure, since they must absorb the axial and radial loads of rotating assemblies. An axial load is transmitted parallel to the bearing shaft, and a radial load is applied perpendicular to the shaft. Bearings are lubricated, cooled, and cleaned by oil. The PW1100G--JM uses three types of bearings, described in the chart. Each bearing type holds engine parts in alignment to transmit their load. Note that tapered roller bearings operate like ball bearings, while requiring less space than standard roller bearings. Bearings support the weight of engine parts and permit one surface to roll over another with minimal friction and wear. The weight of the parts is transmitted through balls or rollers that are contained by two raceways. Bearings are designed from materials that can withstand extreme pressure, since they must absorb the axial and radial loads of rotating assemblies. An axial load is transmitted parallel to the bearing shaft, and a radial load is applied perpendicular to the shaft. Bearings are lubricated, cooled, and cleaned by oil. The PW1100G--JM uses three types of bearings, described in the chart. Each bearing type holds engine parts in alignment to transmit their load. Note that tapered roller bearings operate like ball bearings, while requiring less space than standard roller bearings. Five compartments contain a total of seven bearings. Descriptions are shown at right. Oil--damped bearings use a thin film of oil between the outer race and the bearing support to reduce vibration. Note: the low pressure and high pressure rotors are often referred to as N1 and N2, respectively Bearing types are illustrated in the graphic below.


Bearings 1

Type

Oil Damped

Support Function

n

1,5

Tapered Roller

Fan Rotor and Fan Drive G Gear S System t (FDGS)

2

Ball

n

Front N1 Rotor (LPC)

3

n

Front N2 Rotor (HPC)

4

n

Rear N2 Rotor (HPT)

5 6

Roller

Rear N1 Rotor (LPT) n

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Engine Bearings Page: 15

AIRBUS A-- 320 PW1100 NEO 72 -- 00

BEARING COMPARTMENT HARDWARE Description Each compartment uses carbon seals to prevent oil leakage. * Carbon face seals seat axially against a rotating ring. * Some carbon face seals are called lift--off seals because extra air is used to physically lift the seal off the ring during engine operation. Bearings and their seal types are shown at right. Each compartment has a scupper that returns any start--up leakage oil.

Bearing No 1 -- 1,5 -- 2 -- 4 -- 5 -- 6 3

Seal Type Face Lift--off



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Main Engine Bearing Compartment Hardware Page: 17

AIRBUS A-- 320 PW1100 NEO 72 -- 00

ENGINE MODULES General A module is the largest assembly of engine parts that can be treated in one of two ways: -- removed or installed from the engine as a unit -- disassembled or preassembled, independently of other modules. PW1100G--JM assembly modules are as follows. * Fan rotor (including inlet cone) * Fan case * Fan Drive Gear System FDGS * Fan Intermediate Case FIC * Low Pressure Compressor LPC * Compressor Intermediate Case CIC * High Pressure Compressor HPC * Diffuser/Combustor/ High Pressure Turbine nozzle * High Pressure Turbine HPT * Turbine Intermediate Case TIC * Low Pressure Turbine LPT * Turbine Exhaust Case TEC * Angle Gearbox AGB * Main Gearbox MGB



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Engine Modules Page: 19


POWER PLANT ENGINE FAN ROTOR Purpose: The fan rotor draws in ambient air and provides the first of 12 stages of compression necessary to yield more than 90% of the thrust produced by the engine. Location: The fan rotor is located at the front of the engine. The inlet cone and cover are at the front of the fan rotor.


AIRBUS A-- 320 PW1100 NEO 72 -- 00

Description: The fan rotor includes the inlet cone and 20 fan blades with integrated fairings and reinforced leading edges. Fan diameter is 81 inches. The fan rotor is supported by bearing nos. 1 and 1.5, which are tapered roller bearings. The fan rotates in a clockwise direction as viewed from the aft end of the engine looking forward. The fan is connected to the fan drive shaft. Power to turn the fan is supplied by the fan drive gear through the fan drive shaft. Description of the fan rotor assembly. See Figure 1 (Sheet 1 of 2) and Figure 1 (Sheet 2 of 2). The PW1100G engine has a 20--blade fan with an 81--inch tip diameter. The blades are a hollow aluminum design that has a dovetail root that engages slots in the fan hub. The blade design (which includes a reinforced leading edge) decreases impact damage (birdstrike or FOD) and noise. No mid-- or part--span shrouds are necessary. Front and rear lock rings hold the fan blades in their correct axial position in the hub. Composite under--root spacers beneath the fan blades provide a radial preload of the blades to reduce fan rotor vibration. It is possible to replace fan blades in as assembled engine and in a fan rotor removed from the engine. Counterweights can be added to the outer rim of the fan hub when necessary. The fan rotor is supported by two tapered--roller bearings with an oil--sealed compartment between them.


The fan is driven by the LPC/LPT through the Fan Drive Gear System (FDGS), a gear reduction unit which decreases the fan speed. Because the fan speed is lower than that of the LPC/LPT, there is increased efficiency across the full engine operating range. At the front of the fan rotor is the inlet cone and the inlet cone cover. These parts are made of composite material and keep the flow of air into the engine smooth. The fan rotor is attached to the gearbox shaft from the Fan Drive Gearbox (FDG) with a nut and lock at the end of the shaft.

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Fan Rotor Page: 21


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

INLET CONE Purpose: At the front of the rotor, the inlet cone and its cover smooth the flow of air to the engine. Location: The inlet cone is attached to the fan hub.


Description: The inlet cone is made of composite material. Its front flange provides attachment for the inlet cone cover. The inlet cone s rear flange is bolted to the fan hub and is part of the fan blade retention system. The inlet cone is anti--iced with a continuous flow of 2.5 bleed air. The cone has holes around its circumference for venting anti--ice air.


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Inlet Cone Page: 23


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN BLADE Purpose: Fan blades accelerate the air entering the engine, producing the majority of thrust and providing airflow to the primary gaspath to be used for combustion and cooling. Location: Fan blades are located on the fan hub.


Description: The 20 fan blades are partially hollow and made of aluminum, with a dovetail root to engage slots in the fan hub. Composite Teflon wear strips are bonded to the pressure surfaces of each fan blade dovetail to prevent wear on the blade root pressure surfaces and to reduce fan rotor vibration. Axial retention of the blades is provided by front and rear lock rings. Composite fan blade spacers are installed beneath the fan blades to provide a radial preload of the blades, which also reduces fan rotor vibration. The spacers are mechanically trapped by the front and rear lock rings. An erosion coating is applied on the airfoil and a titanium strip is bonded to the leading edge of each blade to reduce leading edge erosion.


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Fan Blades Page: 25


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN CASE ASSEMBLY Purpose: The Fan Case Assembly contains and directs the fan airstream, sending part of the air directly through the gaspath and the majority of air outside the gaspath as bypass air. The fan case also provides the structural link between the inlet cowl and the core engine. In case of 1st Stage fan blade failure, the fan case will contain the liberated blade. Location: The Fan Case Assembly is located between the inlet cowl and the Fan Intermediate Case (FIC).


Description: The Fan Case Assembly is made up of the fan case, Fan Exit Guide Vanes (FEGVs), fan exit liner segments and fan exit fairing and support. The fan case is a one--piece, composite case with an acoustically treated inner surface that decreases noise. A fan blade rub strip area protects fan blades from contact with the fan case. An ice liner protects the case against ice shed by fan blades. An aluminum support ring at the top rear of the fan case has a Vgroove that provides alignment and support for the thrust reverser doors. On the inner surface of the fan case, 48 hollow aluminum composite Fan Exit Guide Vanes extend to the Fan Intermediate Case. The stationary FEGVs straighten the fan air and also provide radial support between the FIC and the Fan Case Assembly. A fan exit liner assembly goes around the outer area of the LPC. Louvers in these fan exit liner segments release 2.5 bleed air from the Low Pressure Compressor into the fan stream at the correct angle.


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Fan Case Page: 27


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN EXIT GUIDE VANE Purpose: The Fan Exit Guide Vanes (FEGVs) straighten and direct the fan discharge airstream. They supply structural support of flight and blade--out loads and also provide radial support of the engine. Location: The Fan Exit Guide Vanes are located aft of the fan blades.


Description: A total of 48 hollow aluminum composite FEGVs extend diagonally rearward from the outer diameter of the Fan Intermediate Case to the inner diameter of the fan containment case. The aluminum composite material was chosen for its strength and weight savings. The FEGVs straighten fan bypass air and provide radial support for the Fan Case Assembly. A titanium strip is bonded to the leading edge of each FEGV to protect against erosion. Each FEGV is attached to the Fan Intermediate Case along the inner platform and to the fan containment case at the outer platform.


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

SUPPORT

FAN

RING

CASE

FAN DRIVE GEARBOX


Fan Exit Guide Vane Page: 29


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN DRIVE GEAR SYSTEM (FDGS) Purpose: The Fan Drive Gear System allows the fan and low spool (LPC/LPT) to operate at different speeds, improving performance and efficiency, respectively.


Location: The FDGS is located between the fan rotor and the Low Pressure Compressor (LPC) and is attached to the Fan Intermediate Case. Description: The FDGS is made up of a central sun gear surrounded by five star gears which are supported by journal bearings and an outer ring gear. A torque frame and flex mount help with alignment of the FDGS input coupling and with fan alignment to the fan drive gear, also reducing extreme loads that can be transferred to the FDGS from the fan and LPC rotors, such as at takeoff. The front of the FDGS is supported by bearings nos. 1 and 1.5. The FDGS has an auxiliary oil supply to lubricate the journal bearings during a negative g--force or windmill event that would affect normal oil flow. The auxiliary system uses a windmill/auxiliary pump during these conditions. The FDG is a ary gear reduction unit that takes the torque from the LPT and uses it to turn a sun gear. This sun gear then turns star gears against an outer ring gear that is connected to the fan hub. There is a ratio of LPC/LPT to fan hub speed of approximately three to one, and the fan turns in the opposite direction from that of the LPC/LPT. Note: The fan hub turns clockwise as seen from the rear of the engine, and the LPC/LPT system turns counterclockwise as seen from the rear. This gives the effect of a straighter gaspath.


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AIRBUS A-- 320 PW1100 NEO 72 -- 00


L P C


Fan Drive Gear Page: 31


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN DRIVE SYSTEM Operacion The No. 1 and 1.5 bearing support assembly holds the No. 1 and 1.5 bearing compartment. The No. 1 and No. 1.5 bearings are tapered roller bearings that hold the gearbox shaft from the fan drive gearbox (FDG) axially and radially. There are labyrinth and carbon--face seals at the No. 1 bearing side to control oil and airflow. The No. 1 bearing is oil--damped and uses a pressurized oil film around its outer race to absorb fan vibration. In the PW1100G design, the FDG gears and the carrier to which they are attached do not move around the sun gear. The sun gear engages the star gears and turns them in a direction that causes the ring gear (and fan) to turn in the opposite direction from the sun gear (and LPC/LPT) at a lower speed. This design keeps rotating masses to a minimum and keeps centrifugal loads away from the star gear bearings. The most important effect of this is a lower fan speed and a higher LPC speed, which increases compressor efficiency. The FDG uses journal bearings at each star gear position. With these bearings the cylindrical surfaces of the gear inner diameters turn against their mating pivots with no balls or rollers. These bearings are lubricated with pressure oil from a manifold at the rear of the FDG. Oil which goes through the gear bearing areas is caught in an oil channel around the circumference of the FDG and goes by centrifugal force to the auxiliary oil tank in the No. 1 bearing support and to the sump. A windmill pump is installed in the No. 1 bearing support. This pump engages a pump drive gear on the gearbox shaft and turns at fan speed. Under conditions when it is possible that the flow of oil from the main oil pump to the FDG will not be continuous (for example after shutdown when wind could turn the fan), oil from the sump is distributed to the windmill pump and from there to the FDG to keep oil pressure in the gear bearings.


During usual engine operation the main oil pump pressure supplies the FDG bearings sufficiently. Under zero and negative--G conditions, oil from the auxiliary oil tank and the sump go through the windmill pump to the FDG gears. Struts in the No. 1 bearing support give access to the bearing compartment for oil, air, and instrumentation components. See Figure The FDG is attached at its rear side to the fan intermediate case by a torque frame and flex mount. This makes it possible for the FDG to align easily with the LPC at the rear and the fan at the front. See Figure

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FDG Operation Page: 33

AIRBUS A-- 320 PW1100 NEO 72 -- 00

FAN INTERMEDIATE CASE Description The FIC module is the support structure between the fan and Fan Drive Gearbox (FDG) and the Low Pressure Compressor (LPC). This group includes the inlet vanes at the front of the intermediate case, the intermediate case, the fan input coupling, and the LPC Variable Inlet Guide Vanes (VIGV) at the rear. The FIC and fan exit fairing divide the airflow from the fan into the bypass airstream and the engine core airstream. The core airstream is straightened by the fan exit stators and FIC vanes to increase efficiency for the LPC. The rear of the fan intermediate case contains the No. 2 bearing support assembly (which contains the No. 2 bearing and seals). The No. 2 bearing is a ball bearing and holds the LPC and fan input coupling axially and radially. The No. 2 bearing inner race halves are installed on a coupling which supplies the torque of the LPC and Low Pressure Turbine (LPT) to the FDG. There are knife edge and carbon seals at the No. 2 bearing compartment rear side to control oil and airflow. This bearing has oil damping features. A pressurized film of oil around the No. 2 bearing support assembly absorbs rotor radial movement and decreases vibration. The LPC inlet guide vanes point the core airflow in the direction that is most efficient for LPC and engine operation.



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Fan Intermediate Case Page: 35


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

LOW PRESSURE COMPRESSOR (LPC) MODULE Description The LPC is to the rear of the Fan Intermediate Case (FIC). This compressor increases the pressure of the gas path air from the rear of the fan that goes through the LPC inlet vanes in the FIC. The LPT shaft supplies torque to the LPC. The LPC turns counterclockwise as seen from the rear, and the fan (forward of the LPC) and High Pressure Compressor (HPC) (rear of the LPC) turn clockwise. The result of this is a straighter airflow through the compressor and higher efficiency. The LPC is connected at its front to the FDGS Input Coupling which drives the fan rotor through the Fan Drive Gear System (FDGS) which is a reduction gear unit. Because the FDGS lets the fan turn at a lower speed, the LPC is permitted to turn at a higher speed for increased efficiency. The higher speed makes possible an increased pressure ratio per stage. Because of this, fewer LPC parts are necessary. There is one variable inlet vane stage that is part of the FIC group, followed by three LPC rotor and two LPC stator stages. The disks and blades of each rotor stage are of the Integral Blade Rotor (IBR) design, in which the blades and disk are one piece. An LPC rotor hub is attached to the inner diameter of the 2nd stage rotor with tie rods and is connected to the LPT shaft to the rear of the No. 2 bearing. Counterweights are available at the front and rear of the LPC rotor for balance. Note: Compressor stage numbers start at the front of each rotor and stator assembly. There is a fan stage but it has no stage number. The first stage of the LPC is the LPC 1st stage rotor, followed by the 2nd and 3rd stages. The low compressor blades have coated abradable tips to increase efficiency. There are instrumentation ports at the rear of each vane stage through which borescope inspection or airfoil repair is possible. To increase operational flexibility there is an annular LPC (2.5) bleed valve at the rear of the LPC controlled by an actuator, rod, and bellcrank linkage. It is possible to remove the LPC as a module (either after the fan and the fan drive gear system are removed, or as a fan/LPC module assembly). SCL /JGB / REV.00 / Jun--2016

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Low Pressure Compressor Module Page: 37


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

COMPRESSOR INTERMEDIATE CASE (CIC) MODULE Description The CIC is the support structure between the Low Pressure Compressor (LPC) and the High Pressure Compressor (HPC). The CIC contains two primary engine mounts, a redundant engine mount, and two thrust link mounts. This case contains the No. 3 bearing, seals, and the oil pressure and scavenge tubes related to this bearing compartment. At the front of the gas path area in the CIC is the LPC exit stator. This stator turns the air from the LPC and directs it to the HPC. At the rear of the CIC is the HPC inlet case and vane assembly. This is a variable--vane stator assembly in which the angle of each vane is controlled by an arm at its outer end, attached to a synchronizing ring. A linkage from this ring connected to an external actuator turns the vanes to an angle which is best for the engine operation condition. Note: The first three vane stages of the HPC are variable--vane type also. An actuator controlled by both an engine vane control and by signals from the Electronic Engine Control (EEC) causes vane angle movement in the CIC and HPC variable vane stages. At the rear of the CIC is the No. 3 bearing and bevel gear assembly. The No. 3 bearing is a ball bearing and is the thrust bearing for the HPC which holds the front of the HPC radially and axially. This bearing has oil--damping features. A pressurized film of oil around the No. 3 bearing support assembly absorbs rotor radial movement and decreases vibration. The No. 3 bearing bevel gear is forward of the No. 3 bearing. This gear engages a gearbox drive bevel gear and supplies torque from the HPC/HPT through the gearbox drive shaft to the angle gearbox. At the rear of the CIC is the fire seal which closes the opening between the fan exit liner of the Fan Intermediate Case (FIC) and the front of the CIC. The No. 3 bearing area uses buffer air to control airflow. With this type of pressurization, air from the HPC is used to keep internal bearing compartment pressures at the correct level.


The No. 3 bearing front face seal and support assembly at the front of the CIC controls air and oil leakage. Note: The No. 3 bearing front and rear face seals are wet--face carbons seals of a lift--off configuration. Grooves in the faces of the sealing seats push away the carbon seals at high speed, with decreased friction and wear and no increased leakage. A labyrinth seal around LPC rear hub knife--edges controls airflow into the compartment. The No. 3 bearing front and rear face seals control oil in the bearing compartment.

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Compressor Intermediate Case Page: 39


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

HIGH PRESSURE COMPRESSOR (HPC) MODULE. Description The HPC increases the speed and pressure of the primary gas path air and supplies this air to the diffuser and combustor. There are eight stages to the HPC, with numbers that start with stage 1. The first seven rotor stages are of the Integral Blade Rotor (IBR) design, in which the disk and blades are one piece to decrease air leakage. The 8th stage rotor is a disk and blades assembly. The eight rotor stages are held together with a tie shaft that extends from front to rear through the rotor stages. The first three stages of HPC stator vanes are variable--vane design, in which the vane angles are adjusted as necessary for engine operation conditions. An actuator controlled by the Electronic Engine Control (EEC) moves a linkage to turn the vanes of each stage at the same time to the necessary angle. Knife--edge seals on the 2nd through 4th stage rotors seal against the inner seal lands of the variable stators to control the flow of the air. The remaining 4th through 8th stage stators are of a fixed--vane type in which each vane extends inward against an abradable surface on the adjacent rotor, with no vane inner foot or air sealing ring. The HPC gets its power from the High Pressure Turbine (HPT) system. The HPT is attached to the rear of the HPC rotor shaft. The HPC is held radially and axially at the front by the No. 3 ball bearing in the Compressor Intermediate Case (CIC). It is held radially at the rear by the No. 4 roller bearing at the end of the HPC rotor shaft. Between the 7th and 8th stage rotors are eight HPC rotor tubes held in a ring. These tubes turn with the rotor and push air to the center of the rotor where this air can flow to where it is necessary for cooling and pressure balance. These tubes prevent vortices which can cause blockage of necessary internal rotor air movement. Air from the HPC is used to keep downstream parts (for example the HPT) cool and is used as buffer air to pressurize the engine bearing compartments. At the outer case walls there is an instrumentation port in the plane of each blade leading edge for borescope inspection and possible blending of all blades.


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High Pressure Compressor Page: 41

AIRBUS A-- 320 PW1100 NEO 72 -- 00

DIFFUSER CASE Description The diffuser case extends from the High Pressure Compressor (HPC) to the High Pressure Turbine (HPT). This case contains the HPC air pressure and directs it to the combustion chamber. The diffuser inner case attaches to the HPC exit stator at the front section. It controls the direction of the HPC air with an HPC air seal to keep air loss to a minimum. The diffuser inner case has a divergent gas path contour in which the gaspath air from the HPC is expanded and decreased in velocity, to supply static pressure to the combustion chamber. The outer wall of the diffuser case holds the fuel nozzles and the igniter plugs, which extend into the combustion chamber. The HPC exit air temperature and burner pressure are monitored by the T3 probe and P3 sensor which are also installed on the diffuser case wall. The diffuser case has three borescope plugs, where two provide access to the 7th and 8th stage HPC blades and one provides access to the 1st stage HPT and combustion chamber.



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Diffuser Case Page: 43


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

COMBUSTION CHAMBER AND 1ST STAGE NGV Description The combustion chamber and High Pressure Turbine (HPT) 1st stage nozzle assembly is to the rear of the diffuser case. The combustion chamber is a sheet--metal structure in which fuel and air are mixed and burned. High Pressure Compressor (HPC) gas path air comes in from the front through the diffuser case and exit stator, and a spray of fuel from the fuel nozzles makes a fuel--air mixture which can burn and expand the gas pressure. Technology for Advanced Low NOx (TALON) combustion chambers are designed for decreased emissions (NOx and smoke). The combustion chamber is a TALON X type combustion unit. The combustion chamber assembly contains an inner and outer liner and a hood and bulkhead at the front. The bulkhead has ports for the fuel nozzles and openings for HPC air. The inner and outer liners have combustion holes (which supply combustion air used to burn the fuel) and dilution holes (which mix the air and fuel correctly). This type of combustion chamber is a floatwall design. The inner and outer liners have cast segments which it is possible to replace and which have a thermal barrier layer. There are ports in the sides of the outer liner for igniter plugs which are used to start and continue fuel combustion. The combustion chamber mixes air and fuel and burns this mixture to expand the volume of gas. The gas then moves to the HPT 1st stage nozzle, which turns the gases through vanes at the correct angle to go into the HPT. The 1st stage nozzle contains 32 vanes. The vanes are hollow and let air that flows around the combustion chamber through holes in the vane airfoils and across the airfoil surface. This air keeps the vane surfaces cooler. There is also a thermal barrier layer on the vanes to help stop heat damage. A 1st stage cooling air duct at the rear of the 1st stage nozzle collects cooling air from the diffuser area to use in the 1st stage vanes and in the 1st stage blades of the HPT.


This cooling air duct, of the Tangential On--Board Injection (TOBI) type, supplies cooling air to the HPT 1st stage seal and hub at the correct angle for their contour. Seal surfaces at the rear of the duct assembly and mating knife edges on the HPT 1st stage front air seal control air leakage.

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Combustion Chamber and 1St Stage NGV Page: 45


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AIRBUS A-- 320 PW1100 NEO 72 -- 00

HIGH PRESSURE TURBINE Description The HPT is to the rear of the diffuser case. The HPT case holds the Turbine Intermediate Case (TIC) at the rear of the turbines. The HPT uses the gas pressure from the High Pressure Compressor with the increased fuel energy from the combustion chamber to turn a two stage turbine. The torque of this turbine turns the HPC through the HPC shaft. The HPT and HPC turn clockwise while the Low Pressure Compressor (LPC) and Low Pressure Turbine (LPT) turn counterclockwise for increased efficiency. The HPT has a 1st stage hub and a 2nd stage hub held together with a retaining nut on the HPC rotor shaft. The 1st stage hub engages a spline in the HPC rear hub at the front and the 2nd stage hub is attached by a nut on the HPC rotor shaft at the rear. A 1st stage seal at the front of the HPT assembly and a 2nd stage seal at the rear of the assembly have knife--edge seals which turn against sealing rings to control air leakage. These 1st and 2nd stage seals have counterweight flanges to make it possible to correct unbalance of the HPT assembly. There are 44 blades in each of the HPT 1st and 2nd stage hubs, which have a controlled clearance with HPT case seals at their tips. The seals are in segments that let them move as a result of temperature changes. An Active Clearance Control (ACC) system uses fan air manifolds around the outer HPT case to change blade tip clearances. When necessary (for example during climb or cruise operation), fan discharge air in the manifold applied against the HPT case outer rails makes the case cooler and decreases the inner diameter of the 1st and 2nd stage seal segments. This decreases the 1st and 2nd stage blade tip clearances and increases the efficiency of the turbine. The HPT 1st stage blades are kept at as low a temperature as possible with diffuser area air from around the combustion chamber.


This air flows through the cooling air duct and into openings in the hub and blade roots and from there into the blade airfoils. Air from the HPC 7th stage and air released into the diffuser flows downstream through the diffuser area to the rear of the HPT 2nd stage to the roots of the 2nd stage blades. 1st and 2nd stage blade retaining plates (at the HPT 1st stage hub rear and the 2nd stage hub front) keep the blades in position and let cooling air through into the hubs and into the blades. The HPT 2nd stage vane internal temperatures are kept as low as possible with a Turbine Vane and Blade Cooling (TVBC) system in which 6th stage HPC air is supplied to the HPT 2nd stage vane leading edges. The rear of the HPT is held radially by the No. 4 bearing in the aft of the HPT case.

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High Pressure Turbine Page: 47


POWER PLANT ENGINE

AIRBUS A-- 320 PW1100 NEO 72 -- 00

TURBINE INTERMEDIATE CASE Description The TIC is between the High Pressure Turbine (HPT) and the Low Pressure Turbine (LPT). The TIC is aft of the HPT and held internally by the HPT case, which attaches directly to the LPT case. A stator assembly between the inner and outer cases has 16 hollow vanes, each with an airfoil contour, which turn the HPT gas path airflow to align with the LPT (the HPT turns clockwise as seen from the rear, and the LPT turns counterclockwise). The eight TIC rods connect the TIC inner case to the HPT case through hollow stator vanes. Pressure, scavenge, and drain oil tubes from the No. 4 bearing compartment go through three of the remaining hollow vanes, and two buffer air tubes go through a remaining boss into the compartment to supply cooling air to the outer No. 4 bearing housing. The stator vanes give protection to the rods and tubes from gas path heat. Attached to the TIC inner case is a No. 4 bearing support which contains the No. 4 bearing. This roller bearing is the axial support for the rear of the HPT and is of an oil--dampened type in which a pressurized oil film around the bearing outer race absorbs rotor vibration. The No. 4 bearing is installed to the rear of the HPT. This lets the bearing and its related oil system operate in a cooler area of the engine. Pressure oil lubricates the bearing and a scavenge tube removes the oil. Buffer air to keep oil in the compartment comes through two buffer air tubes. Seal seats to the front and rear of the bearing inner race turn against these carbon seals and have a lift--off configuration. There is No. 4 bearing compartment oil that is controlled by wet--face carbon seals at the front and rear of the compartment lift--off grooves on the seal seat faces which cause the carbon seal to move away from the seat by air pressure at high rotor speeds. As a result, there is less friction and wear, with no increased oil leakage.


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Turbine Intermediate Case Page: 49


POWER PLANT ENGINE

AIRBUS A-- 320 PW1100 NEO 72 -- 00

LOW PRESSURE TURBINE Description The LPT is to the rear of the High Pressure Turbine (HPT) case. Gas flow from the HPT is turned by the Turbine Intermediate Case (TIC) vanes and is changed to torque by the LPT. The LPT is connected to the Low Pressure Compressor (LPC) through the LPT shaft and uses this torque to turn the LPC. The LPT and LPC turn counterclockwise as seen from the rear. The HPT and High Pressure Compressor (HPC) turn clockwise. The LPT has three stages, an LPT 1st and 3rd stage attached to an LPT 2nd stage hub with a single circle of tiebolts. Note: Each compressor and turbine rotor has stages whose number starts at one. For the LPT the stages are the 1st through 3rd. Gas leakage at the LPT blade tips is controlled by outer air seal segments. These segments keep a controlled clearance with the adjacent blade tip knife edges. The LPT has an Active Clearance Control (ACC) system. With this system, a Turbine Cooling Air (TCA) valve (energized by the EEC) controls air from the 4th stage of the HPC and sends it to a panel assembly around the LPT case. This HPC air from the ACC cooling panels makes the LPT case cooler and decreases its diameter. This decreases the diameter of the outer air seal segments which causes tighter clearances at the 1st through 3rd stage blade tips and increased engine efficiency. The LPT blades are cooled by the Turbine Vane and Blade Cooling (TVBC) system. Gas leakage between the rotor stages is controlled by fin--type inner air seal segments attached to inner seal supports at the LPT 2nd and 3rd stage vane inner feet. These segments keep a controlled clearance with knife edges on the 1st and 3rd stage turbine rotor flanges.


Borescope inspection of all LPT stages is made possible by instrumentation bosses with access to all three blade leading edges. An instrumentation boss in the HPT case gives access to the LPT 1st stage blades, and bosses in the LPT case in the planes of the LPT 2nd and 3rd stage vanes give access to the LPT 2nd and 3rd stage blade leading edges. The LPT 1st through 3rd stage blades are solid.

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Low Pressure Turbine Page: 51

AIRBUS A-- 320 PW1100 NEO 72 -- 00

TURBINE EXHAUST CASE Description The TEC Module is to the rear of the low pressure turbine (LPT). This module is the static structure for the rear of the LPT and also contains the rear mount for the engine. The TEC is a welded structure in which airfoil--contour vanes are also struts which hold the bearing compartment at the inner center of the assembly. The turbine exhaust case contains the No. 5 and No. 6 bearings. These are roller bearings which hold the rear of the LPT radially. The No. 6 bearing is of the oil--damped type in which a film of pressurized oil around the bearing outer race absorbs engine vibration. Oil leakage is controlled by ring--type carbon seals at the No. 5 bearing side. No. 5 and No., 6 bearing pressure oil and scavenge oil tubes go through the bottom vanes of the TEC. There are mount lugs at the top of the TEC, and ground handling holes in the sides of the TEC. There are bosses in the outer wall of the TEC for eight (8) exhaust gas temperature (EGT) probes to extend into the gaspath.



POWER PLANT ENGINE


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Turbine Exhaust Case Page: 53


POWER PLANT ENGINE

AIRBUS A-- 320 PW1100 NEO 72 -- 00

MAIN GEARBOX ASSEMBLY Description The main gearbox (MGB) is a cast--aluminum housing installed below the engine that contains gearsets to connect the engine rotor torque to different components. The High Pressure Compressor/High Pressure Turbine (HPC/HPT) rotors supply torque to the MGB. This is done with a gearbox drive shaft in the Compressor Intermediate Case (CIC) which engages the No. 3 bearing bevel gear on the front end of the HPC. The HPC/HPT rotational speed (N2) is decreased with a sequence of gearsets along the gearbox drive train from the tower shaft area through the Angle Gearbox (AGB) to the Main Gearbox (MGB). This decreased shaft speed and increased torque is used in the MGB to turn components of the electrical, fuel, oil, hydraulic, and starter systems. There are eight gearshaft axes in the MGB. As much as possible the MGB uses internal casting passages to supply oil to the different bearings and components and to direct fuel to different components. This makes many external tubes not necessary and decreases potential for leakage. Where the component drives go through the MGB front or rear wall, there are carbon seals to prevent oil leakage. The MGB has a deoiler at the left rear which removes oil vapor from the internal breather air before this air is released out of the engine. The MGB internal gear shafts have roller bearings at their ends except for the Integrated Fuel Pump and Control (IFPC) and starter which have ball bearings. A starter installed at the rear of the MGB can (when necessary) turn the HPC/ HPT rotors through the AGB and No. 3 bearing bevel gear. A crank port is supplied at the rear of the MGB housing to make it possible to turn the HPC/HPT rotors when necessary for repair or inspection. The crank port can be used for borescope inspection to trouble shoot the internal parts of the MGB.


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Main Gearbox Page: 55


AIRBUS A-- 320 PW1100 NEO 72 -- 00


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MGB Components Page: 56


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MGB Components Page: 57

AIRBUS A-- 320 PW1100 NEO 72 -- 00

ANGLE GEAR BOX Description The Angle Gearbox (AGB) connects with the gearbox drive shaft that comes out of the Compressor Intermediate Case (CIC) area. Power from the High Pressure Compressor/High Pressure Turbine (HPC/HPT) rotors is used to operate components of the electrical, oil, fuel, hydraulic, and starter systems attached to the main gearbox. The AGB is a 120--degree gearset that sends the HPC/HPT rotor torque from the CIC gearbox drive shaft to the main gearbox to which the oil, hydraulic, electrical, and starter components are attached. The starter turns the HPC/HPT rotors through the AGB and the CIC gearbox drive shaft when necessary. The AGB is installed on the engine core at the approximate 6 o’clock position as seen from the rear. There is a spline in the upper gear shaft and a mating spline on the CIC gearbox drive shaft. The upper drive gear engages the lower drive gear which sends torque to the main gearbox. Roller and ball bearings hold each drive gear in the axial and radial positions. A gearbox drive shaft (layshaft) extends from the lower drive gear to the mating drive spline of the main gearbox and has a cover. The AGB bearings are lubricated with oil from external oil supply tubes connected to the angle gearbox housing. Scavenge oil is returned to the oil tank by an external oil tube to the lubrication and scavenge pump. The AGB has a borescope port which makes internal inspection of the gearbox possible.



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Angle Gearbox Page: 59

AIRBUS A-- 320 PW1100 NEO 72 -- 00

NOTES:



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AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL SYSTEM



POWER PLANT OIL SYSTEM


Page: 1


POWER PLANT OIL SYSTEM OIL SYSTEM General Oil has many functions in a jet turbine engine. Because parts turn against each other, oil is necessary to lubricate them. Oil between mating metal parts keeps the surfaces apart, and the result is less friction and wear. Because parts operate in hot areas, it is also necessary to keep parts as cool as possible. Oil that flows across hot surfaces removes heat from those surfaces, and in an oil--oil heat exchanger, cool oil can remove heat from hot oil. A pressure oil system puts oil where it is necessary. A scavenge oil system must then remove the oil, and (where necessary) a breather system must release bearing compartment air pressure that was the result of oil pressurization. It is important to keep oil clean, and there must be an oil filter system that removes unwanted material from the oil before it is used. To keep too much heat from the oil system, a thermal management system is necessary. This system uses heat exchangers which use available flows of air, fuel, or oil to remove heat from the oil that comes out of the engine. Oil system Control The flow of oil into the bearing compartments and in and out of the heat exchangers is controlled by a central Oil Control Module (OCM). This OCM is attached to the left side of the main gearbox. Oil flows into and through this module through external oil tubes and internal gearbox and OCM housing openings. Attached to the OCM are the Main Oil Temperature Sensor (MOT), Low Oil Pressure Switch (LOP), and Main Oil Pressure Sensor (MOP) which measure oil pressures and temperatures. Also attached are the Active Oil Damper Valve (AODV) and Fuel/Oil Cooler (FOC) bypass valve which, when necessary, are controlled by signals from the Full Authority Digital Engine Controls (FADEC). Attached to the Variable Oil Reduction Valve (VORV) manifold is a Journal Oil Shuttle Valve (JOSV) to control oil to the Fan Drive Gear System (FDGS) journal bearings, and a VORV which schedules oil flow to the FDGS gear teeth and the No. 1, No. 1.5, and No. 2 bearings. The VORV manifold also contains the Auxiliary Oil Pressure Sensor (AOP) and lube trim check valve. SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 79 -- 00 Bearing Lubrication The PW1100--JM engine rotors have frictionless (ball and roller) bearings. These bearings must have oil to lubricate them and to keep them cool.The FDGS uses journal bearings in which cylindrical bearing surfaces, isolated by a constant film of oil, turn against each other without balls or rollers. It is necessary to keep oil in these journal bearings at all times as a protection from friction and heat. A special pump system is used to keep sufficient oil pressure in these journal bearings during windmill and negative G conditions. FDGS Journal Bearing Auxiliary Oil Supply To keep sufficient oil flow to the FDGS journal bearings during windmill rotation, or negative G conditions, an auxiliary oil supply is used. An auxiliary oil tank is part of the No.1 bearing support, at the right side of the support structure. A gear on the rear of the FDGS gearbox shaft engages a two stage auxiliary pump. One stage supplies oil from the sump and the other stage supplies oil from the auxiliary oil tank. This pump supplies oil from either the auxiliary oil tank or the No. 1 and 1.5 bearing compartment sump, through the JOSV, back to the oil tank during engine operation. During windmill rotation, the JOSV supplies the windmill pump oil to the FDGS journal bearings to keep these bearings lubricated. If the engine and aircraft are on the ground and wind conditions cause the fan and LPC rotor to turn (windmill), the JOSV lets bearing compartment sump oil go to the windmill pump stage to the FDGS bearings through the JOSV. If the engine and aircraft are in flight in a zero or negative G condition, oil from the auxiliary oil tank goes to the auxiliary pump stage and from there to the JOSV and to the FDGS journal bearings. During engine operation, oil used by the FDGS flows out by centrifugal force to collect in an oil channel (”gutter”). This collected oil then flows to the auxiliary oil tank in the No.1 bearing support where it is available when necessary.Auxiliary oil tank oil drains down into the compartment sump at shutdown. FDGS Gear Oil Supply The FDGS is a planetary gear system in which a sun gear turns stationary star gears which then cause an outer ring gear (attached to the fan hub and gearbox shaft) to turn. The gears must have oil to lubricate them, but there are conditions in which increased oil flow to the gears is not necessary.The variable oil reduction valve (VORV) is FADEC controlled and valve position is primarily based on N2 speed and oil pressure. When scheduled by the FADEC the VORV controls oil flow to the FDGS and front bearing compartments. Page: 2


AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil System Schematic Page: 3


POWER PLANT OIL SYSTEM JOSV JOSV senses oil pressure to the FDGS journal bearing and when necessary (if main oil pressure to the journal bearings decreases) supplies auxiliary oil from the windmill pump to these bearings. Bearing Oil Damping The No.1, No. 2, No. 3, No. 4, and No. 6 bearings in the engine are oildamped. With these bearings, a layer of pressurized oil between the bearing outer race and its support absorbs rotor vibration and keeps engine vibration to a minimum. The No.3 bearing damping pressure is controlled by the AODV which is attached to the OCM. The AODV is FADEC controlled. Air Turbine Starter The Air Turbine Starter (ATS) on the PW1100--JM shares the engine oil so there are no scheduled ATS oil changes necessary.


Oil Breather Oil pressurization of bearing compartments causes the air pressure in these compartments to increase. In the Compressor Intermediate Case section, oil pressurization of the No. 3 bearing compartment and Angle Gearbox (AGB) causes an increase in air pressure which it is necessary to release. This released air pressure is breather. External tubes supply No. 3 and No. 5/6 bearing breather air (air mixed with oil vapor) to a deoiler unit in the main gearbox. The No. 5/6 bearing breather tubes contain a siphon break used for anti--coking features. Added to this bearing compartment breather air is breather from the main gearbox and from the oil tank. The deoiler removes the oil droplets from the air, and the air then goes overboard. Oil Sealing The engine controls oil in its bearing compartments with labyrinth seals (which control air pressure from adjacent engine areas) and carbon face seals (which control oil leakage). The labyrinth seals (knife edges on rotating parts that touch stationary lands on adjacent support structures) control airflow from higher pressure areas into a bearing compartment. Carbon face seals prevent oil leakage out of a bearing area. The seal plates in the No. 3 bearing area are of a liftoff type in which grooves in the seal plate surfaces put an air layer between the plate and the mating carbon seal. The result is less friction and leakage. SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 79 -- 00 Oil Filtration To keep oil as clean as possible, oil is put through a main oil filter to remove contamination and unwanted material. Note: The engine main oil filter has two stages, a 30--micron inner primary stage and a 150--micron outer secondary stage. Flow through the filter is reverse--type in which the oil flows from the inner (primary) element to the outer (secondary) element. If there is blockage of the primary stage, an oil filter bypass valve opens and supplies pressurized oil directly to the secondary filter. A differential (delta) pressure sensor is installed on the oil pump supplies a signal to the FADEC for maintenance action. There are also last chance screens at specified locations in the oil system to catch material that could get into the oil flow between the filter and (for example) a bearing compartment. The engine has chip collectors at specified positions for each bearing compartment which can collect metallic particles for analysis. Oil Tank To have a unit for oil storage and supply to the engine, an oil tank is installed on the left side of the engine on the fan case. The approximately 10.00 gallon (37.85 liter) oil tank supplies oil to the pressurization system and receives oil from the oil scavenge system. A deaerator in the tank removes air bubbles from oil before the oil goes into the tank. A pressure valve in the tank releases air/oil pressure to the deoiler in the gearbox when necessary. To make it possible to monitor oil level in the tank, there is a sight gage on the side and an oil quantity sensor. Oil Pressurization A lubrication and scavenge pump is installed on the left front of the main gearbox and is connected to the gearbox gear train. This pump has eight pump stages on two rotors. It uses six of these stages to remove scavenge oil from bearing compartments and the gearboxes and two to supply pressurized oil to the oil control manifold. The six scavenge pump stages pull scavenge oil from the No. 1, No. 1.5, No. 2 bearing, FDGS compartment, No. 3 bearing compartment, AGB, main gearbox, the No. 4 compartment, and the No. 5/6 bearing compartment. There are chip collectors internal to each scavenge compartment to catch metallic particles. Only the No. 4 bearing chip collector is installed in the oil return lines.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil System Components Page: 5


POWER PLANT OIL SYSTEM Oil Scavenge Scavenge oil from the No.1, No.1.5, No.2 bearings, and the FDGS bearings goes to the lubrication and scavenge pump and from there to the oil tank deaerator. The scavenge pump pulls scavenge oil from the No. 3, No. 4, No. 5/6 bearing compartments, AGB, and MGB, and supplies this scavenge oil to the oil tank also. All scavenge oil flows back to the oil tank through one main scavenge line. Between this main scavenge line and the oil tank is the Oil Debris Monitor (ODM). The ODM supplies ”real time” monitoring of oil for metallic deposits and is monitored by the FADEC. The oil tank pressure valve lets some of this air and oil vapor flow to the gearbox and overboard through the deoiler.

AIRBUS A-- 320 PW 1100 NEO 79 -- 00 An oil/oil heat exchanger is installed downstream of the air/oil heat exchanger and fuel/oil heat exchanger and uses engine oil to cool IDG oil. In some conditions, such as hot day ground idle, heat is transferred from the oil/oil heat exchanger to the engine oil. The oil/oil heat exchanger has a bypass valve which is mechanically operated.

Thermal Management The heat created in bearing compartments (and its removal by oil flow) makes oil hotter and makes it necessary to control this high oil temperature. This is done by heat exchangers which use available flows of air, fuel, or other oil to decrease oil temperature. A fuel/oil heat exchanger uses fuel to decrease oil temperature. An air/oil heat exchanger uses air, and an oil/oil heat exchanger uses cooler oil to remove heat from a hotter oil. Hot scavenge oil is supplied to the oil tank by the six scavenge chambers of the lubrication and scavenge oil pump. From the tank the oil goes to the pressure chamber of the pump and from there to the OCM. Pressurized oil is supplied to a sequence of heat exchangers. A fuel/oil heat exchanger uses fuel to make the oil cooler and an air/oil heat exchanger uses fan air to remove more heat from the oil. Note: A fuel/oil heat exchanger bypass modulating valve controls oil flow between the fuel/oil heat exchanger and the air/oil heat exchanger. Operation of this bypass valve is controlled by the FADEC (fuel and oil temperature data is supplied to the FADEC by the oil temperature sensor and the fuel temperature sensor). The fuel/oil heat exchanger has an internal bypass valve which is mechanically operated and lets oil flow around the cooler elements during cold starts (when it is possible that oil can be thicker) or if there is blockage of the heat exchanger. An air/oil heat exchanger is installed downstream of the fan discharge and uses this fan air to remove heat from the oil. The air/oil heat exchanger also has an internal bypass valve which is mechanically operated and lets oil flow around the cooler elements during cold starts or if is blockage the heat exchanger.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




JOSV Operation Page: 7



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL STORAGE SYSTEM Description The engine oil tank is the container for the engine’s supply of hot oil. The oil storage system is referred to as a hot tank oil system because hot scavenge oil is sent to the tank. The oil tank is attached to the fan case with a forward bracket and aft upper and lower bracket. The oil tank capacity is 8.75 gallons (33.12 liters). The oil tank has a filler cap which is locked into the top of the manual gravity fill opening. Oil Tank The oil tank is a line replaceable unit. The oil tank is located on the fan case at approximately the 8:30 o’clock position. The tank is made of aluminum with a formed heat shield that protects the fan case assembly from the hot oil. The heat shield is held on by four straps and also gives fire protection to the oil tank assembly. The heat shield is not a line replaceable unit. The oil tank has a swirl--type deaerator as part of the oil tank assembly. The deaerator is used to isolate air from the returning scavenge oil. The deaerator sends the air into the oil tank and then the tank releases the air through a pressurization valve (internal to the tank) to the gearbox. The pressurization valve keeps the tank at a minimum pressure of 12 psi (0.8 Bar) to help push the oil from the tank to the lubrication and scavenge oil pump. The oil tank servicing is done through the oil tank cap. To close the lock, turn the cap handle 45 degrees clockwise to the CLOSE position and press down on the handle. To release the lock feature, lift the handle and turn it 45 degrees CCW to the OPEN position. The cap handle has a pointer which indicates if the cap is in the opened or closed position. A flapper valve in the manual gravity fill opening is used to stop a large quantity of oil loss if the cap is not correctly installed. The scupper drain is used to drain oil spills overboard through the drain mast. The drain plug is on the bottom of the tank.


Oil Tank Sight Gauge The sight gauge gives a visual indication of the oil level. The sight gauge is a line replaceable unit. Oil Quantity Sensor The oil quantity sensor is on the top of the oil tank. The sensor is used for flight deck indication of the quantity of the oil in the tank. The oil quantity sensor is a line replaceable unit.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil Tank Page: 9



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL DISTRIBUTION SYSTEM Description Oil system is used to supply pressurized oil to lower the temperature, lubricate, and clean the engine’s main bearings, seals, gears, splines, and accessory drives, and to heat the engine fuel to prevent ice in the fuel lines. An auxiliary oil supply subsystem is used to keep continuous oil flow to the FDG when there are periods of low main oil pressure because of maneuvers or when windmill operation occurs, the engine fuel/oil heat exchanger bypass valve is located on the fuel/oil manifold at the 10 o’clock. The No. 5 and 6 bearing magnetic chip collector is located on the lubrication and scavenge oil pump at the 8 o’clock. The No. 4 bearing magnetic chip collector is located in the tube (LR40) near the High Pressure Turbine (HPT) case at the 7 o’clock. The main gearbox magnetic chip collector is located on the lubrication and scavenge oil pump at the 8 o’clock. The angle gearbox magnetic chip collector is located on the lubrication and scavenge oil pump at the 8 o’clock. The No. 3 bearing magnetic chip collector is located on the lubrication and scavenge oil pump at the 8 o’clock. The FDG system magnetic chip collector is located on the lubrication and scavenge oil pump at the 8 o’clock. The lube trim check valve is located in the OCM at the 9 o’clock. The Variable Oil Reduction Valve (VORV) is attached to the OCM on the left of the engine at the 9 o’clock. The Active Oil Damper Valve (AODV) is attached to the OCM at the 9 o’clock. The Integrated Drive Generator (IDG) oil/oil heat exchanger is attached to the Thermal Management System (TMS) Manifold at the 8 o’clock. The air/oil heat exchanger is located on the diffuser case at the 10 o’clock. The engine fuel/oil heat exchanger is located on the TMS at the 8 o’clock. The lube and scavenge oil pump is located on the MGB at the 7 o’clock. The oil filter is located on the lub and scavenge oil pump at the 8 o’clock. The deoiler is located on the rear of the MGB at the 8 o’clock.


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Oil Distribution System Page: 11



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

Operation Oil flows from the oil tank to the lubrication pressure area of the lubrication and scavenge oil pump. Oil tank pressure of approximately 12 psi (0.8 Bar) is kept to increase the inlet pressure to the pump for good pump section and to prevent pump cavitation. The lubrication and scavenge oil pump is used to pressurize the oil and move it through the main oil filter. From the main oil filter, the filtered oil flow is split by a metering orifice and flows to the fuel/oil heat exchanger bypass valve or to the system’s main supply point (uncooled). The fuel/oil heat exchanger bypass valve divides the oil flow between the engine fuel/oil heat exchanger and air/oil heat exchanger. The heat exchangers control the supply oil temperature to the engine and components. After the pressurized oil temperature is decreased, part of the oil flow is supplied to the FDG journal bearings. The remaining cool oil is mixed with the uncooled oil by a metering orifice and is supplied to the areas that follow: FDG gears by the VORV No. 1, 1.5, and 2 Bearing Compartments by the VORV No. 3 Bearing Compartment No. 4 Bearing Compartment No. 5 and 6 Bearing Compartments Main Gearbox (MGB) and Angle Gearbox (AGB) Nozzles in the main bearing compartments and gearboxes supply the oil (at the correct flow rates) to the different bearings, seals, gears, and accessory drive splines. Last chance strainers are supplied at the entrances to the mpartments e oil distribution system downstream of the main oil filter out of the oil nozzles.

A mechanically activated static anti--leak valve to prevent back--flow during shutdown In normal operation, the oil flows from the lubrication and scavenge oil pump and into the outer part of the filter to the inner part through the primary and secondary filter elements. The element is also a dual element design. The dual element oil filter is a filter in a filter design, with a primary filter bypass valve installed in the filter assembly. Because the filter function is redundant (primary and secondary filters) the filter function cannot fail. The dual element design has a coarse filtration secondary element in the finer filtration primary element that will continue to supply a level of protection to the engine. The oil filter differential pressure sensor gives an indication if the primary oil filter has clogged. The differential pressure sensor is installed on the lubrication and scavenge oil pump, adjacent to the oil filter housing. Two pressure connections, one upstream and one downstream of the main oil filter element, supply the pressures used to monitor differential oil pressure. The main oil differential pressure sensor is connected to the two pressure connections. When the differential pressure across the filter is more than a specified limit, a maintenance flag is set. When the differential pressure across the primary oil filter element is more than a specified limit, pressurized oil is sent directly to the secondary filter. An oil filter bypass signal is also sent to the flight crew through the Electronic Engine Control (EEC).

Main Oil Filter Element The main oil filter element removes solid contaminants from the pressurized oil sent from the lubrication and scavenge oil pump. The filtered oil goes from the main oil filter to the Oil Control Module (OCM). The main oil filter element is in the main oil filter housing attached to the lubrication and scavenge oil pump. The main oil filter housing contains: A dual element (primary and secondary) oil filter assembly A primary oil filter bypass valve A check valve (shut--off valve) to prevent oil leakage during filter service

Oil Control Module (OCM). The OCM gets pressurized, filtered oil from the lubrication and scavenge oil pump and distributes the oil to different engine compartments and heat exchangers through internal cored passages. The OCM greatly reduces the number of external oil lines necessary to supply pressurized oil. The OCM is located adjacent to the MGB.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil System Component Location Page: 13


POWER PLANT OIL SYSTEM Engine Air/Oil Heat Exchanger The engine air/oil heat exchanger cools the engine oil with fan bypass air. The engine oil that is cooled by the air/oil heat exchanger decreases the quantity of heat that must be moved from the oil to the fuel in the fuel/oil heat exchanger. The engine air/oil heat exchanger is located at flange H on the diffuser case at the 11:00 o’clock position. The engine air/oil heat exchanger, which receives pressurized oil from the fuel/ oil heat exchanger bypass valve, has plates through which the engine oil flows. There are fins brazed onto the plates with passages. The fins are used to move heat from the plate walls to the fan bleed air that flows between the plates. The heated air then flows to the nacelle area. An open bypass valve lets the oil bypass the air/oil heat exchanger when the oil is cold or the heat exchanger core is clogged. The EEC controls the flow of engine oil to the air/oil heat exchanger by commanded input to the fuel/oil heat exchanger bypass valve. It increases the oil flow if the fuel temperature is more than a specified value. The increased oil flow decreases the oil flow through the fuel/oil heat exchanger which decreases the quantity of heat transfer to the fuel in the fuel/oil heat exchanger. Fuel/Oil Heat Exchanger The fuel/oil heat exchanger moves heat from the engine oil to the engine fuel. The fuel/oil heat exchanger is attached to the High Pressure Turbine (HPT) case at the 9:00 o’clock position. The heat exchanger has stacked fuel and oil plates which let the heat transfer. The heated fuel flows through the fuel temperature sensor. The cooled engine oil that flows through the Thermal Management System (TMS) manifold lubricates and decreases the temperature of the engine bearings, gears, and accessory drives. An oil bypass valve causes oil to bypass the fuel/oil heat exchanger if there is a clog. Integrated Drive Generator (IDG) Oil/Oil Heat Exchanger The IDG oil/oil heat exchanger moves heat from the IDG oil to the engine oil from the engine air/oil heat exchanger outlet. The IDG oil/oil heat exchanger is connected to the TMS manifold.


AIRBUS A-- 320 PW 1100 NEO 79 -- 00 The heat exchanger, which receives cooled, pressurized oil from the engine air/oil heat exchanger, has plates through which the IDG oil flows. There are fins attached to the plates. The fins move heat from the plate walls to the engine oil that flows between the plates. The engine oil at increased temperature then flows downstream to mix with engine oil discharged from the engine fuel/oil heat exchanger. A bypass valve lets the oil bypass the heat exchanger when the oil is cold or the heat exchanger is clogged. The IDG oil/oil heat exchanger is a line replaceable unit. Fuel/Oil Heat Exchanger Bypass Valve. The fuel/oil heat exchanger bypass valve is an electromechanical device that variably splits the supply flow from the fuel/oil heat exchanger to the air/oil heat exchanger to control the temperature of the oil and fuel. The valve is located on the OCM. The fail--safe position is 92.5% flow to the fuel/oil heat exchanger and 7.5% to the air/oil heat exchanger. The valve is fully modulated with no position feedback necessary to the EEC, which commands the valve. The fuel/oil heat exchanger bypass valve is a line replaceable unit.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil System Heat exchanger Page: 15


POWER PLANT OIL SYSTEM VORV The VORV is an electromechanical device that regulates the flow of oil to the No. 1, No. 1.5, No. 2, and FDG for different flight conditions. Full oil flow to lubricate the gear faces is only necessary at high engine power conditions. At cruise, a small quantity of direct oil flow is necessary and oil bypasses the loads to decrease the oil heat load and increase gearbox efficiency. Valve position is directly related to N2 speed and oil pressure. The valve is fully modulated by a servo valve controlled by the EEC on channels A and B. Feedback is given by a Linear Variable Differential Transducer (LVDT). The fail--safe valve position is maximum oil flow. The VORV includes a second valve contained in the same housing called the Journal Oil Shuttle Valve (JOSV). The VORV is a line replaceable unit.


AODV The AODV controls oil supply flow to the No. 3 bearing oil damper for high rotor (N2) vibration control. The AODV is located on the OCM. The valve is made to let oil flow to the damper at engine ignition and shuts the oil supply off at high power. The fail--safe position for the valve is open. The AODV is a line replaceable unit. Last Chance Oil Strainers Last chance oil strainers are used to keep large particles from the bearing compartments and prevent blockage of the oil nozzles. The strainers keep contamination away from the oil nozzles that gets into the system downstream of the main oil filter. The metal mesh last chance oil strainers are externally mounted on the engine. * The FDG and No. 1, 1.5, and 2 bearing oil strainer is installed at the VORV manifold outlet. * The two No.3 bearing and bevel gear oil strainers are located in the oil pressure tubes that go to the No. 3 bearing compartment on the Compressor Intermediate Case (CIC).


AIRBUS A-- 320 PW 1100 NEO 79 -- 00 * The No. 4 bearing oil strainer is in the lubrication pressure tube that goes to the No. 4 bearing compartment. * The No. 5 and 6 bearing oil strainer is located in the lubrication pressure tube that goes to the No. 5 and 6 bearing compartment. * The gearbox oil strainer is located in the OCM housing. The last chance oil strainers are line replaceable units. Journal Oil Shuttle Valve (JOSV) The JOSV is a mechanical device that continuously supplies oil to the fan drive journal bearings from the main oil supply or the auxiliary oil supply. The JOSV operates independently of the VORV but is located in the same housing. In normal conditions, the JOSV sends oil from the main oil pump to the journal bearings and scavenge oil from the No. 1 and No. 1.5 bearing support to the OCM. In windmill and zero or negative gravity (G) conditions, scavenge oil from the No. 1 and No. 1.5 bearing support is sent back to the journal bearings. The JOSV is not controlled by the EEC and gives no feedback. Valve position changes with oil pressure. Description of the Engine Oil Scavenge System The engine oil scavenge system sends back the hot lubrication oil to the oil tank through the lubrication and scavenge oil pump. The oil that is sent back is used by the engine oil distribution system to lubricate and decrease the temperature of the engine’s main bearings and gearboxes. The lubrication and scavenge oil pump has six scavenge pump stages that are used to pull oil from the: * Front bearing compartment * No. 3 bearing compartment * No. 4 bearing compartment * No. 5 and 6 bearing compartment * Main gearbox * Angle gearbox The six scavenge pump stages send the scavenge oil to the deaerator in the oil tank.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




VORV - AODV and Fuel /Oil Heat Exchanger Page: 17


POWER PLANT OIL SYSTEM The oil tank deaerator removes the air from the scavenge oil and lets the hot, deaerated scavenge oil flow into the oil tank. The engine oil scavenge system has six magnetic chip collectors that catch ferrous metal particles in the scavenge oil. Scavenge oil pump housing has five chip collectors -- one for each of the scavenge oil lines that go into the pump. These lines cycle scavenge oil from: * The front bearing compartment (No. 1, 1.5, 2, and FDG) * No. 3 bearing compartment * No. 5 and 6 bearing compartment * Main gearbox * Angle gearbox The No. 4 bearing compartment chip collector is located on the No. 4 bearing scavenge line. The scavenge oil flows around the magnetic chip collectors and is pumped to the deaerator in the oil tank. The six chip collectors are bayonet--type plugs that can be removed and examined at regular time intervals or on condition. The collectors are in housings with self--closing check valves, so there is no oil leakage when a chip collector is removed. The six magnetic chip collectors are line replaceable units.


Engine Oil Breather System and Components The engine oil breather system removes air from the bearing compartments, isolates the breather air from the oil, and releases the air overboard. Deoiler The breather air from the oil tank deaerator, the No. 3 bearing compartment, and the No. 5 and 6 bearing compartment flows to the deoiler which isolates the scavenge oil from the air. The oil drains into the MGB and mixes with the gearbox scavenge oil. Breather air from the deoiler goes through the deoiler air exit duct to be sent overboard. The air/oil scavenge mixture from the front bearing compartment (No. 1, 1.5, 2, and FDG), and No. 3, 4, and 5 and 6 bearing compartments flows to the lubrication and scavenge oil pump. The pump sends the mixture to the oil tank deaerator which isolates the breather air from the scavenged oil. The oil flows into the oil tank. The breather air flows through the oil tank pressurization valve and goes to the MGB. The MGB breather then sends air/oil.


AIRBUS A-- 320 PW 1100 NEO 79 -- 00 Deoiler Air Exit Duct The deoiler, which is on the MGB, sends isolated breather air overboard through the deoiler air exit duct. The air exit duct is attached to the rear of the deoiler housing with bolts. The deoiler air exit duct is a line replaceable unit. Auxiliary Oil System An auxiliary lubrication system is included to make sure that the journal bearings will get oil while a flight or windmill condition occurs. System Components Fan Oil Pump JOSV Auxiliary Oil Reservoir FDG Gutter Gravity Valve Operation This system has a single function fan oil pump located in the front bearing compartment. It is operated by the fan rotor that supplies pressure to the JOSV which usually will bypass the oil back to the oil tank. If the JOSV senses a low pressure (when a windmill operation or a zero or negative G event occurs), it will send this oil to the journal bearings to make sure there is sufficient lubrication of the journal bearings. The pump will get oil from the compartment sump or an isolated auxiliary reservoir through the gravity valve. The gravity valve has a small weight which is sensitive to pressure differentials and the gravity vector. For zero or negative G conditions, the pump will get oil through the gravity valve from the isolated auxiliary reservoir. This reservoir is located internally in the No. 1 and 1.5 bearing compartment adjacent to the gear assembly. To make sure the reservoir is always full, the total oil flow for the gear system is thrown off the ring gear flange into a 360 degree FDG gutter. The high tangential velocity forces the oil into the reservoir. All the oil from the gear is not necessary to make sure the journal bearings are sufficiently lubricated, so the unwanted oil is bypassed to the compartment sump. For ground windmill operations, the pump will get oil through the gravity valve from the compartment sump, which is continuously filled by the recycled oil that moves from the gear system and main shaft bearings. The small quantity of oil supplied in ground windmill operation will prevent journal bearing damage. Page: 18


AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Lub Trim and DeOiler Page: 19

AIRBUS A-- 320 PW 1100 NEO 79 -- 00

PRESSURIZATION AND SCAVENGE Lubrication and Scavenge Oil Pump The lubrication and scavenge oil pump sends pressurized oil from the oil tank to the engine bearings, seals, gears, and accessory drives. The pump also sends scavenge oil back to the oil tank. The lubrication and scavenge oil pump connects to the front of the main gearbox at the 7:00 o’clock position. The pump has seven positive displacement gear--type pump stages. Each stage is turned by the MGB at a speed proportional to N2. The seven pump stages are: A lubrication main pressure stage that sends oil to the main oil filter Six scavenge stages that scavenge oil from the engine compartments that follow, and send it to the oil tank deaerator: FDG and No. 1, 1.5, and 2 bearings No .3 bearing Angle gearbox Main gearbox and deoiler No. 4 bearing No. 5 and 6 bearings The lubrication and scavenge oil pump has chip collectors that remove metal particles from the scavenge oil to find problems with oil--lubricated components. The collectors are found at the inlets of five scavenge stages that receive oil through external tubes. The lubrication and scavenge oil pump has a relief valve that limits the maximum main oil pressure. The main oil filter housing is included in the lubrication and scavenge oil pump housing.





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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Lube and Scavenge Pump Page: 21



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

SCAVENGE SYSTEM General The scavenge system pumps the oil from the bearing compartments and gearboxes back to the oil tank. The system consists of a Lube and Oil Scavenge Pump (LSOP) .The pump has six stages that return oil from the areas listed below. * Front bearing compartment servicing the FDGS and bearing nos.1, 1.5 and 2. * No. 3 Bearing compartment * No. 4 Bearing compartment * Compartment for bearing nos. 5 and 6 * Main Gearbox * Angle Gearbox The stages send scavenged oil to the oil tank, where a deaerator separates the air that has mixed with the oil. Air that is separated from the oil pressurizes the oil tank. Note that the Main Gearbox requires no external scavenge line. The system also has six magnetic chip collectors. Magnetic Chip Collectors The Lubrication and Scavenge Oil Pump has six magnetic chip collectors. The collectors catch ferrous metal particles in the scavenge oil which are used to diagnose system problems.


Location Five of the collectors are located on the LSOP and a sixth is found on the No.4 Bearing scavenge return tube. Description The collector assembly consists of a collector probe and probe housing. The probe housing has a spring--loaded check valve so there is no leakage when a detector is removed. The six chip collectors are bayonet--type plugs that can be removed and examined at regular intervals or on--condition. Operation When the probe is inserted to the housing, the check valve is forced open, exposing the probe tip to the oil flow. When the probe is removed for inspection the check valve closes, preventing oil from leaking out of the system.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Scavenge and Chip Detectors Page: 23



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

LAST CHANCE STRAINER General Large particles can enter the oil supply beyond the main oil filter on the Oil Control Module. Last chance oil stariners prevent these particles fron entering bearing compartments and cloggin oil nozzles.


Description Strainers are metal mesh type that fit inside the oilk pressure supply tubes and are referred to as in line strainers. They can be removed, inspected, cleaned and replaced during line maintenance, but this is not required under normal engine operation. The most of the last chance strainers are located on the left side except for the No 4 bearing strainer wich is located at the right side eight o’clock.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Last Chance Strainer Page: 25



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

AUXILIARY LUBRICATION SYSTEM General The Auxiliary Lubrication System protects the FDGS journal bearings from low oil pressure conditions that could cause loss of oil. These include windmilling operation (in flight or on ground) and zero gravity or negative gravity events. The system is located on the support housing for bearing nos.1/1.5 and consists of the following components: * fan drive gear train * windmill/auxiliary pump * sprag clutch gear assembly * auxiliary reservoir. The dedicated windmill/auxiliary dual--stage fan oil pump is located in the front bearing compartment and driven by the fan rotor. The pump continuously draws oil from a dedicated auxiliary reservoir and compartment sump located in the front bearing compartment. The auxiliary reservoir is part of the casting of the support for bearing nos.1/1.5. In normal conditions the pump sends the reservoir and sump oil to the Journal Oil Shuttle Valve (JOSV), which directs the oil back to the oil tank. In low pressure conditions the JOSV directs the oil to journal bearings, ensuring their lubrication. During zero or negative gravity events the pump draws oil from the auxiliary reservoir, which is continuously replenished by oil slung from the gear system into the reservoir. During windmill operations, the pump draws oil through the compartment sump, which is continually replenished by oil cast off from the gear system and main shaft bearings. Fan Drive Gear Train The Fan Drive Gear Train is a system of gears that ensures the windmill/ auxiliary pump is protected from reverse windmill conditions. The fan drive gear train is located inside the support for bearing nos.1/1.5 and on the fan shaft. The fan drive gear train connects the windmill/auxiliary pump to the fan shaft. Whenever the fan is turning, the windmill/auxilliary pump is sending oil to the FDGS journal bearing operation. Torque transfers from the fan shaft gear through the fan drive gear train. Torque is transferred next to the windmill/ auxiliary pump gear and to the pump itself through a splined shaft.


Windmill/Auxiliary Pump The windmill/auxiliary pump is a dedicated dual stage fan oil pump that supplies oil to journal bearings during low pressure oil conditions. The pump is located on the support for bearing nos.1 and 1.5. The pump consists of two stages with separate feeds and a common discharge. One stage feed is connected to the auxiliary oil reservoir in the support for bearing nos.1 and 1.5. The other stage feed is connected to the oil in the bottom, or sump, of the bearing compartment. Operation See the table at right for a summary of oil flow in various operating conditions. In normal conditions when the Lubrication and Scavenge Oil Pump pressure supplies the FDG bearings sufficiently, oil from the windmill/auxiliary pump is directed back to the main oil tank by the JOSV. In low oil pressure conditions, such as zero gravity or negative gravity events, the pump stage connected to the auxiliary reservoir pumps oil to the journal bearings through the JOSV. Under windmill conditions, oil from the sump stage is pumped by the windmill/ auxiliary pump to the journal bearings through the JOSV. Sprag Clutch Gear Assembly The sprag clutch gear assembly keeps the auxiliary oil pump gear turning in the same direction during windmill conditions, regardless of the direction in which the fan shaft turns. The assembly is mounted at 7:00 inside the 1/1.5 bearing support housing. The sprag gear clutch assembly consists of a steel housing, two sprag clutch bearings and one roller bearing. All three bearings are pressed onto shafts turning spur gears that mesh with each other. Auxiliary Oil Reservoir Oil in this reservoir is directed to journal bearings during zero gravity or negative gravity events. The reservoir is located in the compartment for bearing nos. 1 and 1.5. Centrifugal action of the ring gear set drives oil into the gutter and then into the auxiliary oil reservoir. Operation Oil is directed out of the reservoir to the auxillary oil pump and to the journal bearings through passageways cast into the bearing support.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Auxiliary Lubrication System Page: 27


POWER PLANT OIL SYSTEM BREATHER SYSTEM General During engine operation, sealing air flows into the bearing compartments. The sealing air must be vented to allow a continuous flow. The sealing air that vents is referred to as breather air. The Breather System removes air from the bearing compartments, separates the breather air from the oil, and vents the air overboard. Components are shown below. * Deoiler * Deoiler vent duct * No.3 Bearing breather vent tube * Main oil tank deaerator vent tube * Anti--siphon tube for bearing nos.5 and 6 Deoiler and Deoiler Vent Duct The deoiler separates air from scavenge oil. The deoiler is integral to the Main Gearbox and is located on the left side. Description Torque is applied to the deoiler rotor drive gear from the MGB. The rotor captures oil mist residing in the MGB, and through centrifugal action the oil is separated from the air. The oil--soaked breather air is vented into the MGB from the compartments for bearing nos. 5, 6 and 3, and from the oil tank. The separated oil flows into the MGB sump to the LSOP, and the air flows out of the MGB through the deoiler vent duct.


AIRBUS A-- 320 PW 1100 NEO 79 -- 00 Main Oil Tank Deaerator Vent Tube The main oil tank deaerator vent tube vents tank pressure greater than 12 psi from the main oil tank. The tube is connected to the oil tank deaerator and the Main Gearbox. Description Sealing air for all bearing compartments excluding bearing nos. 5 and 6 mixes with scavenge oil and flows back to the oil tank. A static deaerator in the oil tank separates the oil from the air. The released air pressurizes the oil tank. Operation Pressure in the tank is controlled by a spring--loaded closed, mechanical poppet valve. The valve opens to release excess pressure in the tank and sends the excess air/oil mist to the deoiler that is internal to the Main Gearbox. Breather air from the Main Gearbox flows internally to the deoiler vent tube. Anti--Siphon Tube for Bearing Nos. 5, 6 The anti--siphon tube allows some oil to remain present in the pressure tube for bearing nos. 5 and 6 after the engine is shutdown. The anti--siphon tube is on the left side of the engine core. It attaches to the oil pressure “T” fitting for bearing nos. 5 and 6, and to the No. 3 Bearing breather tube. The effect of allowing oil to remain in the anti--siphon tube after shutdown prevents coking in the bearing pressure tube.

No. 3 Bearing Breather Vent Tube The No.3 Bearing breather vent tube sends breather air directly from the bearing compartment to the deoiler in the Main Gearbox. The vent tube is attached to the CIC at 10:00 and to the rear of the Main Gearbox. Description Air enters the No. 3 Bearing compartment from between the carbon seal and face seal, flowing through the compartment and removing heat. The airflow carries some of the oil that is used in the bearing compartment, lubricating the bearings in the form of a mist. The No. 3 Bearing vent tube vents this breather air directly to the deoiler in the Main Gearbox. SCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00

DEOILER


DEOILER VENT DUCT

MAIN GEARBOX BREATHER AIR


Breather System Page: 29

AIRBUS A-- 320 PW 1100 NEO 79 -- 00

INDICATING SYSTEM General The Indicating System monitors Lubrication System conditions and alerts the flight crew to potential problems. Components in the system send signals to the Electronic Engine Control (EEC), which in turn notifies the flight deck Electronic Centralized Aircraft Monitoring System (ECAM). Sensors are listed below. * Oil Level OLS * Oil Filter Differential Pressure OFDPS * Oil Debris Monitor ODM * Main Oil Temperature MOT * Main Oil Pressure MOP * Low Oil Pressure LOPS * Auxiliary Oil Pressure AOPS





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AIRBUS A-- 320 PW 1100 NEO 79 -- 00


B

B Figure 15 SCL /JGB / REV.00 / Jun--2016

Page: 31


POWER PLANT OIL SYSTEM Oil Level Sensor (OLS) The Oil Level Sensor indicates the oil level within the oil tank. The sensor is internal to the oil tank. The sensor is a single channel transducer with a magnetic float and reed switch configuration. A hollow tube is welded to the top mounting plate and has an integral bottom mounting flange that fits into a mating flange inside the bottom of the oil tank assembly. The hollow tube contains a magnetic ball float and a circuit board. The length of the circuit board contains a series of switches. A single electrical connector is attached to the top mounting plate. The mounting plate is secured to the top of the oil tank assembly with three bolts. An O--ring beneath the mounting plate prevents leakage. The OLS must be replaced if it is removed on--wing or during a shop visit.

AIRBUS A-- 320 PW 1100 NEO 79 -- 00 Unit (PHMU) to differentiate between a ferrous and non--ferrous particle. Ferrous material passing through the electromagnetic field strengthens the field, and non--ferrous material passing through weakens the field. This effect creates two unique signatures used by the PHMU to differentiate the types of particles. The PHMU processes the signal from the ODM and issues a chip generation rate (the number of chips counted in a given period of time). The chip generation rate signal is then sent to the EEC, where the rate is compared to predetermined values and the appropriate maintenance message or cockpit signal is sent to the EIU.


Operation The magnetic field produced by the magnetic float closes and opens each switch as it passes them while floating on the oil surface. The sensor then outputs a single channel signal to the EEC using a DC voltage that correlates to the oil level in the tank. Oil Debris Monitor (ODM) The Oil Debris Monitor (ODM) detects and measures metallic debris in the Lubrication System. The ODM is installed between the main oil scavenge line and the deaerator in the oil tank assembly. The ODM is a single channel, in--line sensor that is non--repairable. It consists of a sensing element, a stainless steel body that shields the sensing element from damage, a mounting flange, and an electrical connector. O’rings prevent oil from entering the unit. The ODM is secured by three bolts which also secure the oil scavenge line to the deaerator. An O’ring provides oil sealing at the deaerator interface and a face seal provides oil sealing at the oil scavenge line interface. Both the O’ring and face seal must be replaced if the ODM is removed on--wing or during a shop visit. Operation The ODM creates an electromagnetic field through which scavenge oil flows. When metallic particles are present in the scavenge oil, the sensing element produces a characteristic signal. The amplitude of the signal is proportional to the particle size. Its phase allows the signal processing electronics of the Prognostics Health and Management SCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil Level Sensor and Debris Monitor Page: 33



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

Oil Filter Differential Pressure Sensor (OFDPS) The Oil Filter Differential Pressure Sensor measures the difference in oil pressure upstream and downstream of the oil filter. The OFDPS is secured to the Lubrication and Scavenge Oil Pump. Two bolts secure the OFDPS to the LSOP. Two O--rings are installed beneath the mounting flange to prevent oil leakage. The Orings must be replaced If the OFDPS is removed on--wing or during a shop visit. The dual channel sensor consists of an electrical connector and two independent, electrically isolated sensing elements contained within a sealed stainless steel body that protects the sensing elements from damage. Each sensing element consists of a diaphragm with strain gages bonded to the surface and is connected to the electrical connector. Operation When pressure is applied, the strain gages change resistance altering the output voltage. This output voltage for each sensing element correlates directly to oil differential pressure and is sent to the Electronic Engine Control (EEC). The EEC uses this signal to set a maintenance message or an “oil filter clogged” message, depending on the differential pressure value.


Main Oil Temperature (MOT) Sensor The Main Oil Temperature (MOT) Sensor measures the temperature of the oil. The sensor is installed on the OCM. The dual--channel sensor consists of two independent, electrically isolated sensing elements; a stainless steel body with mounting flange and a protective tube that shields the sensing elements from damage; and one electrical connector. Components are assembled as a hermitically sealed unit that is non--repairable. Oil temperature is measured by two independent Resistance Temperature Detector (RTD) sensing elements. The MOT sensor is secured to the OCM with two bolts. An O’ring is installed beneath the mounting flange to prevent oil leakage. The O’ring must be replaced if the MOT sensor is removed on--wing or during a shop visit. Operation As the temperature of the sensing element changes, the electrical resistance alters, causing the voltage across the element to change proportionally. Each sensing element is connected to a single electrical connector and sends the oil temperature signal (voltage) to the EEC over separate channels A and B. Both channels share the same electrical connector.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL PRESSURE


ELECTRICAL CONNECTOR J35

MGB


Main Oil Temp Sensor and Oil Differential Pressure Sensor Page: 35


POWER PLANT OIL SYSTEM Main Oil Pressure (MOP) Sensor The Main Oil Pressure sensor measures oil pressure on the supply side of the Lubrication System. The sensor is installed on the Oil Control Module. The dual--channel sensor consists of two independent, electrically isolated sensing elements; a stainless steel body with a mounting flange, and which shields the sensing elements from damage; and one electrical connector. Each sensing element consists of a diaphragm with strain gages bonded to the surface. The MOP sensor is secured to the Oil Control Module (OCM) with two bolts. An O’ring is installed beneath the mounting flange to prevent oil leakage. The O’ring must be replaced if the MOP sensor is removed on--wing or during a shop visit.

Low Oil Pressure Sensor (LOPS) The Low Oil Pressure Sensor sends a low oil pressure signal directly to the Engine Interface Unit (EIU) when oil pressure has been reduced to a level below which engine operation is not recommended. The LOPS is mounted to the bottom of the OCM on the supply side of the Lubrication System downstream of the MOP sensor. The LOPS consists of a diaphragm sensing element, a spring, a mechanical switch, a stainless steel body, and one electrical connector. The stainless steel body, chosen for its strength and resistance to corrosion, has a mounting flange and shields the internal components from damage. An O’ring is installed beneath the mounting flange to prevent oil leakage. The O’ring must be replaced if the LOPS is removed on--wing or during a shop visit. Operation The mechanical switch is always in the closed (actuated) position during engine operation and when oil pressure is applied to the diaphragm sensing element. However, when the engine is operating and the applied oil pressure decreases below a predetermined design value, the spring force on the diaphragm is greater than the applied oil pressure on the diaphragm--sensing element. This allows the spring to displace the diaphragm and open the mechanical switch to the deactuated position. The low oil pressure electrical signal is then sent to the EIU, bypassing the EEC.


Operation When pressure is applied, the strain gages change resistance, altering the output voltage. This output voltage correlates directly to oil pressure. Each sensing element is connected to a single electrical connector and sends the oil pressure signal to the Electronic Engine Control (EEC) over separate channels A and B. Both channels share the same electrical connector.

AIRBUS A-- 320 PW 1100 NEO 79 -- 00


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL PRESSURE



OIL PRESSURE



Main Oil Pressure and Low Oil Pressure Sensor Page: 37



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

Auxiliary Oil Pressure Sensor (AOPS) The Auxiliary Oil Pressure Sensor detects latent failures in the Journal Oil Shuttle Valve or the windmill/auxiliary pump. The AOP Sensor is secured to the manifold for the Variable Oil Reduction Valve and the Journal Oil Shuttle Valve. The dual--channel sensor measures the pressure of the oil being delivered to the journal bearings in the fan drive gearbox under normal, windmill, and negative--G conditions. The measurement detects latent failures in the JOSV or the windmill/auxiliary pump. The sensor consists of two independent, electrically isolated sensing elements, one electrical connector, and a stainless steel body with a mounting flange, which shields the sensing elements from damage. The sensor is secured with two bolts to the VORV/JOSV manifold. The O’ring must be replaced if the sensor is removed on--wing or during a shop visit.


Operation Each sensing element consists of a diaphragm with strain gages bonded to the surface. When pressure is applied, the strain gages change resistance, altering the output voltage. This output voltage correlates directly to oil pressure. ach sensing element is connected to a single electrical connector and sends the oil pressure signal to the EEC over separate channels A and B. Both channels share the same electrical connector.


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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Auxiliary Pressure Sensor Page: 39


POWER PLANT OIL SYSTEM INDICATING SYSTEM Flight Deck Display The flight deck ECAM Secondary Engine Parameters displays Lubrication System conditions for both engines, using three separate display pages depending on the type of information, Engine Indicating, Engine Status and Maintenance Mode. The table gives details about the display for each page. PAGE Secondaryy Engine g Page g

Status Page g

MCDU


AIRBUS A-- 320 PW 1100 NEO 79 -- 00

DISPLAY TYPE

DISPLAY CONDITION

Oil Pressure Temperature p and Quantity

Engine g is in operation p

Details about the cause of the ECAM message g

Must be selected byy the crew in order to be viewed

Faults for oil pressure temperature temperature and quantity

Can only be displayed in interactive mode

Oil pressure displayed on the Secondary Engine page is measured in pounds per square inch (PSI) and temperature in degrees Centigrade (C). These parameters will change color on the display if they start to go outside of the normal range. Green Normal range Green pulsing If pressure exceeds OIL HIGH PRESSURE ADVISORY Amber Approaching red line limit that signall low oil temperature Red Red line limit Oil quantity displayed on the Secondary Engine page is measured in quarts (QTS). The display changes color depending on condition. Green Normal range Green pulsing If quantity drops below advisory level Amber Quantity below limit Red Red line limit SCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil Indicating System Page: 41


AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Secondary Engine Display Oil Parameters Page: 42


AIRBUS A-- 320 PW 1100 NEO 79 -- 00




ECAM and MCDU Message Page: 43



AIRBUS A-- 320 PW 1100 NEO 79 -- 00

OIL SERVICE Oil Tank Procedure To check and refill the engine oil to the correct level, open the oil tank access door on the left fan cowl door at approximately 9:00. If the oil level in the sight glass is below the FULL mark, replenish the oil as follows. 1. Lift up the T--handle on the oil tank cap and open the cap. 2. Put a fluid drain collector/container (approximately 5 gal/(20 L) under the end of the scupper drain line under the engine. 3. If necessary, insert a small screw driver through one of the 0.25 inch (6.35 mm) holes in the oil tank inlet screens and open the oil tank flapper valve. 4. Continue to hold the flapper valve open and add the correct engine oil into the filler neck until no more oil can be added without overflow into the scupper drain. 5. The stable oil level in the sight glass will now be in the full mark range. Note: if the aircraft is not parked on a level, the oil level indication in the sight glass is affected and may be above the full mark. This is acceptable. The oil system is serviced correctly if no more oil can be added without overflow into the scupper drain. 6. Close the cap to seat and push down on the T--handle until the locking pin fully engages the lock pin hole. 7. Close the cap to seat and push down on the T--handle until the locking pin fully engages the lock pin hole. Note: The handle will be parallel to the cap when the pin is completely engaged. 8. Record the amount of oil you added. 9. Wipe clean the area with a lint--free cotton cloth. 10. Close the oil tank access door on the left.


WARNING:

BE CAREFUL WHEN YOU WORK ON THE ENGINE AFTER SHUTDOWN, THE ENGINE AND ENGINE OIL CAN STAY HOT FOR A LONG TIME. IF YOU DO NOT OBEY THIS WARNING, INJURY CAN OCCUR. REFER TO THE MSDS FOR ALL MATERIAL USED AND THE MANUFACTURER S SAFETY INSTRUCTIONS FOR EQUIPMENT USED. IF YOU DO NOT OBEY THIS WARNING, INJURY CAN OCCUR.

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AIRBUS A-- 320 PW 1100 NEO 79 -- 00




Oil Service Page: 45

AIRBUS A-- 320 PW 1100 NEO 79 -- 00

NOTES:





Page: 46

AIRBUS A-- 320 PW 1100 NEO 73 -- 00

ENGINE FUEL SYSTEM



POWER PLANT FUEL SYSTEM

SSCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 73 -- 00

FULL AUTHORITY DIGITAL ENGINE CONTROL Description The Full Authority Digital Engine Control (FADEC) System has an Electronic Engine Control (EEC) that is designed to be an interface with aircraft displays and computers, control engine fuel flow, and it keeps engine operations stable while in all phases of operation. The engine fuel and control system supplies calculated fuel flow to the fuel nozzles for combustion, servo fuel pressure for valves and actuators operated by fuel, as well as monitors the system and indicates fault warnings or displays in the cockpit. It also adjusts engine thrust control for stable condition and temporary engine operation as related to applicable systems which include cockpit controls, aircraft configuration information, and aircraft air data. Find the PMA rotor attached to the front of the MGB at the 7 o’clock. Fuel Nozzles Find and mark the notch located at the top 12 o’clock on the diffuser case. The T2 probe is attached to the inlet cowl at the 1 o’clock. The Fuel Filter Differential Sensor is installed on fuel manifold, next to fuel filter. The Fuel Flow Transmitter is installed above the fuel filter on the high pressure turbine case at the 3 o’clock. The PMA Stator is attached to the front of the main gearbox at the 8 o’clock. The Fuel Return Pump is located on the Compressor Intermediate Case (CIC) at the 5:30 o’clock. The return to tank module is located under the Fuel/Oil Heat exchanger Bypass Valve at the 5 o’clock. Find the P2.5/T2.5 probe attached to the CIC at the 1 o’clock. Find the T3 probe attached to the diffuser case, forward of the fuel nozzles, at approximately the 1 o’clock. Find the PB sensor attached to the CIC bulkhead at the 10 o’clock. The Fuel Filter is attached to the fuel manifold on the top of the main gearbox at the 4 o’clock. Find the IFPC attached to the fuel manifold on the top of the main gearbox at the 4 o’clock. The Fuel Flow Divider is located under the Fuel/Oil Heat exchanger Bypass Valve at the 5 o’clock.


The EEC is located at the 2 o’clock attached to the fan case assembly. Find the Data Storage Unit attached to the EEC at the 2 o’clock on the fan case assembly. Find the harness WC08 attached to the right hand side of the engine core. Find the harness WF01, WH02 and WF07 attached to the RH of the fan case. The EEC Wiring Harness WC11 is located on the left hand side of the of the engine core. Find the EEC wiring harness (W03) attached to the left and right side of the fan case and to the left and right side of the engine core. Find the EEC wiring harness (W04) attached to the right side of the fan case and to the left and right sides of the engine core. The EEC wiring harness (WF06) attached to the right hand side of the fan case. The EEC Wiring Harness WC10 is located on the left hand side of the of the engine core. The fuel recovery tank is located under the fuel/oil heat exchanger bypass valve at the 5 o’clock. The Fuel Manifold is located approximately at the 4 o’clock attached to the main gearbox.

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AIRBUS A-- 320 PW 1100 NEO 73 -- 00



Figure 1 SSCL /JGB / REV.00 / Jun--2016

Fuel Componets Page: 3

AIRBUS A-- 320 PW 1100 NEO 73 -- 00

ENGINE FUEL AND CONTROL General Fuel and Engine Control systems operations are subdivided among three integrated systems: Engine Control, Fuel Distribution, and Fuel Indicating. * Engine Control This system commands engine systems and fuel flow. It maintains stable engine thrust and performance during all phases of operation. It also interfaces with aircraft displays and computers. * Fuel Distribution This system supplies metered fuel for combustion and servo fuel pressure for component actuators. * Fuel Indicating This system monitors the condition of the Fuel System. It displays warnings in the cockpit that alert the flight crew to Fuel System status and to potential problems. The primary components of the fuel system are on the RH side of the engine. The Integrated Fuel Pump and Control (IFPC) is attached with bolts to the fuel manifold on the right side of the main gearbox at the 3 o’clock.





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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Integrated Fuel Pump and Control Page: 5


POWER PLANT FUEL SYSTEM Fuel Distribution System The Fuel Distribution System supplies calculated, filtered fuel to the engine at the pressure and flow rate necessary to meet all engine requirements while in operation. The fuel is also heated to prevent ice accumulation. The system supplies calculated fuel to the fuel injectors for combustion and pressurized fuel to engine component actuators for stable pressure. The Fuel Distribution System consists of these components. * Integrated Fuel Pump and Control (IFPC) * Fuel Filter * Fuel/Oil Heat Exchanger (includes bypass valve) * Fuel Manifold * Fuel Nozzles The IFPC is an electronically controlled unit which uses a dual coil torque motor and solenoids to control fuel flow in relation to the hydro--mechanical valves. The dual coil torque motor controlled by the EEC positions the valve that measures fuel. The meter valve position is sent back to the EEC by a dual coil Linear Variable Differential Transformer (LVDT). Two solenoids in the IFPC can be used to operate the valve that meters fuel. They are the overspeed/shutdown solenoid and the pressure equalization solenoid. The EEC directs the overspeed/shutdown solenoid to position the Minimum Pressurizing and Shut Off Valve (MPSOV) to the shut off position. The dual coil shutoff solenoid is directed by the EEC to open the equalization bypass valve to make equal pressure in the primary and secondary fuel manifolds downstream of the flow divider valve. This will make sure that there is an equal fuel to air ratio in the combustor. The fuel distribution valve is part of the IFPC. The fuel distribution valve sends calculated fuel through the supply tubes to two fuel supply manifolds around the diffuser case. A valve that controls pressure internally and prevents fuel from going into the secondary fuel manifold until a set fuel pressure is applied. When the valve that controls pressure is closed at engine shut down, the primary fuel manifold goes to the injectors drain back to the diffuser case where the fuel is applied. The entire fuel manifold stays full to help the next time the engine starts. There is also a fuel temperature sensor installed on the IFPC. SSCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 73 -- 00 Integrated Fuel Pump and Control (IFPC). The IFPC consists of a fuel pump and control. The fuel pump supplies low pressure boost stage fuel to the Fuel/Oil Heat Exchanger and fuel filter and then to the gear stage inlet port of the fuel pump. The gear stage pump sends high--pressure fuel through a valve that measures fuel in the IFPC. The boost stage is a centrifugal pump and the main stage is a gear type positive displacement pump. The fuel pump is attached with bolts to the fuel manifold on the top right of the main gearbox. The IFPC supplies measured fuel flow to the engine as scheduled by the EEC. The IFPC is an electronically controlled unit which uses a dual coil torque motor and solenoids to control fuel flow in relation to the hydro--mechanical valves. The dual coil torque motor controlled by the EEC positions the valve that measures fuel. The meter valve position is sent back to the EEC by a dual coil Linear Variable Differential Transformer (LVDT). Two solenoids in the IFPC can be used to operate the valve that meters fuel. They are the overspeed/shutdown solenoid and the pressure equalization solenoid. The EEC directs the overspeed/shutdown solenoid to position the Minimum Pressurizing and Shut Off Valve (MPSOV) to the shut off position. The dual coil shutoff solenoid is directed by the EEC to open the equalization bypass valve to make equal pressure in the primary and secondary fuel manifolds downstream of the flow divider valve. This will make sure that there is an equal fuel to air ratio in the combustor. The fuel distribution valve is part of the IFPC. The fuel distribution valve sends calculated fuel through the supply tubes to two fuel supply manifolds around the diffuser case. A valve that controls pressure internally and prevents fuel from going into the secondary fuel manifold until a set fuel pressure psi is applied. When the valve that controls pressure is closed at engine shut down, the primary fuel manifold goes to the injectors drain back to the diffuser case where the fuel is applied. The entire fuel manifold stays full to help the next time the engine starts. There is also a fuel temperature sensor installed on the IFPC. Page: 6


AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Fuel Flow Transmitter Page: 7



AIRBUS A-- 320 PW 1100 NEO 73 -- 00

Fuel Filter The main fuel filter element is used to remove solid impurities from the pressurized fuel sent from the IFPC. The filtered fuel goes from the main fuel filter to the main fuel pump contained within the IFPC and then splits in two directions -- to the flow divider valve and fuel nozzles, and to the turbine Active Clearance Control (ACC), High Pressure Compressor (HPC), Low Pressure Compressor (LPC), and 2.5 bleed actuators. The main fuel filter element is in the main fuel filter cover installed on the forward side of the fuel manifold near the 4 o’clock engine position. An isolated fuel filter bypass valve is installed on the fuel manifold, adjacent to the Fuel Filter. The fuel flow passage is from the outside diameter of the filter to the inside. The fuel filter differential pressure sensor is used to find if the primary oil filter element is clogged.


Fuel Manifold The IFPC is installed and sealed on the fuel manifold, adjacent to the fuel flow transmitter, and the fuel filter. The IFPC is used to transfer fuel, and for connection points for external fuel lines. Since the fuel manifold moves fuel to and from the IFPC, the IFPC can be removed and replaced without the removal of any fuel lines. The fuel manifold is installed on the front face of the gearbox on the left side. Fuel/Oil Heat Exchanger The fuel/oil heat exchanger is used to help the movement of heat from engine oil to engine fuel. The fuel/oil heat exchanger bypass valve adjusts the amount of oil that flows through the heat exchanger as a function of fuel temperature. The fuel/oil heat exchanger heats engine fuel to prevent ice collection. The fuel/oil heat exchanger cools the engine oil for proper engine bearing/seal operation. The fuel/oil heat exchanger is installed at the 8:00 position on the Thermal Management System (TMS).


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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Thermal Management System Page: 9



AIRBUS A-- 320 PW 1100 NEO 73 -- 00

Fuel Controlling -- Governing System The EEC also controls components in different systems to hold oil and fuel temperatures within specified limits and increase engine performance. The EEC monitors the health of the different systems and reports this information to the aircraft in the form of maintenance messages and cockpit caution, status, and warning messages if necessary. The PW1000G FADEC System provides these capabilities: * Closed loop control of engine thrust parameter (N1) * Control of engine transients and stability * Control of performance features * Control of fuel and oil temperatures * Protection of engine limits (N1, N2, Pb) * Automatically starts sequence * Control of thrust reverser * Redundancy management * Efficient communication with aircraft systems * Monitors health and engine diagnostics * Overspeed/Thrust Control Malfunction (TCM) System * Components of the Controlling System Fuel Nozzles There are 18 fuel nozzles that are used to atomize fuel for combustion inside the combustor. There are 12 dual port nozzles that have both a primary and secondary fuel flow path and six single port nozzles that give only a secondary fuel flow path. The primary fuel flow path is used in the process of the start and the secondary fuel flow path is used in the process of all other power options. The nozzles are a single vent, moderate fuel pressure drop airflow design, that uses combustor inlet air to atomize the fuel. Each fuel nozzle inlet connection contains a filter inlet cover and a flow vent. Each fuel nozzle has its own support cover.


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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Fuel Nozzles Page: 11


POWER PLANT FUEL SYSTEM Electronic Engine Control (EEC). The EEC controls these functions: * Starting * Monitors engine condition * Measures fuel flow * Engine stability * Fault detection and maintenance * Electrical Power Management * Engine performance * Ignition * Heat Management * Aircraft Interface * Thrust reverser system * Overspeed/Overthrust (TCM) System The EEC is installed on the fan case at the 2:00 position, with four bolts next to the Prognostic Health Monitoring Unit (PHMU). The cast aluminum EEC cover uses natural convection to cool. The EEC has dual electronic channels, channel A and channel B, each with its own processor, power supply, program memory, input sensors and output actuators. The EEC has four installed pads that isolate vibration, an EEC data storage unit receptacle, inter--channel communication, two pneumatic pressure ports, and a ground strap for electromagnetic compatibility. The EEC uses these to improve the stable operation of the engine: * Dual channel control * Automatic fault detection * Automatic fault accommodation * Redundant inputs and outputs When fully operational, the EEC acts in active/standby mode where all engine capabilities are controlled by one channel. Upon detection of an output failure the EEC acts in an active/standby mode. Active/standby permits the EEC to switch control of the fault loop to the other channel and maintains control of the capabilities that stay in the channel given priority for that flight. Power for the EEC is supplied by the gearbox pushed Permanent Magnet Alternator (PMA). Aircraft power is available in flight to allow EEC operation in case of a PMA failure. SSCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 73 -- 00 Full self--test and fault isolation logic is programmed into the EEC and used continuously. Self--test and Line Replaceable Unit (LRU) test cycles make on--wing corrections and shop maintenance easier to isolate faults in the EEC and the input and output devices. The EEC controls these redundant solenoid systems: -- Overspeed solenoid in the IFPC -- Starter evaluation through a speed sensor (NS) -- Air Starter Valve controls pneumatic power to the air starter. -- 2.8 High Compressor Bleed Solenoids control pneumatic pressure to the bleed valve -- Directional control valve solenoid controls the hydraulic pressure for the nacelle thrust reverser system -- Isolation control valve solenoid isolates the hydraulic pressure for the nacelle thrust reverser system. -- Evaluation is determined by the strain gauge pressure sensor installed in the isolation control valve unit. The EEC controls these redundant outputs that use alternate interfaces: -- On/Off command of the dual channel ignition exciter is completed by control from each channel of the FADEC. -- On/Off of T2 heater is completed by a FADEC command to the airframe. The FADEC senses current from the heater circuit by a software discrete in each channel. The EEC controls these redundant torque motor systems: -- IFPC torque motor controls the measured fuel flow to the engine. Position analysis is from a dual LVDT. -- Stator vane actuator controls the position of the variable stator vanes. Analysis is from a dual LVDT. -- The 2.5 bleed valve actuator controls the position of the 2.5 bleed valve. Analysis is from a dual LVDT. -- The HPT case cooling and LPT case cooling are controlled by 1 actuator. The actuator positions a valve to control fan air flow for cooling the turbine cases. A single channel LVDT is used for analysis of the actuator. -- The Variable Oil Reduction Valve controls oil bypass into the gearbox. Analysis is from a dual LVDT. -- The fuel/oil heat exchanger bypass valve keeps oil and fuel temperatures within the ranges that they operate in. Page: 12


AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Fuel Nozzless Page: 13


POWER PLANT FUEL SYSTEM Data Storage Unit (DSU). The DSU has the subsequent purposes: * Selection of engine thrust rating * Records the engine serial number * Selection of N1 modification data * Selection of engine performance package The DSU is installed on the EEC A channel cover. The DSU has a single Electronically Erasable Programmable Read--Only Memory (EEPROM) and is installed on the EEC channel A. The engine gets a correct DSU in the process of the initial test cell operation. The N1/thrust relation is compared. The DSU must remain with the engine if the EEC is replaced. The DSU is installed to the engine fan case by a lanyard. The DSU is installed to the engine fan case by a lanyard. The DSU has a hardware part number, software part number, engine serial number, N1 class, and the class it is rated in stamped on it. If the DSU is not installed, the EEC will not allow the engine to start. If the DSU disconnects in flight, the EEC will use the information that is stored in the EEC memory. The engine serial number is aided by the DSU to help monitor engine condition. PHMU -- refer to PW1000G--C--77--00--00--00A--040A--D.

AIRBUS A-- 320 PW 1100 NEO 73 -- 00 -- The fuel flow transmitter is a single element device wired to channel A of the EEC. The signal is hard wired to channel B internal to the EEC. -- The fuel flow transmitter is installed above the fuel filter on the high pressure turbine case at the 3 o’clock. -- The EEC sends the fuel flow rate signal to the Engine Indicating and Cockpit Awareness System (EICAS) for display on the flight deck primary engine display. * Fuel filter differential pressure sensor -- The fuel filter differential pressure sensor measures the differential pressure across the fuel filter. If the fuel filter becomes clogged, the fuel filter differential sensor sends a signal through the EEC that warns the flight deck. -- The sensor has one independent circuit that is installed on the side of the fuel manifold, next to the fuel filter. * Fuel temperature sensor -- The fuel temperature sensor is a dual element Resistance Temperature Device (RTD). -- Each channel of the EEC is given a fuel temperature signal by one of the two elements. -- The EEC uses the fuel temperature to control oil flow through the fuel/oil heat exchanger. -- The sensor is installed on the fuel manifold.

Fuel Indicating System The Fuel Indicating System monitors fuel system condition and tells the flight crew about the fuel system status and possible problems. These conditions are monitored one after the other: * Fuel flow * Fuel filter clog warning * Engine fuel temperature The Fuel Indicating System has the components that follow. * Fuel flow transmitter -- The fuel flow transmitter sends a signal to the EEC where it is used to calculate fuel flow to the combustor.


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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Electronic Engine Control Page: 15



AIRBUS A-- 320 PW 1100 NEO 73 -- 00

FUEL DISTRIBUTION General The fuel distribution system supplies fuel from tanks to the engines. The fuel is metered, filtered and supplied at the pressure and flow rate necessary to enable stable engine operations during all the phases. The metered Fuel Flow (FF) is sent to the fuel nozzles for combustion, and pressurized fuel is supplied to the fuel--operated actuators of the engine (e.g. Air valves). The fuel is also heated to prevent ice formation and used to cool engine oil and Integrated Drive Generator (IDG) oil. The distribution system consists of: -- The Integrated Fuel Pump and Control (IFPC), -- A fuel manifold, -- Fuel/Oil Heat Exchangers (FOHEs), --A fuel filter, -- Flow Divider Valve (FDV). -- Fuel nozzles, -- Ecology collector tank, -- Return--To--Tank (RTI’) valve.

Integrated Fuel Pump and Control The IFPC is an electronically controlled unit wich integrates the fuel metering components and the fuel pumps in a single unit to limit the space and the number of the external tubes required for the system. The IFPC uses a dual coil torque motors and solenoids to control hydro--mechanical valves in relation to the fuel flow. The MGB turns the IFPC input shaft wich drives the fuel pump boost satge, the main fuel pump and servo pump. Fuel Filter and Main Pump The heated fuel from the engine FOHE is directed through the fuel filter. The filter element is a disposable filter located in a housing attached to the fuel manifold. The filter is monitored by a differential pressure switch. The filter housing is fitted with a bypass valve in case of filter element clogging. The filter element is a disposable 25 micron filter. The fuel exits the fuel filter and flows to the inlet port of the main fuel pump. The main fuel pump is a single -- satge gear pump, wich increases the fuel pressure and sends the pressurized fuel to Fuel Metering Valve (FMV).


Fuel Feed from Aircraft When the engine master lever is selected ON, the LP Shut Off Valve opens and fuel from the aircraft tanks flows through the main fuel supply line to the inlet port of the boost pump in the IFPC Heat Exchanger and Fuel Return To Tank The boost pump sends LP fuel from engine fuel supply line to the IDG FOHE. Fuel flow is used to cool down the IDG oil through the IDG FOHE an the engine oil through the engine FOHE. It turn, fuel is heated and de--iced. Fuel from the engine FOHE is then send to the fuel filter. The Fuel Return To Tank (FRTT) module contains the the FUEL Return Valve (FRV) and the FRTT sensor. The FRV controls fuel to flow back to the aircraft tanks from downstream of the IDG FOHE and before it enters the engine FOHE as part of the fuel heat management system. The FRV is controlled by the Electronic Engine Control (EEC) depending of the fuel temperature.


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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




Fuel Distribution Page: 17


POWER PLANT FUEL SYSTEM Fuel Metering Valve and High Pressure Shut--Off Valve The EEc controls a dual Torque Motor (TM) wich positions the FMV in the desired position. The close loop monitoring is ensured by the EEC using the valve LVDT feedback signals. The fuel from the FMV is directed to the High Pressure Shut--Off Valve (HPSOV). The fuel pressure at the back side of the HPSOV is controlled by the Thrust Control Malfunction (TCM) / Overspeed TM and allows the valve to open or close.


Pressure Regulating Valve and Bypass Directional Control Valve Inside the IFPC, thye fuel from the main pump is directed to the FMV and to the Pressure Regulating valve (PRV). The purpose of the PRV is to maintain a constant fuel pressure drop across the FMV to ensure the correctfuel flow and acceleration for the engine. The TCM / Overspeed TM controls the fuel pressure to the back side of the PRV to modulate fuel flow between the FMV and the Bypass Directional Control Valve (BDCV). Pressurized fuel that passes through the PRV is directed to the BDCV. The BDCV The BDCV directs fuel By--passed by the PRV to the engine FOHE at low engine power or when the fuel temperature is low to help in maintaining the engine oil and fuel within operating limits. At high power, the BDCV returns the recirculation flow downstream of the FOHE. EEC Control The EEC controls the dual TCM / Overspeed TM for HPSOV positioning. It monitors the valve fully closed position with the two proximity switches. The EEC also controls the FMV position via a dual channel Torque Motor (TM). A dual channel Linear Variable Differential Transducer (LVDT) provides the FMV position to the EEC. For the Air system, the EEC controls the fuel operated actuators with dual channel TMs and it monitors their position thanks to LVDT position feedback.

AIRBUS A-- 320 PW 1100 NEO 73 -- 00 The EEC commands the FDV opening during starting to improve fuel atomization. During engine start, the FDV sends most of fuel to the primary manifold. Above idle, the FDV evenly divides metered fuel flow between the primary and secondary manifolds. At shutdown, the FDV es springloaded closed to allow primary and secondary drainage. The FDV is fitted with a metal screen stariner that can be bypassed in case of blockage. There are 18 fuel nozzles mounted to the outer diffuser case. All the nozzles atomize fuel inside the combustor. Twelve of them are duplex nozzles featuring both a primary and a secondary fuel flow paths while six others are simplex nozzles providing only a secondary fuel flow path. Servo Fuel and Servo Minimum Pressure and Pump Sharing Valve The servo pump housed in the IFPC is a gear stage pump wich sends pressurized fuel to a wash filter. Fine filtered, pressurized fuel from the wash filter is supplied to the engine air system actuators where it is used as servo and muscle pressure to position the actuator pistons. The actuators are: -- the Low Pressure Compressor (LPC) Stator Vane Actuator (SVA) -- the LPC (2.5) Bleed Valve Actuator (BVA) -- the turbine Acitve Case Cooling (ACC) valve -- and the High Pressure Compressor (HPC) SVAs (primary and secondary) The fuel from the actuators is filtered again before it returns back to main pump and servo pump inlet. The Servo Minimum Presssure and Pump Sharing Valve is a spring loaded valve that provides the five air system actuators with main pump fuel pressure when servo pump fuel pressure is not enough during start.

Fuel Flow Transmitter, Flow Divider and Fuel Nozzles The metered fuel from the FMV crosses the HPSOV and flows to the Fuel Flow Transmitter (F/F Tx). The F/F Tx sends the fuel flow rate to the EEC channel A and directs fuel to the Flow Divider Valve(FDV). SSCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 73 -- 00






POWER PLANT FUEL SYSTEM ECOLOGY SYSTEM General At engine shutdown, residual fuel in the manifolds dowstream of the FDV is drained back through the FDV to an ecology collector tank. The collected fuel remains in the ecology tank until the next engine start when the fuel is drawn back into the fuel system. During shutdown, the fuel pressure from the IFPC is reduced and the FDV closes to prevent fuel from entering the combustor and to drain any fuel remaining in both the primary and secondary fuel lines to the ecology collector tank. The ecology collector tank has enough space to receive fuel from a single engine shutdown. The tank has an inlet float valve wich closes when the tank has reached its maximum capacity. This prevents the tank from overfilling and spilling fuel out following and aborted start. At next start up, the ejector pump draws the fuel from the ecology collector tank back to the IFPC boost pump. The tank has an outlet float valve wich closes when the tank has reached its minimum capacity and a chec valve to avoid fuel tarnsfer from the suction line. Starting During starting, the servo pump fuel pressure is not enough to control the air system actuators and to close the Servo Minimum Pressure and Pump Sharing Valve. In this position, the Servo Minimum Pressure and Pump Sharing Valve directs a portion of pressurized fuel from the main pump to five actuators. The other portion of fuel from main pump is sent to the PRV and to the FMV. The PRV opens partially and directs the excessof fuel flow to the BDCV wich is spring loaded to send it to the engine FOHE. The EEC opens the FMV and let the fuel to flow to the HPSOV wich aldo opens and sends fuel to the fuel flow transmitter. The pressurized fuel opens the FDV. The FDV partly opens and sends most of fuel to the primary and secondary fuel nozzles.

AIRBUS A-- 320 PW 1100 NEO 73 -- 00 The FDV also opens more and evenly divides metered fuel flow between the primary and secondary fuel nozzles. In parallel, the fuel pressure from the servo pump increases and pushes the Servo Minimum Pressure and Pump Sharing Valve, segregating the burn flow from the servo fuel. Normal Shutdown During a normal engine shutdown, the Master Lever controls the LPSOV to close and sends a shutdown signal to the EEC. As a consequence, the EEc controls the TCM / Overspeed TM that directs fuel pressure to the back side of the HPSOV to close it and stop the fuel flow to the engine. In the same time, the PRV is controlled fully open to bypass the main pump fuel flow away from the FMV to the FOHE. In turn when the related fuel pressure drops, the FDV closes to let the remaining fuel in the nozzle manifold to drain in the ecology drain tank, and the Servo Minimum Pressure and Pump Sharing Valve reopens. After the HPSOV is confirmed closed by the proximity switches, the EEC tests the FMV via its TM then closes it. Abnormal Shutdown The abnormal shutdown is initiated in case of an overspeed (N1 or N2) , shaft shear (Fan, LP or HP) or Thrust Control Malfunction (TCM) even detected on ground. In such case, the TCM / Overspeed TM directs fuel pressure to the back side of the HPSOV and of the PRV. This cause the PRV to open and stop fuel flow to the FMV, allowing rapid closure of the HPSOV and rapid engine shutdown. Fuel flow through the PRV is directed to the BDCV and then to the engine FOHE. This shutoff method is independent from the FMV control.

Acceleration As the pumps rotation speed increases with the engine acceleration, the fuel pressure also increases. The FMV opens more and as a consequence the fuel pressure pushes the BDCV out of its rest position to direct the excess fuel flow to the fuel filter. SSCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 73 -- 00







AIRBUS A-- 320 PW 1100 NEO 73 -- 00

FUEL INDICATING General The engine fuel indicating monitors the system condition and provides the system status to the cockpit displays. The fuel flow transmitter send signal to the EEC wich enables the calculation pf the fuel flow to the combustor. The fuel flow is a primary engine parameter and is displayed on the EWD permanently. The EEC also send this data for the fuel used computation and display on the System Display (SD). The Fuel Filter Differential Pressure (FFDP) sensor measures the differential pressure across the fuel filter. This helps to detect if the filter is partially or totally clogged. According to the received value, the EEC will generate various warnings on the EWD: * ENG X FUEL FILTER DEGRAD or * ENG X FUEL FILTER CLOG or * ENG X FUEL SENSOR FAULT and on the SD * CLOG The IDG Fuel Oil Heat Exchanger (FOHE) differential pressure sensor is used to sense the differential pressure on the fuel side of the FOHE and send a signal to the EEC in case of clogging detection. According to the status, the EEC will generate various warnings on the EWD: * ENG X HEAT EXCHANGER CLOG or * ENG X FUEL SENSOR FAULT. For monitoring and Thermal Management System control by the EEC, the fuel temperature is sensed by two dual channel temperature sensors. The fuel temperature sensor is used for the controlof the Heat Exchangers Fuel /Oil Heat Exchanger Bypass Vakve (FOHEBV) and BDCV. The Fuel Return To Tank (FRTT) temperature sensor is used for the RTTV control. The engine fuel temperature is not directly displayed in the cockpit but, according to the status, the EEC will generate various warnings on the EWD: * ENG X HOT FUEL or * ENG X FUEL HEAT SYS or * ENG X HEAT SYS DEGRADED or * ENG X HEAT SYS FAULT.


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AIRBUS A-- 320 PW 1100 NEO 73 -- 00




EWD and MCDU Page: 23

AIRBUS A-- 320 PW 1100 NEO 73 -- 00

NOTES:





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AIRBUS A-- 320 PW 1100 NEO 76 -- 00

ENGINE CONTROL SYSTEM



POWER PLANT ENGINE CONTROL


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POWER PLANT ENGINE CONTROL PROPULSION CONTROL General The Propulsion Control System (PCS) consists in EIU and FADEC System wich includes EEC and Pregnostic s and Healt Management Unit (PHMU). Each EIU is decicated to an engine. EIU 1 and 2 are located in the aircraft avionics bay 80VU. The EEC and PHMU are attached to the engine fan case assembly at 2:30 O’clock. Engine Interface Unit Each EIU is an Interface Concentrator between the airframe and the corresponding EEC on the engine. It ensure the segregation of the 2 engines and aircfraft electrical power supply to the FADEC. In concentrates data from or to the cockpit panels and displays It gives Logics and information to or from other aircraft systems as Flight/Ground from Landing Gear Control and Interface Unit (LGCIU). FADEC The FADEC consists in a dual channel EEC with corsstalk and failure detection, a PHMU and sensors used for control and monitoring. The FADEC system manages the engine thrust and opotimizes the performance. The EEC interfaces with most of the A/C systems through the EIU. The FADEC controls the engine parameters displayed in the cockpit. The primary parameters (N1, N2, EGT and FF are sent by the EEC to the ECAM through Display Management Computers (DMCs). The engine sytem page shows secondary parameters: oil quantity, pressure and temperature. The vibration figures are communicated by the PHMU to the EEC. The Flight Warning System (FWS) will gather necesary information directly from EEC, EIU, System >Data Acquisition Concentrator (SDAC) and generates associated messages on Engine Warning Display (EWD).


AIRBUS A-- 320 PW 1100 NEO 76 -- 00

Engine Limit Protection The FADEC ensures engine integrity protection. It provides overspeed protection for N1 and N2 or rotor shaft shear by driving to close the Thrust Control Malfunction (TCM) / Overspeed torque motor in the Integrated Fuel Pump and Control (IFPC). Shaft shear detection logic is only active at high power setting. It ensures overheat protection by monitoring EGT, nacelle and EEC temperature.


Power Management The FADEC provides automatic engine thrust control and thrust parameter limit computation. The EEC uses air data parameters from Air Data/Inertial Refference System (ADIRS) for rating calculations. The FADEC manages power according to two thrust modes: -- manual mode depending on Throttle Lever Angle (TLA). -- autothrust mode depending on autothrust function generated by the AFS. The FADEC also provides two idle mode selections: -- minimum idle and -- approach idle. If the aircraft is on ground and extend the slats the engine will stay at minimum idle but in flight it will go to approach idle. The idle can also be modulated up to approach idle depending on: -- airconditioning demand -- wing anti ice demand -- engine anti ice demand and --oil temperature (for IDG cooling). Engine System Control The FADEC provides optimal engine operation by controlling: -- combustor metering valve and fuel flow, -- compressor airflow and turbine case cooling, -- thermal management (oil cooling, fuel heating) -- control and monitoring sensors, -- BITE (fault detection, isolation, annunciation and transmission to the A/C). -- nacelle anti ice. Starting and Ignition Control The FADEC controls the engine start sequence in automatic or manual mode when initiated from the control panels. It monitors N1, N2, EGT and oil parameters and then can abort or recycle an engine start. Thrust Reverser The Fadec supervises the thrust reverser operation. Page: 2


AIRBUS A-- 320 PW 1100 NEO 76 -- 00




Propulsion Control Page: 3



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

FADEC ARQUITECTURE General The FADEC consists in the Electronic Engine Control (EEC), the Prognostic and Healt Monitoring Unit (PHMU) and perispherals (sensors and output drivers). EEC The EEC is a microprocessor controlled digital unit with two independentcontrol channels identifiednas channel A and B. Each channel has its own porccessor, power supply, program memory,selected input sensors and output drivers. In addition to input/output redundancy (for comparision and backup), data is sent internallybetween the two channels by a crosstalk data link. Each channel receives inputs from the A/C and FADEC system sources. Thus, each channel can monitor and contorl the operation of the engine and transmit engine data to the A/C and to engine subsystem duplicated controls (torque motors and solenoids). Each channel A and B are housed in one assembly but arte phisically divided by a two piece modular desihn. Each channel module has one printed circuit board module, the input/output interconnector modules and one pressure sensor. Five electrical connectors are used in each channel module to connect wiring from the engine, aircraft and nacelle. The EEC also has a connector to test the unit and a connector for the Data Storage Unit (DSU).

PHMU The PHMU is a single channel component with internal software that performs the following engine health monitoring functions: * Vibration analysis * Oil Debris Monitoring (ODM) * Auxiliary Oil Pressure (AOP) signal conversion. It uses data provided by several engine sensors and by the EEC and sends back the computed data to the EEC through CAN buses. Two connectors are used for the data exchange


DSU The DSU is a data memory plug attached to the engine case bracket by a lanyard connected on the EEC channel A for engine identification and rating, engine trim data storage and detected failures storage.


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FADEC Arquitechture Page: 5



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

PROCESS General Most ofthe FADEC operations are based on the same principie, they respond to a demand from the A/C or from the EEC interna! schedules, and they take into account input parameters from the A/C and from the engine sensors. Most ofthe sensors and output drivers are duplicated for redundancy and segregated to each EEC channel. For a control loop, one EEC channel elaborates a single command signal sent to an engine subsystem control and it makes sure that its command has been followed by monitoring the dual feedback from this engine subsystem. The EEC also continuously perfom integrity test of its control circuits. When fully operational, the EEC starts and operates in an Active--Standby mode. Under this control scheme, only one channel of the EEC has full authority over all engine functions and is identified as the preferred channel. The preferred channel is altemated upon every engine shutdown for the next engine start. If a feedback fault is detected in the preferred channel, the data is retrieved from the standby channel via the crosstalk data link. If an output driver fault is detected, the EEC switches from Active--Standby mode to Active--Active mode. This allows either channel to control any ofthe output drivers independently, regardless ofwhich channel is the preferred channel. This control mode allows both channels to be engaged simultaneously and to manage different engine functions, providing an effective fault accommodation strategy. If the crosstalk data link is lost, each channel maintains its current controls prior the failure. If the engine subsystem control loop is no more possible (by any channel), the subsystem control is set to its failsafe position .


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AIRBUS A-- 320 PW 1100 NEO 76 -- 00




Schematic Process Page: 7


POWER PLANT ENGINE CONTROL FADEC INTERFACES General In order to provide a full range of engine control and monitoring, the Propulsion Control System (PCS) exchanges data within its own computers Engine Interface Unit (EIU), Electronic Engine Control (EEC), Prognostic and Healt Monitoring Unit (PHMU) and with the other aircraft systems computers. The EIU is the main interface with the aircraft systems. Inputs or outputs are transmitted on a digital, analog or discrete format. PCS Interfaces The EIU performs the following bus transfer. EIU digital inputs from: GCU #: for idle modulation based on IDG load. DLRB: for EIU dataloading. ACSC 1/2: for bleed decrement computation. CFDIU: for BITE purposes (Normal Mode and Menu Mode). BMC 1/2: for bleed computation. LGCIU 1/2: for flight/ground status computation. FCU: for Autothrust function and TCM protection in flare. EIU digital outputs to: ADIRU 1/2: for air data correction. CFDIU: for BITE purposes (Normal Mode and Menu Mode). DLRB: for EIU dataloading. SDAC 1/2: for engine parameters acquisition. FDIMU (ACMS): for condition monitoring and troubleshooting purpose. BMC #: for bleed computation FWC 1/2: for warnings display. The EIU performs the following discrete exchange. EIU discrete inputs from: From cockpit controls: -- Master lever ON/OFF -- Throttle position (switches): for thrust reverser operation. -- Rotary selector Ignition/Auto/Crank SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 76 -- 00 -- Wing De--Ice P/B OFF: for bleed decrement computation. -- Nacelle AntiIce ON/OFF: for NAI control and bleed decrement computation. -- Fire handle ON: for engine isolation. Manual Engine Start P/B ON -- FADEC Ground Power ON -- Bump ON/OFF -- APU Master Switch ON/OFF: for bleed decrement computation. From LGCIUs: -- LH Landing Gear compressed: for flight/ground status computation. -- RH Landing Gear compressed: for flight/ground status computation. -- Nose Landing Gear (NLG) compressed: for flight/ground status computation. From SECs: -- Ground Spoiler OUT -- TLA <--3 deg From SFCC: -- Flaps and Slats lever retracted From engine: -- FRTTV Selected OFF (EEC) -- Low Oil Pressure sensor: for OIL LO PRES Warning. -- Engine position and type -- Latch Door Monitoring Proximity Switches. EIU discrete outputs: Fuel HPSOV Closed N2 Not Below Idle TLA in Take Off Position Start Valve Closure APU Boost Command Master Lever Fault Light Oil Low pressure and Ground NAI P/B Fault Light Latch Door Monitoring Proximity Switches The EIU provides the following power supplies. Page: 8


AIRBUS A-- 320 PW 1100 NEO 76 -- 00




PCS Interfaces Page: 9


POWER PLANT ENGINE CONTROL EIU power supply outputs to: PHMU (28V DC). Hydraulic pump depressurization solenoid (28V DC). EEC channels (28V DC). Igniters (115V AC). Thrust reverser Valves (28V DC for ICV & DCV). Unless specified differently, signals are dual (from/to both EEC channels). The EEC performs the following bus transfer. EEC digital inputs from: EIU # (channel A): for aircraft data exchange. ADIRU 1/2: for engine control (alt, TAT, PT, CAS, Mn). PHMU: for vibration monitoring and trim balancing. EEC digital outputs to: EIU #: for engine data exchange. FMGC 1/2: for Autothrust function and TCM protection in flare. PHMU: for vibration monitoring and trim balancing. DMC 1/2/3: for parameters, faults and warnings display. FWC 1/2: for warnings display. GCU #: for power supply management. The EEC performs the following discrete/analog exchange. EEC discrete/analog inputs from: Cockpit controls: Master lever OFF: for shutdown and reset. Throttle position (resolvers): for manual and auto thrust control. Autothrust disconnect P/B (ch B) FADEC Ground Power OFF Nacelle Anti--Ice ON/OFF: for NAI control and bleed decrement computation. FCU: Autothrust engagement (ch B) SECs: TCM ground operation Engine: SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 76 -- 00 Engine sensors and subsystems feedbacks Engine position (ch A). EEC discrete/analog outputs to: PHMU: Nf (ch B), Ni (ch A), N2 (ch A) Engine subsystems: Control signals EIU: FRTTV Selected OFF.

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AIRBUS A-- 320 PW 1100 NEO 76 -- 00

LGCIU

ACSC

BMC

SDAC

DLRB

FDIMU

CFDIU

SECs

SFCC

EIU

FCU DMC

FWC

ADIRU

GCU

FMGC

EEC


PHMU


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POWER PLANT ENGINE CONTROL FADEC INTERFACES FADEC Interfaces Unless specified diffcrently. signals are dual (from/to both EEC channels). The EEC is the main contro!Jer and monitoring devicc over the enginc subsystems. Air System For the air system management, the EEC sends and receives the following data. Compressor Stator Vane Control System: -- LPC SVA TM control signal, -- HPC master SVA Torque Motor (TM) control signal , -- LPC SVA. HPC master and slave SVAs LVDT feedback signal. Compressor Bleed Control System : -- LPC Bleed Valve Actuator (BVA) TM control signal, -- LPC BVA LVDT feedback signal, -- HPC BY solenoid control signal, -- HPC active and passive bleed pressure sensors. Turbine Active Case Cooling Control System: -- TACC Valve TM control signal -- TACCV LVDT feedback signal (ch A). Buffer/ Ventilation Control System : -- HPC Buffer Shut OffValve (SOY) solenoid feedback signal, -- Buffer Air Pressure Sensor (BAPS) feedback signal. Fuel System For the fuel system management the EEC sends and receives the following data. Fuel Supply for combustion: -- Fuel Metering Valve (FMV) TM control signal, -- FMV LVDT feedback signal. -- TCM /Overspeed TM control signal. -- HP Shut Off Valve proximity switch feedback signal, -- Fuel Flow Meter (FFM) control signal (eh A), -- Flow Divider Valve (FDV) solenoid control signal, SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 76 -- 00 -- Fuel Filter Differential Pressure Sensor feedback signal. Thermal Management System : -- Bypass Direction Control Valve (BDCV) solenoid control signal, -- Fuel Temperature sensor feedback signal, -- Fuel Retum To Tank (FRTT) Val ve solenoid control signal, -- FRTTV Proximity Switch feedback signal, -- FRTT Temperature Sensor feedback signal, -- IDG Fuei/Oil Heat Exchanger Diff. Pressure Sensor feedback signal. Oil System For the oil system management the EEC sends and receives the following data: Oil Supply: -- Oil Filter Differential Pressure sensor feedback signal, -- Fuel Oil Heat Exchanger Bypass Valve (FOHEBV) TM control signal, -- Active Oil Damper Valve (AODV) solenoid control signal, -- Variable Oil Reduction Valve (VORV) TM control signal, -- VORV LVDT feedback signal. Oil Monitoring: -- Oil Level (OL) sensor feedback signal (ch B), -- Main Oil Pressure (MOP) sensor feedback signal, -- Main Oil Temperature (MOT) sensor feedback signal, -- Auxiliary Oil Pressure (AOP) sensor feedback signal via PHMU, -- Oil Debris Monitoring (ODM) sensor feedback signal (ch A) vía PHMU. -- Low Oil Pressure (LOP) switch feedback signal to the ElU. Ignition and Starting System For the ignition and starting systems management, the EEC sends and receives the following data: Ignition: -- Ignition Exciter control signal (2 pairs). Starting: -- Starter Air Valve (SAV) solenoid control signal, -- Air starter speed sensor feedback signal. Page: 12


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FADEC Interfaces Page: 13


POWER PLANT ENGINE CONTROL FADEC INTERFACES Nacelle Anti--Ice System For the Nacelle Anti Ice system management, the EEC sends and receives the following data: NAI: -- Upstream PRSOV solenoid control signa! (ch B), -- Downstream PRSOV solenoid control signal (ch A), -- Upstream pressure sensor feedback signal (ch B), -- Downstream pressure sensor feedback signal, -- Dual temperature sensor feedback signal.


Thrust Reverser For the thrust rteverser system management, the EEC sends and receives the following data. Thrust Reverser -- Isolation Control Valve (ICV) solenoid control signal by EIU and EEC. -- ICV pressurized proximity switch feedback signal. -- ICV inhibition proximity switch feedback signal. -- Directional Control Valve (DCV) solenoid control signal by EIU and EEC. -- Locking Feedback Actuators primary lock proximity Sw feedback signal. -- Locking Actuators primary lock proximity Sw feedback signal. -- Locking Feedback Actuators LVDT feedback signal. . Track Locks proximity switch feedback signal. Note: Tertiary Lock Valve (TLV) solenoid are controlled independently by SEC.


AIRBUS A-- 320 PW 1100 NEO 76 -- 00 Engine Sensors For the engine control and monitoring, the EEC receives the following data. Engine Sensors: -- N1 feedback signal. -- Nf feedback signal. -- N2 feedback signal. -- P ambient feedback signal (ch A) -- Ps14 feedback signal (ch B) -- P2 feedback signal. -- P25 feedback signal (ch A). -- P3 feedback signal (2 pairs). -- T2 feedback signal. -- T25 feedback signal (ch A). -- T3 feedback signal. -- Core Nacelle Temperature feedback signal (ch B). -- NAI Temperature feedback signal. -- EGT feedback signal (2 pairs). -- Forward Vibration feedback signal to PHMU. -- Aft Vibration feedback signal to PHMU. Cockpit Controls For the engine control, the EEC receives the following data, Cockpit Controls: -- Master Lever position, -- Thrust Lever Resolver Angle, -- Auto Thrust (A/THR) Disconnect P/B (ch B), -- Flight Control Unit (FCU) A/THR engagement (ch B), -- FADEC Ground Power P/B, -- NAI P/B.

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AIRBUS A-- 320 PW 1100 NEO 76 -- 00

LPC SVA HPC SVA HPC SLV SVA LPC BVA HPC BV HPC ACTIVE & PASSIVE BLEED PRESS SENSOR TACC VALVE HPC BUF SOV BAPS

UPSTREAM PRSOV


DOWNSTREAM PRSOV UPSTREAM PRESS SNSR DOWNSTREAM PRESS SNSR DUAL TEMP SNSR

FMV TCM/OVRSP HPSOV FFM FDV DP SNSR

OIL LEVEL MOP SNSR MOT AOP SNSR ODM SW LOP

BCDV FUEL TEMP FRTT FRTTV PROX SNSR FRTT SNSR IDG/OHE nP SNSR

ICV ICV PRESS ICV INHIB DCV LOCKING FDBK ACTR LOCKING ACTR TR SLEEVE POSITION TRACK LOCKS


OFDP SNSR TM FOHEBV SOL AODV TM LVDT VORV

N1 Nf N2 Pamb Ps1.4 P2 P2.5 P3 T2

T2.5 T3 CORE NACELLE TEMP NAI TEMP EGT FWD VIB AFT VIB

FADEC Interfaces Page: 15



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

FADEC ELECTRICAL POWER SUPPLY EEC The Electronic Engine Control (EEC) is electrically supplied by the aircraft electrical network when high pressure rotor speed (N2) is below 10% or when the dedicated Permanent Magnet Alternator (PMA) has failed, and then by its dedicated PMA when N2 is above 10%. Aircraft Power The EEC is supplied by the A/C electrical power network when N2 is below 10%. Each channel is independently supplied by the A/C 28V DC through the Engine Interface Unit (EIU). The aircraft 28Y DC permits the EEC to: -- automatic ground check of the FADEC system when the engine is not running, that is to say FADEC GrouND PoWeR ON for interactive tests and data loading. -- control starting: MASTER lever ON or mode selector on IGNition or CRANK, Starter Air Valve (SAV), -- control reverser system. Note: The EIU takes its power from the same bus bar as the EEC.

Manual Re--Powering For maintenance purposes and MCDU engine tests, the ENGine FADEC GrouND PoWeR panel permits FADEC power supply to be restored on the ground with engines shutdown. When the corresponding ENGine FADEC GrouND PoWeR P/B is pressed ON the EEC recovers its power supply. Note: The FADEC is also repowered as soon as the engine start selector is in IGNition/START or CRANK position, or the MASTER lever is selected ON. Subsystems Power Supply The PHMU receives aircraft 28VDC directly from the the aircraft normal DC power bus through the EIU. The de--powering conditions are the same as the EEC. The Fan cowl door proximityare supplied by another bus in 28 VDC. Power is also transferred to the reverser system valves for Directional Control and Isolation. Each starting igniter is independently supplied with 115 VAC


PMA Supply As soon as the engine is running above 10% of N2, its PMA directly supplies each EEC channel with three--phase AC power. Two transformer rectifiers provide 28V DC power supply to channels A and B. Switching between the A/C 28V DC supply and the dedicated alternator power supplies is done automatically by the EEC. Auto De--Powering The FADEC is automatically depowered on the ground, through the EIU, after engine shutdown. The EEC automatic depowering occurs on the ground: -- 5 min after A/C power--up, -- 5 min after engine shutdown. Power is not cut--off if CFDS EEC menus are active or Data Loading going on (soft\vare upload/memory dump). Note: An action on the ENGine FIRE P/B provides EEC power cut--off from the A/C network.


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FADEC Power Supply Page: 17



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

THROTTLE CONTROL SYSTEM Throttle Control Lever The Throttle control handle comprises: * a throttle controllever which incorporates stop devices, autothrust, instinctive disconnect pushbutton switch. * a graduated fixed sector * a reverse latching lever. The throttle controllever is linked to a mechanical rod. This rod drives the input lever ofthe throttle control artificial feel unit. The throttle controllever moves over a range from --20 deg.TLA (Reverser Full Trottle stop) to +45 deg.TLA: --20 degrees TLA corresponds to Reverser Full Throttle stop +45 degrees TLA corresponds to Forward Full Throttle stop An intennediate mechanical stop is set to 0 deg. TLA. This stop is overridden when the reverse latching lever is pulled up for selection of the reverse power. This stop is reset as soon as the throttle control lever is selected back to forward thrust area. In the fonvard thrust area, there are two detent points, the MAX CLIMB detent point set to 25 deg.TLA and the MAX CONTINUOUS/FLEX TAKE--OFF detent point set to 35 deg.TLA. In the reverse thrust throttle range. there is one detent point at -- 6 deg.TLA . This position agrees with the selection of the thrust reverser command and the Reverse Idle setting. In the middle throttle range (0 deg. To 35 deg.TLA), the autothrust function can be active if engaged. This range agrees with the selection of MAX CLIMB or MAX CONTINUOUS thrust limit mode (in single operation). If the autothrust is not engaged, the engine control is manual. In the forward range (35 deg. To 45 deg.TLA), the autothrust function cannot be activated (except in alpha fioor condition).This range agrees with the selection of FLEX TAKE--OFF/MAX TAKE--OFF Mode.


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Throttle Control Lever Page: 19



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

THROTTLE CONTROL UNIT General A mechanical rod transmits the throttle control lever movement. It connects the throttle artificial feel unit to the input lever of the throttle control unit. The control unit comprises: * An input lever * Mechanical stop, wich limit the angular range. * 2 resolvers (one per each channel) * 6 potentiometers * A device , wich drives the resolver and potentiometer * A pin device for rigging the resolver and potentiometers *1 switch whose signal is dedicated to the EIU * 2 output electrical connectors The input lever drives two gear sector assembled face to face, Each sector dirves itself a set of one resolver and three potentiometers. The relationship between the throttle lever angle and throttle resolver angle (TRA) is linear and 1 degrees TLA = 1,9 degrees TRA. The accurancy of the throttle control unit (error between the input lever and the resolver angle) is 0,5 degrees TRA. The maximum discrepancy between the signals generated by two resolvers is 0,25 degrees TRA. The TLA resolver operates in two quadrants. The first quadrant is used for positive angles and the second quadrant for negative angles. Each resolver is dedicated to one FADEC channel (EEC) and receives its electrical excitation current (6 VAC) from the related FADEC channel. The EEC considers a throttle resolver angle value: * or less than --47.5 degrees TRA * or greather than 98,8 degrees TRA as resolver position signal failure. The EEC includes a resolver fault accomodation logic. This logic allows engine operation after a failure or a complete loss of the throttle resolver position signal.


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TCU Page: 21

AIRBUS A-- 320 PW 1100 NEO 76 -- 00

THROTTLE CONTROL SYSTEM BUMP Function If an airline requests the bump function , this function is selected in the aircraft by guarded pushbutton svitch with TLA at TOGA position (one on each throttle control lever). With this switch, a signal can be sent to the two FADEC units at the same time through the Engine Interface Unit (EIU) . Thrust bump can be used to obtain additional thrust capability during takeofflt can be used either with two engines or in single engine operation. With the throttle levers at TOGA and the Bump P/B pushed, ’B’ appears on the right side of the EPR/N 1 dial on the EWD.





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Bump Function Page: 23



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

ENGINE THRUST MANAGEMENT General The engine thrust is controlled under the management of the Electronic Engine Controller (EEC). The engine thrust can be set: -- manually from the throttle control lever or, -- automatically from the Auto Flight System (AFS) The engine thrust parameters are displayed on the ECAM. The main thrust monitoring parameters is the N1 speed (LP shaft). The main thrust demand parameter is the engine Fuel Flow (FF) .


Thrust Limit Mode The throttle levers are used as thrust limit mode selectors. Depending on the throttle lever position, a thrust limit mode is selected and appears on the upper ECAM display. The throttle levers are set between two detent points, the upper detent will determine the thrust limit mode. An additional Soft Go--Around (SGA) mode is available. lt is automatically selected if during approach, the TOGA detent is set and the thrust levers are then moved back to the FLX/MCT detent. Note: -- On the ground with the engines running, the displayed N1 rate limit corresponds to the TO/GA thrust limit whatever the thrust lever position is. -- On the ground with the engines runnig and if FLEX mode is selected. FLEX N1 is displayed whenever the thrust lever position is between IDLE and FLX/MCT. N1 LIMIT For each thrust limit mode selection, an N1 rating limit is computed by the EEC according to Thrust Lever Angle (TLA) and the air data parameters from the Air Data Inertial and Reference Units (ADIRUs). This indication is displayed in green on the upper ECAM display near the thrust limit mode indication. PREDICTED N1 The predicted N1 is indicated by a blue circle on the N1 indicator and corresponds to the value determined by the TLA. SCL /JGB / REV.00 / Jun--2016

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Engine Thrust Management Page: 25



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

ENGINE THRUST MANAGEMENT ACTUAL N1 The actual N1 is the actual valu e given by the N1 speed sensor and is used as a reference for the engine thrust control loop. This actual N1 is displayed in green on the N1 indicator. N1 COMMAND The N1 command, used to regulate the fuel flow, is: N1 TARGET In A/THR mode, the FMGCs compute an N1 target according to the AFS command, the air data and the engine parameters and send this demand to the EECs .


AUTOTHRUST CONTROL MODE The A/THR function is engaged manually when the A/THR P/B is selected or automatically at take--off power application . AUTOTHRUST ACTIVE When engaged, the A/THR function becomes active when the throttle levers are set to CLimb detent after take--off. The N1 command is the FMGC N1 target. The A/THR function is normally active when the throttle levers are set between IDLE and CLimb (including CLimb). The AITHR active range is extended to MCT in the case of single engine operation. When the throttle levers are set between two detent points, the N1 command is limited by the throttle lever position. Note: In case of Alpha Floor detection, the A/THR function becomes active automatically and the N1 target is to TOGA


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Autothrust Control Mode Page: 27



AIRBUS A-- 320 PW 1100 NEO 76 -- 00

ENGINE THRUST MANAGEMENT AUTOTHRUST NOT ACTIVE When engaged, the A/THR function becomes inactive when the throttle levers are set above CLimb with both engines running. In this case, the N1 command corresponds to the N1 throttle (TLA). Note: -- The A/THR function is inactive above MCT in case of single engine operation. -- The A/THR function is disengaged when the throttle levers are set at IDLE stop .


MANUAL CONTROL MODE The engines are in manual control mode when the A/THR function is not engaged or engaged and not active (throttle levers are not in the A/THR operating range and no Alpha Floor detected).


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Autothrust Not Active Page: 29

AIRBUS A-- 320 PW 1100 NEO 76 -- 00

ENGINE THRUST MANAGEMENT THRUST CONTROL MALFUNCTION The Thrust Control Malfunction (TCM) is a FADEC protection function against un--commanded and uncontrollable excessive power excursion in which the normal thrust control becomes inoperative. Note: The FADEC logic uses TCM penmission data from FMGCs to FCU to automatically reduce engine thrust during flare .





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AIRBUS A-- 320 PW 1100 NEO 76 -- 00

FLIGHT MANAGEMENT and GUIDANCE COMPUTER (1)

FLIGHT CONTROL UNIT (FCU)

FLIGHT MANAGEMENT and GUIDANCE COMPUTER (2)

ELECTRONIC ENGINE CONTROL (EEC)

ENGINE INTERFACE UNIT (EIU)

CHANNEL A

CHANNEL B

TCM OVER SPEED TORQUE MOTOR

SPOILER ELEVATOR


COMPUTER (1)

SPOILER ELEVATOR COMPUTER (2)


Thrust Control Malfunction Page: 31

AIRBUS A-- 320 PW 1100 NEO 76 -- 00

NOTES:





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00

AIR SYSTEM



POWER PLANT AIR SYSTEM


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00

AIR SYSTEM General The engine air system make sure that the compressor airflow and turbine clearances are controlled. The air system is used to improve engine operability and stability by bleeding air from the Low Pressure Compressor (LPC) and the High Pressure Compressor (HPC). External and internal ducts are used to achieve functions such as ccoling, pressurizing and ventilation airflows The air system also removes dirt, water, and hail from the LPC air stream and prevents the debris from entering the HPC. The air system supplies the necessary cooling airflow to keep the temperature of the engine compartments in limits. The air system increases compressor stability during starting, transient, and reverser thrust operation. The air system consists of the engine cooling system and the compressor control system.





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Air System Page: 3



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

COMPONENTS Component Location The air ejector valve control and actuator (upper) is located on the diffuser case at the 11 o’clock position. The air ejector control valve and actuator (lower) is located on the HPC at the 10 o’clock position. The HPC Stator Vane Actuator (SVA) primary unit is mounted on the HPC case on the right side of the engine at approximately the 2 o’clock engine position. The HPC SVA secondary unit is attached to the HPC and mounted on the HPC case located at the 7 o’clock engine position. The LPC SVA is located at approximately the 9 o’clock position. The turbine Active Clearance Control (ACC) air valve is located on the diffuser case at the 1 o’clock position. The buffer air valve solenoid is located at approximately the 4 o’clock position. The buffer air valve pressure sensor is attached to the Compressor Intermediate Case (CIC) bulkhead at the 3 o’clock position. The HPC bleed valve air pressure sensor is attached to the CIC bulkhead at the 2 o’clock engine position. Air ejector pressure sensor The air ejector valve solenoid (upper) is located on the HPC at the 12 o’clock position. The air ejector valve solenoid (lower) is located on the HPC at the 12 o’clock position. The LPC 2.5 bleed valve actuator is mounted at the rear of the CIC fire seal at approximately the 10 o’clock position. The buffer air heat exchanger is mounted on brackets attached to the diffuser case at the 11 o’clock position. The HPC bleed valve is attached to the diffuser case at the 1 o’clock position.


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Air Ejector Valve Page: 5


POWER PLANT AIR SYSTEM ENGINE AND ACCESSORY COOLING Engine Bearing Cooling a-- The bearing compartments are cooled using two passive engine bearing cooling systems. b -- Buffer/Ventilation System The main components of the buffer/ventilation system are the Low Pressure (LP) buffer shutoff valve, LP buffer shutoff valve solenoid, LPC check valve, Buffer Air Temperature Sensor (BATS), buffer air pressure sensor, buffer air manifold, external buffer air tubes, and internal buffer air tubes. Buffer air for the No. 1, 2, 3, 5, and 6 bearing compartments is supplied from the 3rd stage HPC and the LPC by the LP buffer shutoff valve which is controlled by the LP buffer shutoff valve solenoid. The LP buffer shutoff valve solenoid is operated by the buffer air pressure sensor. The BATS is installed to external tubing and a C--seal is used to provide a seal and prevent air leaks. The BATS provides a buffer air temperature signal to the Electronic Engine Control (EEC). Buffer air that makes an exit from the No. 3 bearing forward seal area then flows through the Low Pressure Turbine (LPT) shaft inner diameter to the No. 5 and 6 bearing compartment. c -- Engine Bearing Cooling System -- The main components of the cooling system are the buffer air heat exchanger, external cooling air tubes, and internal buffer air tubes. -- The No. 4 bearing compartment buffer air is supplied from the 3rd stage high compressor bleed plenum by the pneumatically controlled LP buffer shutoff valve. -- The high compressor supply air enters the buffer air heat exchanger through external plumbing. It then flows through the two--pass plate fin, air--to--air heat exchanger. -- Cooling air is supplied to the heat exchanger from the LPC discharge air (station 2.5 air). The station 2.5 air exits the heat exchanger and dischargers into the fan bypass. -- The cooled high compressor discharge air is then routed to the No. 4 bearing compartment by internal and external plumbing. -- At the bearing compartment, the cooled air flows around the compartment wall of the bearing housing to cool the bearing compartment.


AIRBUS A-- 320 PW 1100 NEO 75 -- 00 Turbine Active Clearance Control (ACC) System The ACC system is a dual channel EEC controlled system. This system meters fan air spray onto the turbine case through external manifolds to decrease High Pressure Turbine (HPT) and LPT blade tip clearance for better fuel efficiency. The ACC system actuator controls air flow to the HPT and LPT case cooling manifolds to minimize turbine blade tip clearance over the full engine operating range. The ACC valve is a fuel--actuated butterfly valve that controls cooling air to the turbine cases. The actuator lets fan bypass air cool and control the expansion of the turbine case to match the radial expansion of the rotor. Usually the ACC valve is closed at start and idle, partially open at take--off and climb, and fully open at cruise. The manifolds point fan air onto the turbine case to decrease turbine case growth because of thermal expansion. The HPT manifolds and the LPT manifolds share a single control air valve and collector. Fan bypass air enters the inlet duct through the Inner Fixed Structure (IFS) of the nacelle and is supplied to the ACC air valve. The EEC controls the system on a schedule that is based on engine parameters and altitude. The EEC sends a milliampere signal to the torque motor which controls the fuel muscle pressure to put the actuator shaft in position. This shaft operates to open and close the air valve to control the flow of fan air onto the turbine cases. Feedback of actuator position is reported back to the EEC by a single channel Linear Variable Differential Transducer (LVDT).

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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Buffer Air Valve and Heat Exchanger Page: 7



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

Turbine Cooling Air (TCA) System The TCA System supplies a continuous flow of cooling air to the HPT 2nd stage vanes, inter--stage HPT cavity, Turbine Intermediate Case (TIC) stator vanes (which include OD and ID cavities), LPT case outer cavity, and LPT rotor inter--stage cavities. The TCA System has four HPT cooling air tubes. The TIC/LPT has four main TCA tubes, eight TIC jumper air tubes, and three jumper air tubes that supply air to the LPT case. HPT Cooling Air The HPT TCA system supplies continuous flow of 6th stage bleed air to the HPT 2nd stage vanes for cooling. The HPT TCA system has four TCA air tubes that are installed approximately 90 degrees apart on the engine core. This air is controlled by plates which are located between the HPT case and TCA tubes. The cooling air flows into the hollow vanes and through passages that go out of the trailing edge of the vanes and on the inner and outer platforms. Cooling air is also supplied through the vanes to the inter--stage HPT cavity. TIC/LPT Cooling Air The TIC/LPT TCA tubes supply continuous flow of HPC 3rd stage bleed air to the TIC and LPT case for cooling. This air cools the TIC fairings and the inner and outer TIC walls. Cooling air flows to the TIC through TIC jumper tubes (eight each) located on the HPT case. This air is controlled by plates located between the TIC and the jumper tubes. There are an additional three jumper tubes connected to the TIC/LPT TCA tubes. These tubes supply cooling air to the LPT outer case. This air is also controlled by plates located between the LPT case and the LPT TCA jumper tubes.


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




TCC and TCA Page: 9



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

COMPRESSOR CONTROL SYSTEM Variable Stator Vane System The variable stator vane system is used to adjust the position of the HPC and LPC variable stator vanes during engine operation to optimize engine performance. The system consists of an HPC primary (master) SVA unit, an HPC secondary (slave) SVA unit, and an LPC SVA unit which are controlled by the EEC. The function of the HPC master and slave SVAs is to position the HPC Variable Inlet Guide Vanes (VIGVs), 1st, 2nd, and 3rd stage vanes for the HPC. The HPC master and slave SVAs are fuel actuated and control the HPC variable stator vanes position. The EEC controls the master SVA piston position through a dual channel torque motor which directs pressurized fuel to the actuator piston. The position of the HPC master SVA piston rod is sensed by a temperature compensated single channel Linear Voltage Differential Transducer (LVDT). The LVDT is used to send the piston position to channel A of the EEC. The position of the HPC slave SVA piston rod is sensed by a temperature compensated single channel LVDT. The LVDT is used to send the piston position to channel B of the EEC. The EEC can use the piston position of the HPC master SVA piston rod or the HPC slave SVA piston rod to control the system. If there is a power failure to the master SVA, the actuator puts the vanes in the open position for maximum airflow through the HPC. Vanes are scheduled based on N2 (Core) correct speed. The HPC variable stator vanes are usually closed during engine start up and engine idle, and open during aircraft takeoff. The HPC SVA piston changes the position of the variable stator vanes through a bell crank and linkages. The function of the LPC SVA is to position the LPC stator vanes through the LPC VIGV linkage. The LPC SVA is fuel actuated. The VIGVs, which are the first stator stage of the LPC (the LPC has three stages), are variable and are put in position by the SVA.


The position of the LPC SVA piston rod is sensed by a temperature compensated dual coil LVDT. The LVDT is used to send the piston position to the EEC. If there is a power failure to the SVA, the actuator puts the vanes in the open position for maximum airflow through the LPC. The LPC SVA piston changes the position of the variable stator vanes through a bell crank and linkages. Vanes are scheduled based on N1 (LPC) correct speed. Usually, the LPC variable inlet guide vanes are open during aircraft takeoff and closed during engine start up and idle.

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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




LPC Bleed Page: 11



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

Compressor Bleed Air System 2.5 Bleed Assembly The 2.5 bleed assembly is attached to the LPC assembly just forward of the CIC. The assembly has a 2.5 bleed valve ring, two sealing rings, linkage, and a bell crank which moves the valve ring to the opened or closed position. The 2.5 bleed valve actuator is used to actuate the 2.5 bleed valve assembly. The actuator is mounted at the rear of the CIC fire seal at approximately the 10 o’clock position. A connecting link transmits actuator motion to the 2.5 bleed valve assembly. The EEC controls the 2.5 bleed actuator position through a dual coil torque motor which controls a servo valve. The 2.5 bleed valve actuator position is sensed by temperature compensated dual coil LVDTs. The LVDTs transmit actuator position to each channel of the EEC. The 2.5 bleed valve actuator is connected to the EEC by an electrical harness and is hydraulically actuated (fuel). The actuator is scheduled based on N1 (LPC) correct speed. Usually, the valve is open at start and idle, and closed at aircraft takeoff. HPC Bleed Valve This valve is a two--position (open/close) inverted poppet--type valve which is spring--loaded to the open position. The bleed valve receives pressure from the HPC 6th stage bleed air for closing. The bleed valve bleeds air from the 6th stage high compressor during engine start. The valve is spring--loaded open and will close during engine start when the HPC 6th stage air pressure is large enough to overcome the spring force. HPC Bleed Valve Pressure Sensor The pressure sensor is located approximately at the 2 o’clock position on the CIC fire seal. The sensor provides feedback to the EEC. The sensor is used as a notification if the bleed valve failed to open and the engine core area is exposed to compressor air after engine start. The sensor is also used as notification if the bleed valve failed to close during engine start. The sensor measures the rise in pressure when the bleed valve is open during and after engine start--up and provides feedback to the EEC. The sensor is a dual channel absolute pressure sensor with a single connector. SCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




HPC Bleed Page: 13



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

AIR SYSTEM General The Air System controls engine airflow to perform five major functions: * pressurize and seal bearing compartments * cool engine parts * improve fuel efficiency * remove ingested debris from the airstream. * increase engine operability and stability The Air System is made up of five subsystems. Details are shown below, including air sources that are used to perform system functions. ENGINE BEARING COOLING SYSTEM The Engine Bearing Cooling System provides cooling buffer air to the engine main bearing compartments. It also supplies sealing air to prevent oil leakage. The system consists of the Buffer Cooling and Bearing Ventilation subsystems. Buffer Cooling System * Buffer Air Heat Exchanger BAHE * External cooling tubes Bearing Ventilation System * Low Pressure Buffer Shutoff Valve LPBSOV * Low Pressure Buffer Shutoff Valve solenoid * Low Pressure Compressor check valve * Buffer Air Pressure Sensor BAPS * Buffer Air Manifold BAM

Buffer Air Heat Exchanger (BAHE) The Buffer Air Heat Exchanger uses Station 2.5 bleed air to cool HPC 3rd Stage Air before its delivery to the No. 4 Bearing housing. The BAHE is attached to the diffuser case at 11:00. The Buffer Air Heat Exchanger is a sealed assembly mounted on brackets. The assembly contains a tube matrix consisting of a number of U--shaped tubes brazed into a tube plate. The tubes are held in place by stainless steel baffles. Operation 1. Station 2.5 bleed enters the Buffer Air Heat Exchanger to cool the 3rd Stage HPC air. 2. Station 2.5 bleed air exits the Buffer Air Heat Exchanger and discharges into the fan bypass. 3. Third Stage HPC air enters the Buffer Air Heat Exchanger and flows through the tube matrix and then enters the No. 4 Bearing compartment for cooling.

Buffer Cooling System The Buffer Cooling System cools the No. 4 Bearing compartment by sending 3rd Stage high compressor bleed air through the Buffer Air Heat Exchanger.


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Buffer Heat Exchanger and Valves Page: 15


POWER PLANT AIR SYSTEM Bearing Ventilation System The Bearing Ventilation System controls the flow of HPC Stage 3 air to maintain the proper air pressure at bearing compartment nos. 1, 2, 3, and 5/6, ensuring proper functioning of the carbon seals. During engine start and at high power settings, pressurized air is provided from LPC exit stage 2.5 bleed air. At low engine power settings, pressurized air is provided from the HPC 3rd Stage. Low Pressure Buffer Shutoff Valve (LPBSOV) The Low Pressure Buffer Shutoff Valve provides discrete (on/off) control of HPC 3rd Stage bleed air supply to the Bearing Ventilation System. The valve is located on the HPC split case at 1:00. Bearing compartments that house bearing nos. 1, 1.5, 2, 3, and 5/6 are cooled using the EEC--controlled Low Pressure Buffer Shutoff Valve. The LPBSOV consists of an actuator cover and housing that contains a spring and piston assembly. The valve is spring--loaded closed. It is actuated by HPC 3rd Stage servo pressure that is controlled by the Low Pressure Buffer Shutoff Valve solenoid. If the valve is removed on--wing or during a shop visit, the gasket and C--seal are replaced. Operation When HPC 3rd Stage servo pressure is applied, it pushes the spring loaded piston forward. In the actuator housing, the piston is connected to a shaft through an internal linkage that rotates the shaft when the piston moves forward or aft. A butterfly valve in the valve body is connected to the shaft, allowing it to open or close based on the piston position Low Pressure Buffer Shutoff Valve (LPBSOV) Solenoid The Low Pressure Buffer Shutoff Valve solenoid provides discrete (on/off) control of the HPC 3rd Stage servo pressure sent to the Low Pressure Buffer Shutoff Valve. The solenoid is located on the CIC firewall at 5:00. The solenoid is dual channel and controlled by the EEC. Operation When the solenoid is de--energized, the valve is closed, shutting off the flow of HPC 3rd Stage servo pressure. When the solenoid is energized, the valve is open, allowing HPC 3rd Stage servo pressure to flow to the LPBSOV. HPC SCL /JGB / REV.00 / Jun--2016

AIRBUS A-- 320 PW 1100 NEO 75 -- 00 3rd Stage servo pressure lines direct air from the HPC 3rd Stage to the shutoff valve solenoid, and from the solenoid to the Low Pressure Buffer Shutoff Valve. Low Pressure Compressor Check Valve The Low Pressure Compressor check valve prevents HPC 3rd Stage air from flowing into the LPC flow path while being used by the Bearing Ventilation System. Located inside the Buffer Air Manifold at 4:30, the check valve is not visible. The check valve is a passive device with two flappers that hang from the flapper shaft in the open position at engine start. Operation: LPC buffer air flows past the valve at low power settings. When the EEC commands the Low Pressure Buffer Shutoff Valve to open, the HPC 3rd Stage air forces the flappers to close, preventing backflow into the LPC. The flappers pivot on the flapper shaft due to gravity and air pressure Buffer Air Manifold (BAM) The manifold directs buffer air to the Bearing Ventilation System, including bearing compartment nos. 1, 2, 3, and 5/6. The Buffer Air Manifold is attached to the Compressor Intermediate Case at 4:00. The Buffer Air Manifold receives air directly from the LPC and from the 3rd Stage HPC from tubes attached to the Low Pressure Buffer Shutoff Valve and the CIC. Buffer Air Pressure Sensor (BAPS) The Buffer Air Pressure Sensor provides feedback to the EEC to validate the position of the Low Pressure Buffer Shut--off Valve (LPBSOV). The dual channel sensor is located downstream of the Buffer Air Manifold on the Compressor Intermediate Case at 3:00. The signal provided to the EEC confirms the position (open or closed) of the LPC check valve and Low Pressure Buffer Shutoff Valve.

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AIRBUS A-- 320 PW 1100 NEO 75 -- 00

LOO POWER SETTING


HIGH POWER SETTING


LPBSOV and Check Valve Page: 17

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

ACTIVE CLEARANCE CONTROL SYSTEM (ACC) General The Active Clearance Control System meters fan cooling air that is ducted from the nacelle thrust reverser door and sent to the turbine cases. The cooling air limits turbine case growth during thermal expansion, reducing HPT and LPT blade tip clearance and improving fuel efficiency. The ACC System includes: * Inlet duct * ACC valve and actuator * ACC collector * ACC manifolds for HPT and LPT * ACC air distribution tube assembly for LPT





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




ACC System Page: 19


POWER PLANT AIR SYSTEM Inlet Duct The inlet duct receives and directs fan bypass cooling air to the ACC valve and actuator. The inlet duct is located on the diffuser case at 1:00. The inlet duct is made of stainless steel. To prevent cooling air leakage, the outboard side of the duct has a seal land that contacts a bellow seal attached to the nacelle door. The inboard end of the duct is attached with a clamp to the ACC valve and actuator. Three connecting rods bolted to three brackets provide axial support for the inlet duct. The rods are bolted to the diffuser case at the inboard end and to the inlet duct at the outboard end. Operation Fan bypass air enters the inlet duct via the Inner Fixed Structure of the nacelle. ACC Valve and Actuator The ACC air valve is a fuel--actuated butterfly valve that regulates the flow of cooling air to the turbine cases. The air valve and actuator are located on the diffuser case at 1:00. The dual--channel valve and actuator are controlled by the EEC based on N2 speed and altitude. A fuel--actuated piston opens and closes the butterfly valve. The piston is attached via a link to the butterfly valve shaft. The valve is held in position relative to the shaft by a tapered pin that goes through a hole in the center shaft. The valve and actuator are attached to the ACC collector with a clamp and a seal that prevent leakage between the components. A single--channel LVDT is mechanically coupled to the actuator piston to provide an electrical feedback signal to Channel A of the EEC. Operation 1. During engine operation, the EEC sends electrical command signals to a dual channel torque motor that is part of the valve and actuator. 2. The torque motor uses the signals to direct pressurized fuel to either side of the actuator piston to achieve the commanded position. 3. The piston opens and closes the butterfly valve, sending bypass airflow to the HPT and LPT case cooling manifolds. The airflow cools and actively controls turbine case expansion to match the radial expansion of the rotor.


AIRBUS A-- 320 PW 1100 NEO 75 -- 00 During normal engine operation the valve is closed at start and idle, partially open at takeoff and climb, and fully open at cruise. In the event of electrical power loss the fail--safe mode of the valve and actuator is closed. ACC Collector The ACC Collector distributes fan cooling air to the ACC HPT and LPT manifolds. The collector is located at 2:00 on the HPT case. The stainless steel collector is attached to the HPT ACC manifolds with four bolts, and attached to the LPT ACC tube using a coupling and two packings. The packings prevent cooling air leakage. Operation Fan air is routed from the ACC valve into the collector and directed separately to the HPT and LPT manifolds.

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AIRBUS A-- 320 PW 1100 NEO 75 -- 00


SUPPORT RODS


ACC Inlet Duct and Valve Page: 21

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

ACC HPT Cooling Air Manifolds HPT ACC Manifolds HPT ACC manifolds receive cooling air from the ACC collector and distribute it around the inside of the manifolds. The manifolds are located around the diameter of the HPT case. Two sets of stainless steel manifolds (left and right) are installed on the outside of the HPT case in the same radial plane as the 1st and 2nd stage HPT blades. They are attached with brackets and bolts to the M and N flanges of the HPT case. Operation Cooling air exits the HPT ACC manifolds through small holes on the inner diameter, cooling the HPT case.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




HPT Cooling Manifold Page: 23



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

LPT ACC Manifolds LPT ACC manifolds supply cooling air from the ACC collector to the outside of the LPT case. The manifolds are located around the diameter of the LPT case. The stainless steel manifolds are installed onto 24 studs in the LPT outer case and secured with 24 nuts. Each manifold has an integral tube stand--off that mates with a manifold connector on the LPT Air Distribution Tube Assembly.


LPT Air Distribution Tube Assembly The LPT Air Distribution Tube Assembly receives cooling air from the ACC collector and distributes the air onto the LPT case. The assembly is attached to the LPT case. The assembly consists of three separate sections, six tube couplings and four manifold connectors. One of the three tube sections has an integral T connector that receives air from the ACC collector. Tube couplings connect the tube sections to the manifold connectors. The couplings fit over the outside diameter of the tube sections and manifold connectors, and are secured with 12 clamps. The tube sections are attached to the LPT case with seven clamps installed on seven brackets attached to the case studs.


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




LPT Cooling Manifold Page: 25

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

TURBINE COOLING AIR The Cooling Air System is a passive system yhat provides a continuous flow of cooling air inside the turbine cases. The Turbine Cooling Air System provides continuous cooling air to the High Pressure Turbine, and to the Turbine Intermediate Case/Low Pressure Turbine. The system consists of 19 external tubes or jumpers that direct calibrated HPC bleed air 3nd and 6th stages to the the following engine parts: * HPT 2nd Stage vanes between the 1st and 2nd stage rotors * HPT 2nd Stage blade attachment * Turbine Intermediate Case (TIC) fairings * LPT case * LPT rotor and blade attachments





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




TCA System Page: 27

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

TCA SYSTEM High Pressure Turbine Cooling Air Tubes HPT cooling air provides continuous airflow of 6th Stage bleed air to the HPT 2nd Stage vanes. The airflow tubes are located approximately 90 apart at these positions: 1:00, 5:00, 7:00, 10:00. The system consists of four TCA air tubes that provide cooling airflows into the hollow vanes, and through passages that exit out of the trailing edges of vanes and vane platforms. The air is metered by plates located between the HPT case and tubes.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




TCA Nozzles Page: 29

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

Turbine Intermediate Case/Low Pressure Turbine Cooling Air Tubes TIC/LPT cooling air provides continuous flow of HPC 3rd Stage bleed air to the Turbine Intermediate Case and LPT case and rotor. Four TIC cooling air tubes cool the TIC fairings, as well as the inner and outer TIC walls known as transition ducts. The tubes are positioned at 3:30, 6:00, 8:00 and 11:00. Eight jumper tubes are installed around the radius of the TIC fairing and three more are located on the LPT case at 12:00, 4:30 and 9:00, this jumper tubes that feed off the four cooling tubes send air through the TIC connecting rods to the Low Pressure Turbine rotors and blade attachments.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




TCA Ducts and Jumpers Page: 31

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

Low Pressure Turbine Cooling Air Tubes The LPT TCA jumper tubes direct cooling air between the LPT outer case and Second Stage vanes. The three jumper tubes are located on the LPT case at 12:00, 4:30 and 9:00.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




TCA Jumpers Page: 33



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

COMPRESSOR AIRFLOWS Compressor Variable Vane Control The Compressor Variable Vane Control System uses actuators to move the HPC and LPC Variable Inlet Guide Vanes (VIGVs), adjusting the angle of airflow required for optimal engine operation. The actuators receive commands from the EEC and are positioned hydraulically using pressure fuel (PF) from the Integrated Fuel Pump and Control. Sub--systems include: * Low Pressure Compressor Variable Inlet Guide Vane Control * High Pressure Compressor Variable Vanes System. Both the LPC and HPC Variable Inlet Guide Vanes use actuators to change vane positioning via a bellcrank and linkages. The vanes in the LPC and HPC are commanded by the EEC and use fuel pressure to maintain engine stability. The EEC controls the vanes using schedules based on rotor speeds. It also receives vane position feedback from Linear Variable Differential Transducers (LVDTs) mounted to the actuators. VIGVs for the LPC are positioned by a schedule based on N1 (LPC) speed. VIGVs for the HPC are positioned by a schedule based on N2 (HPC) speed. Components for the LPC Variable Inlet Guide Vane Control system include: * stator vane actuator * bellcrank * linkage * connecting rod. Components for the HPC Variable Vanes System include: * stator vane actuators (2) * primary * secondary * bellcrank * linkages (4).


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Compressor Airflows Page: 35



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

LPC Variable Inlet Guide Vane Control LPC Stator Vane Actuator (LPC SVA). The LPC SVA positions LPC variable inlet guide vanes through the LPC SVA linkage when commanded by the EEC. The LPC SVA is mounted in the Compressor Intermediate Case fire containment ring on the left side of the engine at approximately 9:30. A dual--channel Linear Variable Differential Transducer (LVDT) is mechanically coupled to the actuator piston to provide electrically isolated position feedback signals to each channel of the EEC. Operation During engine operation, the EEC sends electrically isolated drive signals to the dual channel torque motor that is part of the LPC SVA. The drive signals direct pressurized fuel to either side of the actuator piston to achieve the commanded position. If there is a loss of electrical power to the torque motor, the actuator positions the vanes in the full open fail safe position for maximum airflow through the LPC. The actuator has a fuel drain for internal component leakage. LPC Variable Inlet Guide Vane (VIGV) Linkage The LPC VIGV linkage translates the axial position movement of the LPC Stator Vane Actuator into circumferential synchronizing ring movement in order to position the guide vanes. The linkage is located on the LPC case at 9:30. The LPC SVA piston is attached the LPC connecting rod by a bolt through a clevis end. The forward end of the LPC connecting rod is bolted to the LPC bellcrank assembly. The LPC bellcrank assembly is bolted to the bellcrank bracket. The bracket is attached to the Fan Intermediate Case by four bolts. Operation 1. Upon receiving a command from the EEC, the LPC Stator Vane Actuator directs pressurized fuel to the appropriate side of the piston to move the LPC connecting rod. 2. The forward end of the LPC connecting rod moves the LPC bellcrank assembly, which pivots about the bellcrank bracket. This translates the axial movement of the LPC SVA into circumferential movement of the synchronizing rings. The rings move the 61 VIGVs in unison to the correct position.


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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




LPC Stator Vane and Actuator Page: 37

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

High Pressure Compressor Variable Vanes HPC Variable Stator Vane Actuators work in unison to position HPC Variable Inlet Guide Vanes and Variable Stator Vanes in response to EEC commands that optimize engine performance. The system has primary and secondary stator vane actuators and adjusts the vanes using the HPC VIGV and VSV linkage. HPC Primary Stator Vane Actuator The HPC primary Stator Vane Actuator positions the inlet guide vanes of the Compressor Intermediate Case and the 1st, 2nd and 3rd variable vanes of the HPC. The primary actuator is mounted on the HPC case on the right side at 2:00. The primary SVA is a dual--channel, EEC--controlled valve with a fuel actuated piston that moves the HPC bellcrank assemblies. Piston position feedback is provided to Channel A of the EEC. Operation During engine operation, the EEC sends electrically isolated drive signals to the dual--channel torque motor that is part of the primary SVA. The torque motor uses the electrical signals to direct pressurized fuel to either side of the actuator piston to achieve the commanded actuator position. A single channel Linear Variable Differential Transducer (LVDT) is mechanically coupled to each actuator piston to provide a positional feedback signal to the EEC. If there is a loss of electrical power to the torque motor, the actuator positions the vanes to the full open fail--safe position for maximum airflow through the HPC. The EEC can use either the primary or secondary actuator LVDT feedback signals to control the system.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




HPC VSV and Primary Actuator Page: 39

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

HPC Secondary Stator Vane Actuator The HPC secondary Stator Vane Actuator positions the inlet guide vanes of the Compressor Intermediate Case and the 1st, 2nd and 3rd variable vanes of the HPC. The secondary actuator is on the left side of the HPC at 9:00. The secondary SVA is a dual--channel, EEC--controlled valve with a fuel actuated piston that moves the HPC bellcrank assemblies. Piston position feedback is provided to Channel B of the EEC through the secondary SVA’s Linear Variable Differential Transducer (LVDT). Operation Pressurized fuel from the primary SVA is routed via a fuel supply and fuel return tubes to the HPC secondary SVA, positioning the HPC secondary actuator piston. The EEC uses primary and secondary actuator LVDT feedback signals to control the system.





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AIRBUS A-- 320 PW 1100 NEO 75 -- 00




HPC VSV and Secondary Actuator Page: 41

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

COMPRESSOR BLEED AIR SYSTEM General The Compressor Bleed Air System improves engine operability and stability by bleeding air from the Low Pressure Compressor. The system also removes debris from the LPC air stream. The Compressor Bleed Air System consists of the following: * 2.5 Bleed Valve Air System Assembly * HPC Bleed Air System. The compressor bleed control comprises one LPC Bleed Valve Actuator (BVA) and two HPC Bleed Valve. The LPC bleed system is used to control the LPC discharge 3nd stage airflow into the fan discharge. The EEC modulates LPC BVA ellectrically controlled by dual torque motor and a fuel operated EHSV. The actuator LVDT transmits the npiston position to each EEC channel individually. The HPC Bleed system is used to control the HPC 6th stage airflow into the core area. The system has two on--off HPC bleed valves; one is active and the other is passive, and both spring loaded open and pneumatically closed at certain engine operating conditions. The active valve is EEC controlled closed through the HPC bleed valve solenoid thanks to Ps3 pressure. The passive valve closes when the pressure inside the HPC is high enough to force the spring loaded valve closed. Both are monitored by the EEC thanks to two dedicated pressure sensors.





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Engine Bleed System Page: 43

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

2.5 BLEED VALVE AIR SYSTEM ASSEMBLY Description The 2.5 Bleed Valve Air System Assembly discharges LPC exit airflow into the fan bypass airstream. Assembly components are shown below. * Bleed ring * Actuator * Bellcrank * Linkages 2.5 Bleed Valve Actuator The 2.5 bleed valve actuator controls the LPC bleed valve through the 2.5 bleed valve linkage when commanded by the EEC. The actuator is mounted at the rear of the CIC fire containment ring at approximately 9:30. A dual channel Linear Variable Differential Transformer (LVDT) is mechanically coupled to the actuator piston to provide position feedback signals to each channel of the EEC. Operation During engine operation, the EEC sends electrical signals to a dual channel torque motor that is part of the bleed valve actuator. The torque motor uses the electrical signals to direct pressurized fuel to either side of the actuator piston to achieve the commanded position. If there is a loss of power, the actuator positions the bleed valve in the full open fail safe position for maximum bleed air flow.





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2.5 Bleed and Actuator Page: 45

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

2.5 Bleed Valve Linkage The 2.5 bleed valve linkage translates axial movement of the 2.5 bleed valve actuator piston to circumferential movement to open and close the bleed valve. The linkage is located at 12:00 on the LPC case. The bleed valve connecting link is bolted to the 2.5 bleed valve actuator piston at one end and to the bleed valve bellcrank at the other end. The bleed valve bellcrank is also attached to the bleed valve by one bolt and is fastened to the bellcrank support bracket with one bolt and washer. The bellcrank support bracket is mounted to the LPC outer case by two bolts. Two fabric coated, silicone seal rings are installed on the bleed valve to provide sealing when the bleed valve is in the closed position. Idler links support the bleed valve around the circumference of the CIC. Operation Forward and aft movement of the actuator piston is transmitted by way of the bleed valve bellcrank to the bleed valve. This force moves the bleed valve in a spiral motion between an open or closed position, regulating LPC bleed air out the 2.5 bleed ducts in the Compressor Intermediate Case. The LPC bleed air is discharged into the fan bypass airstream.





Page: 46



AIRBUS A-- 320 PW 1100 NEO 75 -- 00


2.5 BLEED SLOTS


2.5 Bleed Components Page: 47



AIRBUS A-- 320 PW 1100 NEO 75 -- 00

HPC BLEED AIR SYSTEM Description The HPC Bleed Air System bleeds 6th Stage HPC air to improve engine startup performance. The system has both active and passive components as shown in the chart.


HPC Passive Bleed Valve The HPC passive bleed valve is a spring--loaded valve that allows HPC 6th Stage air to bleed directly into the core compartment during engine start to help with initial compression of upstream core air flow. The passive bleed air valve is located on the HPC case at 1:00. The HPC passive bleed air valve is attached to the outer diffuser case boss by four bolts. A gasket is installed between the HPC bleed valve and the diffuser case to prevent air leakage. Operation The spring forces the bleed valve open when the pressure within the High Pressure Compressor is low. When sufficient pressure is developed in the High Pressure Compressor the valve is forced closed. HPC Active Bleed Valve The HPC active bleed valve is a spring--loaded valve that allows HPC 6th Stage air to bleed directly into the core compartment during engine start to help with initial compression of upstream core air flow. The active bleed air valve is located on the HPC case at 3:30. The HPC active bleed valve is attached to the outer diffuser case boss by four bolts. A gasket is installed between the HPC active bleed valve and the diffuser case to prevent air leakage. Operation The passive HPC bleed valve closes at sub--idle and the active HPC bleed valve is opened with Station 3 air supplied by the HPC active solenoid valve. At observed idle, the EEC will command the solenoid closed, shutting off P3 air. The active HPC bleed valve closes with Stage 6 HPC air.


Page: 48


AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Bleed Air System Page: 49

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

HPC Passive and Active Bleed Valve Air Pressure Sensors The dual--channel HPC passive bleed valve air pressure sensors measure the outlet air pressure on the HPC bleed valves. Both sensors are located on the CIC. The passive sensor is located at approximately 10:00 and the active sensor is at approximately 2:00. The pressure sensors consist of two independent, electrically isolated sensing elements, a stainless steel body, and an electrical connector. The components are assembled as a hermitically sealed unit. The stainless steel body has a mounting flange and houses the sensing elements. Operation HPC bleed valve sense lines direct pressurized air from the HPC bleed valves to the HPC bleed valve pressure sensors sensing elements. Each sensing element consists of a diaphragm with strain gages. When pressure is applied, the strain gages change resistance, which changes the output voltage. This output voltage correlates directly to air pressure. Each sensing element is connected to the electrical connector and sends the air pressure signal to the EEC over separate channels.





Page: 50



AIRBUS A-- 320 PW 1100 NEO 75 -- 00 PASSIVE VALVE CONNECTOR J53 ACTIVE VALVE SENSOR CONNECTOR J56


HPC 6TH STAGE AIR WHEN IS OPEN


Passive and Active Pressure Sensor Page: 51

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

HPC Active Solenoid Valve The HPC Active Solenoid Valve provides discrete (on/off) control of HPC 3rd Stage servo pressure sent to the HPC active bleed air valve. The active solenoid valve is located at 2:30 on the combustor case. The dual--channel HPC active solenoid valve controls the flow of HPC 3rd Stage servo pressure to the HPC active bleed valve. The active solenoid valve is controlled by the EEC. Operation When the solenoid is de--energized the valve is closed, shutting off the flow of HPC 3rd Stage servo pressure. The valve is open when energized, allowing HPC 3rd Stage servo pressure to flow to the HPC active bleed valve and open it.





Page: 52


AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Active Solenoid Valve Page: 53

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

COOLING COMPARTMENT Description The compartment cooling system ensures the ventilation fan compartment, core compartment and dedicated component inside the core compartment. The cooling of the fan compartment is achieved through a passive ventilation system. Outside airflow circulates from the top scoop around the fan case and exhaust through bottom holes and gabs of the fan cowls. The cooling of the core compartment is achieved trhough a passive ventilation system. Fan bypass airstream is directed to the nacelle core, ignition leads, igniter plugs and Enviromental Control System (ECS) bleed valves through openings on the inner contour of the thrust reversercowl doors and exhaust through bottom holes and gabs of the Inner Fixed Structure (IFS) trailing edge. Additional tubes are dedicated for the cooling of the ACC Valve, Starter Air Valve (SAV) and the Flow Divider Valve (FDV).





Page: 54


AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Cooling Compartment Page: 55


AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Air System Table Page: 56


AIRBUS A-- 320 PW 1100 NEO 75 -- 00




Page: 57

AIRBUS A-- 320 PW 1100 NEO 75 -- 00

NOTES:





Page: 58

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

ENGINE INDICATING



POWER PLANT ENGINE INDICATING


Page: 1

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

ENGINE INDICATING General The engine indicating system is used to sense, transmit, and display engine operation parameters and component/system data collection. It also is used to transmit information about above limit conditions to the cockpit warning gauges. These parameters are supplied to the Electronic Engine Control (EEC) and / or to the Prognostic and Healt Monitoring Unit (PHMU) for computation and transmission. They are sent to the Electronic Instrument System (EIS) for display on the EWD and on the SD-- Engine page. In conjuction with inputs from the ADIRS, they are also used to control and minitoring the engine with the throttle Lever Angle (TLA) position in manual thrust control mode or with the Engine Interface Unit (EIU) inputs in autothrust mode





Page: 2


AIRBUS A-- 320 PW 1100 NEO 77 -- 00




Engine Indicating System Page: 3



AIRBUS A-- 320 PW 1100 NEO 77 -- 00

POWER INDICATING SYSTEM The power section of the engine indicating system is related to low rotor (N1) speed, the fan rotor speed (Nf) and the high rotor speed (N2). N1 Power Indicating Shaft speed indicating is used to sense and transmit low rotor (N1) speed to the EEC. The primary engine thrust control parameter is the low rotor speed. The N1 speed probe is installed in the CIC, which is station 2.5. The probe is a dual channel sensor that has a magnet stack with two isolated coils and electrical connectors. The probe is attached with two bolts. Low rotor speed is sensed by the dual element, N1 speed sensor. The probe sends an N1 speed signal through the EEC to the Engine Indicating and Cockpit Awareness System (EICAS). EICAS displays the N1 speed signal as a virtual gauge on the aircraft flight deck’s main instrument panel. Each channel of the N1 speed probe can command the thrust level through adjustment of the N1 speed. The EEC controls all engine subsystems to get the necessary N1 speed and keep the engine operation in all approved limits. The N1 speed probe makes a waveform that is positive as the target tooth gets near and negative as it goes by. Each one of the 20 teeth on the No. 2 bearing coupling nut, which goes in front of the probe, causes a sudden change in the magnetic field. The probe gives an oscillation signal which is a signal that is in proportion to the speed that the coupling turns. The probe, if correctly installed, has a minimum space between the sensor and the exciter teeth to let a strong signal be permitted and also prevent rub. The maximum limit (or overspeed warning) is set to N1 100%. When the indicator shows N1 100%, the display changes to red, the master caution light illuminates, and a single chime goes off.


Page: 4

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

N1 SPEED SENSOR





N1 Speed Sensor Page: 5

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

Nf -- Shaft Speed Indicating Shaft speed indicating is used to sense and transmit fan rotor (Nf) speed to the EEC. The Nf speed probe is installed to the No. 1/1.5 bearing support. The probe is a one--channel sensor that has a magnet stack, coil, and electrical connector. The probe has a two--bolt flange with jackscrew holes to remove the probe. Fan rotor speed is sensed by the one--element N2 speed probe. The probe sends an Nf speed signal through the EEC to the PHMU where it is used with the N1 vibration pickup to measure fan vibration. This speed signal and vibration data is used for on--wing trim balance. The Nf speed signal is not displayed in the cockpit.





Page: 6

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

B





Nf Speed Sensor Page: 7

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

N2 -- Shaft Speed Indicating Shaft speed indicating is used to sense and transmit high rotor (N2) speed to the EEC. The EEC uses the N2 speed signal for control of fuel and ignition during start, and also to monitor an overspeed condition of the high rotor. The N2 speed transducer is a speed coil that is located on the AGB which transmits the torque from the high rotor to the Main Gearbox (MGB). The N2 speed transducer is dual channel. It makes two isolated electrical signals, with frequencies that are related to N2 speed by N2 rpm. The speed transducer sends an N2 speed signal through the EEC to the EICAS. EICAS displays the N2 speed signal as a three--digit digital value in the lower level engine indications are on the aircraft flight deck’s main instrument panel.





Page: 8



AIRBUS A-- 320 PW 1100 NEO 77 -- 00


C


N2 Speed Page: 9



AIRBUS A-- 320 PW 1100 NEO 77 -- 00

Temperature Indicating System The temperature section of the engine indicating system is related to engine exhaust gas temp (EGT), HPC discharge temperature (T3), and core compartment temperature. The EGT indicating system has four thermocouple probes located around the circumference of the TEC at semi--regular intervals. The four probes are located at the two, four, eight and ten o’clock engine positions. the EGT harness leads connect to each of the four individual EGT porbes. The probes monitor the temperatures of the low pressure turbine gaspath exit temperature. The EGT signal is received by the EEC and is then synthesized and sent to the cockpit display. The EGT virtual gauge in the cockpit displays a virtual needle and a four--digit digital parameter (in degrees Celsius). The display also has an amberline and redline limit. The display changes to amber or red given the measured EGT and the approved limits. The four signals from the probes are electrically averaged into two signals which are sent to the EEC through the EGT (T5/T3) wiring harness (WC05). The two signals from the left side are averaged into the Channel A signal. The two signals from the right side are averaged into the Channel B signal. When received by the EEC, the EEC changes the averaged EGT analog signals to digital signals. The digital signal is then transmitted to the flight deck on the ARINC 429 data bus, where it is displayed in degrees Celsius. Each probe has two studs of different dimensions and a stud/lug design with a captive nut and anti--rotation feature. The probes are Type K thermocouples. There are two sets of K+/K-- wires in the harness that transmit the signals to the cold junction reference located in the Main Oil Temperature Sensor.


Page: 10



AIRBUS A-- 320 PW 1100 NEO 77 -- 00

D


D


EGT Probes Page: 11

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

FUEL PARAMETER General The Fue! Flow Meter (FFM) is installed on the intennediate case right hand side ofthe eugine core at approximately the 3 o’clock position. The FFM is a single element device wired to Channel A of the EEC. The signal is hardwired to Channel B internal to the EEC. The FFM is an in--line sensor located between the IFPC and the Flow Divider Valve. The fuel flow and the fue lused are displayed on the ECAM EWD by digital indications. The FFM is a magnetic drum and impeller type. The fuel used value is computed by the EIS from the fuel flow value sent by the EEC. The fuelused for each engine is computed from the engine start to the engine shutdown. It is reset to 0 at the next engine start. The FFM is an LRU.





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AIRBUS A-- 320 PW 1100 NEO 77 -- 00




Fuel Flow Meter Page: 13



AIRBUS A-- 320 PW 1100 NEO 77 -- 00

OIL PARAMETERS General The Oil Leve! (OL) sensor is located in the oil tank. It sends the oil quantity analog signal to the EEC. The EEC sends the signal for display on ECAM SD ENGINE page . The Main Oil Pressure (MOP) sensor is located on the left hand side of the engine on the Oil Control Module (OCM), rear lower side. lt is a dual channel sensor which sends the signal to the EEC for monitoring. EEC sends the signal for display on SD ENGINE page. The Main Oil Temperature (MOT) sensor is a dual channel sensor and is used to measure the temperature of the scavenge oil returning to the tank This data is monitored by the EEC and is displayed on the SD ENGINE page. The sensor is located on the front face of the OCM. In case of abnonnal condition, sensors send signals to trigger messages on ECAM and / or CFDS. An Oil Filter Differential Pressure (OFDP) sensor is installed adjacent to the oil pressure filter unit on the Lubrication and Scavenge Oil Pump (LSOP) unit. The pressure sensor signal is transmitted by the EEC to thc ECAM system to generate the main oil filter clogging alerts when the oil differential pressure across this filter exceeds the thresholds. Two indications are available: -- DEGRAD or -- CLOG. An Auxiliary Oil Pressure (AOP) sensor is located on the left side of the engine. below the Variable Oil Reduction Valve /Joumal Oil Shuttle Val ve (VORV / JOSV). It measures the pressure of oil delivered to the joumal bearings in the Fan Drive Gear System (FDGS). lt sends a signal to the EEC, where it is used in conjunction with other oil parameters to detecta Fan Drive Gearbox (FDG) auxiliary oil supply malfunction. The Low Oil Pressure (LOP) switch signals the EIU when the oil pressure drops below a threshold. lt is located on the left hand side ofthe engine on the Oil Control Module (OCM).


Oil Debris Monitoring. The Oil Debris Monitoring (ODM) sensor is located on the top front side of the oil tank. It sends signals proportional to size and typc of the pollution particles to the PHMU The PHMU monítors the debris for quantity and identifies whether it is ferrous or non--ferrous debris. The data is transmitted to the EEC for analysis and to generate an ECAM message a the trend monitoring accordingly. The data is also stored in the Data Storage Unit (DSU).

Page: 14



MAIN OIL TEMPEREATURE (MOT)

AIRBUS A-- 320 PW 1100 NEO 77 -- 00 LOW OIL PRESSURE (LOP)

OIL LEVEL SENSOR (OL)

MOT

MAIN OIL PRESSURE SENSOR (MOP)

MOP

LOP


LSOP

AUXILIARY OIL PRESSURE SENSOR (AOP)

LUB AND SCAVENGE OIL PUMP (LSOP)


Oil Parameters Page: 15

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

NACELLE TEMPERATURE Description The naceelle temperature is monitored by a temperature probe installed in the ventilated core compartment. The nacelle temperature is displayed on the ECAM ENGINE SD, exccpt during starting or cranking sequences where it is replaced by starting paramcters. The T--core probe senses the temperature in the core compartment between the nacelle and engine case. The signal is transmitted to the EEC. When the core compartment temperature exceeds the core compartment advisory temperature limit the appropriate signal is sent to the ECU for display in the cockpit.





Page: 16


AIRBUS A-- 320 PW 1100 NEO 77 -- 00




Nacelle Temperature Page: 17

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

ANALYZERS SYSTEM General The analyzers system consists of the prognostic system and vibration monitoring system. The PHMU receives signals from the ODM, engine accelerometers, and other engine sensors and sends prepared output signals to the EEC. The EEC then sends the correct output to interface with aircraft displays and computers. The PHMU has the capability to monitor the following: -- Vibrations -- Oil Debris -- Anomaly Detection Overview The Analyzer System provides critical information about oil debris monitoring, vibration monitoring, and auxiliary oil pressure to the EEC and the flight deck. A primary component of this system is the Prognostics and Health Management Unit (PHMU). The chart shows the roles of the Analyzer System and how the PHMU functions to fulfill each one. The PHMU continuously computes engine trim balance solutions using Nf, N1, and N2 speed signals received from the EEC, and from the vibration signals received from the aft and forward accelerometers. This information is stored by the PHMU in the Data Storage Unit (DSU). A trim balance procedure in the cockpit interprets the stored data using the interactive mode, and provides instructions to trim balance the fan.

Analyzer Roles PHMU Functions Oil Debris Monitoring

Vibration Monitoring

Auxiliary Oil Pressure Monitoring

* Provides early indications of ferrous and non ferrous particles in engine oil, based on data from the Oil Debris Monitor * Detects exceedances based on the rate of particles generation over a period of time * Detects vibration exceedances based on data f from engine i vibration ib ti sensors, plus l N1, N1 N2 and d Nf inputs from the EEC. * Calculates optimum fan trim balance solutions using flight or ground data. * Indicates low oil pressure in Auxiliary Oil System





Page: 18

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

G





PHMU Page: 19

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

PHMU FUNCTIONS Vibration Monitoring The PHMU receives signal analysis from engine accelerometers to assess the health of rotating and mechanical components. The PHMU reports this information to the aircraft as maintenance messages and cockpit caution, status, and warning messages if necessary. The Forward Vibration Sensor is attached to the Compressor Intermediate Case, located at the 3 o’clock position. The Aft Vibration Sensor is attached to the low pressure turbine case, located at the 3 o’clock position.





Page: 20



AIRBUS A-- 320 PW 1100 NEO 77 -- 00


I


FWD Vib Sensor Page: 21



AIRBUS A-- 320 PW 1100 NEO 77 -- 00


J


FWD and AFT Vib Sensors Page: 22



AIRBUS A-- 320 PW 1100 NEO 77 -- 00


MCDU

SD Figure 12 SCL /JGB / REV.00 / Jun--2016

Indications Page: 23



AIRBUS A-- 320 PW 1100 NEO 77 -- 00

PHMU FUNCTIONS Anomaly Detection The PHMU compares engine data to models of engine gas path and subsystem parameters. The information is prepared by the PHMU and the applicable output is sent to the EEC. The EEC then sends this information to the aircraft computer as maintenance messages. The prepared ODM signal will enhance maintenance operations that provide an early indication of debris events which help the magnetic chip collector plug inspections that are required to troubleshoot. The prepared vibration signal will provide the applicable signal for cockpit indication during rotor imbalance. It will also provide additional vibration information for ground maintenance purposes. Processed signal information for anomaly detection is also used for ground maintenance purposes. Other Sensors Various sensors are used by thc EEC for the engine control and monitoring. The T2 sensor measures the air inlet temperature for engine rating, Mach number calculation and bleed scheduling. The P 2.5P/T 2.5 sensor measures the air pressure and temperature downstream of the booster at the High Pressure Compressor (HPC) inlet. It is located on the Compressor lntermediate Case at 1 o’clock position . The Bumcr Pressure (PB) sensor measures the pressure related to the combustion for fuel scheduling, surge recovety, stall detection, idle modulation and continuous ignition logic. lt is located in the LH side Compressor Intermediare case firewall at 11 o’clock position. The T3 sensor measures the compressor discharge temperature for total temperature calculation. It is located on the diffuser case, forward of the fuel nozzles at 1 o’clock position.


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AIRBUS A-- 320 PW 1100 NEO 77 -- 00




Other Sensors Page: 25

AIRBUS A-- 320 PW 1100 NEO 77 -- 00

NOTES:





Page: 26

AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

IGNITION & STARTING SYSTEM



POWER PLANT START AND IGNITION


Page: 1



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

IGNITION AND STARTING SYSTEM General The Ignition system provides the electrical spark needed to satrt or continue engine combustion. The ignition system is made up of two independent systems, includes an ignition exciter, two coaxial shield ignition leads and two igniter plugs. The Starting sytem drives the engine HP rotor at speed high enough for a ground or in flight start to be initiated, the system includes the electrical Start Air Valve (SAV) and the pneumatic starter. Air bleed is taken from the aircraft pneumatic system for engine start, APU bleed, external pneumatic cart or other engine bleed. Control The EEC controls the SAV and ignition during automatic or manual start. 115 Vac from aircraft electrical systemis supplied to the ignition exciter wich provides the necesary voltage to the igniter plugs to generate the saprk for combustion. Indication The SAV and Ignition system are displayed on the ENGINE ECAM page. The main parameters to be monitored during starting are displays on the E/WD such as N1, EGT, N2 and F/F, and on the SD Oil Press, IGN system, SAV position and available pneumatic pressure.

Note: there is not automatic shutdown function or second attemp in MANUAL START Cranking Engine motoring could be performed for dry or wet cranking sequences. Note: during cranking the ignition is inhibited. Continuous Ignition With engine running, continuous ignition can be selected via the EEC either manually using the rotary selector or automatically by FADEC during specific conditions. Safety Precautions Safety precautions have to be taken prior to working in this area. WARNING:

THE IGNITION EXCITER PROVIDES HIGH ENERGY PULSES THROUGH THE IGNITION LEADS TO THE TWO IGNITER PLUGS.

Maintenance To increase A/C dispatch reliability, the SAV is equipped with a manual override. For this manual operation, the technician has to be aware of the engine safety zones.


STARTING MODES Automatic Start During an automatic satrt, the EEC opens the SAV to drive the engine and the ignition is then energized when the HP rotor speed is sufficient for that. The EEC provides full protection during the start sequence. When the automatic start is completed, the EEC closes the SAV and cutoff the ignition. In case of a incident during the automatic start the EEC makes a second attemp or aborts the start procedure. Manual Start During a manual start, the SAV opens when the MANual START P/B is pressed in, then the ignition system is energized when the MASTER control lever is set to the ON position. SCL /JGB / REV.00 / Jun--2016

Page: 2


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Page: 3



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

IGNITION SYSTEM General The ignition system supplies a high energy spark to ignite the fuel/air mixture in the combustor. The ignition system is composed of the ignition exciter, two ignition cables, and two igniter plugs. The system can provide adequate spark rates for burner ignition throughout the engine ignition envelope in continuous duty. Redundant ignition systems are incorporated into the design. The EEC automatically controls the ignition systems. During engine start, one or both ignition systems are energized at the same time as fuel flow is turned on. During an auto--start on the ground, the EEC selects one ignition system (System A or System B) for two starts in a row (one on Channel A and one on Channel B), then for the next two starts in a row, the other ignition system is selected. The ignition exciter is installed on the fan case assembly at the 5 o’clock position. The ignition cables are connected at one end to the ignition exciter attached to the fan case. The other end of the ignition cables are connected to the igniter plugs. The igniter plugs are installed on the diffuser case at the 3 o’clock position. System Description The ignition system is used at engine start, engine relight, and each time ambient conditions require a continuous ignition to prevent the risk of flame out. Each engine has two electrically independent ignition systems. Each system can be used at the same time or alternately to prevent dormant failure. When the manual ignition selection is used, the EEC energizes one or both igniters based on the logic and aircraft inputs of flap position, cowl anti--ice status and/or continuous ignition commands. The EEC includes an automatic ignition system as an integral part of the EEC’s ignition function. The EEC includes an automatic relight system which energizes both igniters within two seconds of when an engine flameout has been detected.


Page: 4


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Igniter Box Page: 5



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

IGNITION COMPONENTS Ignition Exciter The ignition exciter changes 28 VDC aircraft electrical power from the aircraft to high tension voltage necessary for igniter plug operation. The ignition exciter is attached (with anti--shock mounts) to a bracket on the fan case at the 5 o’clock position. The ignition exciter supplies 5 kilovolts to the igniter plugs to ignite the fuel/air mixture in the combustor. The electronic ignition exciter has an energy delivery capacity of 0.9 joules/ channel. Fan compartment air flows over the hermetically sealed exciter box to provide cooling for the unit. The exciter duty cycle varies between 1 spark/sec. to 3 spark/sec. An EEC command voltage activates each exciter channel. The EEC can command either exciter channel from both exciter inputs. Ignition Cable The ignition cables are used to transmit 5 kilovolts from the exciters to the igniter plugs. The cables are connected at one end to the ignition exciter and at the other end to the igniter plugs. The ignition cables are flexible braided steel conduits with ceramic insulated terminals at the plug end and are interchangeable. Igniter Plug The igniter plugs use the 5 kilovolts from the ignition exciters to make an electrical spark across the plug gap. The electrical spark ignites the fuel/air mixture in the combustor. The igniter plugs are attached to the diffuser case at the 3 o’clock position. Each igniter plug goes through a hole in the diffuser case and into the combustion chamber. The igniter plugs have shielded center electrodes. Classified spacers are installed under a mounting boss and are used to control the immersion depth of the igniter plug.


Page: 6


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Igniter Plug and Cables Page: 7



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

ENGINE STARTING SYSTEM General The engine starting system provides the means for cranking the engine on the ground and in flight, to an RPM at which an engine start can occur. Engine ground start can be accomplished using air supplied from the aircraft Auxiliary Power Unit (APU), the cross engine bleed air (air from the other engine), or a ground cart. In flight, windmill starts may need starter assistance, which could require the APU, or cross engine bleed air. The Electronic Engine Control (EEC) will minimize the time that the starter is free--running so that starter life is protected. If the EEC logic detects that the starter shaft has sheared by sensing that starter speed is high without any N2 rotation, it will command the Starter Air Valve (SAV) to close. The starter is attached to the rear of the engine main gearbox at the 5 o’clock position. The starter magnetic chip collector is located on the starter. The starter air control valve is attached to the starter air ducts at the 5 o’clock position. Functional Description The engine starting system consists of a pneumatic starter and an SAV assembly. The operation of the starting system is controlled by the EEC as a function of the cockpit engine start selector switch position (manual or automatic) and the engine start run on/off switch. The FADEC system can handle all aspects of engine start and/or crank , wet and dry). This includes control of the SAV, ignition exciter, and the introduction of fuel into the combustion chamber in response to aircraft command signals. During the start sequence, the EEC energizes the solenoid in the SAV. When energized, the solenoid puts aircraft starter duct air pressure (muscle pressure) to the actuation piston on the SAV and opens the valve. Air is then supplied to the starter where it drives the starter turbine to crank the engine. Once the valve opens, starter rotation is sensed by the speed pickup in the starter, and the signal is processed by the EEC.


Page: 8


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Engine Starter Page: 9



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

STARTING COMPONENTS Starter The starter converts air flow to power (torque) to drive the main gearbox, which in turn rotates the engine. The starter requires a corrected air flow rate of 59 pounds per minute. The starter speed sensor is a mechanically single and electrically dual channel sensor. The target teeth are located on the ring gear of the planetary gear system and reads the starter gearbox output speed into the clutch. This is the same speed as the starter output shaft during the start cycle. The indicated speed of the starter will go to zero when the engine speed is more than the starter clutch disengagement speed. The signal is processed by the EEC and used to find if the starter is rotating (during start) or not rotating (during normal engine operation). The starter has a single--piece output shaft, and Synchronous Engagement Clutch (SEC). The starter gears and bearing are splash--lubricated from a shared oil system (shared with the engine oil system). The starter oil capacity is 473 -- 591 ml. Fill and drain ports are provided in the housing for servicing. A magnetic chip collector plug assembly has an inner magnetic probe and an outer valve. The inner magnetic plug can be removed to check for metallic chips without draining the oil. A check valve in the outer valve prevents the loss of oil when the magnetic probe is removed. Starter Air Control Valve The valve is actuated with a FADEC controlled dual coil solenoid integrated into the SAV design. The position is determined by the high rotor speed (N2). The valve also has a 3/8 inch square drive for manual override capability. The SAV is a pneumatically actuated butterfly valve. It connects to a 4.0 inch stainless steel duct with a 0.035 inch wall thickness. V band flanges (lower starter air duct) and a flex joint coupling (located in the upper starter air duct) are used with the SAV to enhance system maintenance.


Page: 10


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Start Valve Page: 11



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

AUTOMATIC START Description The automatic start mode begins when all the following conditions are true: * the engine is not running and * the selected rotary selector is set to ING/START and * the selected engine MASTER lever is set to ON and * the ENG MAN START p/b is OFF. When the ENG MODE rotary selector is set to ING/START position, FADEC is powered, the ENGINE page is automatically shown on the SD and displays the IGN indication, SAV position and bleed pressure during this sequence. The APU bleed demand will increase and the pack valves will close. When the MASTER lever is set to ON position, the LPSOV opens and the automatic starting sequence begins. The EEC will automatically control the: -- Thrust Control Malfunction (TCM) cutback test, -- HPC active bleed valve (opening and closing), -- Hydraulic pump depressurizing (vía EIU) if necessary during in flight restart, -- SAV (opening and closing), -- Igniters ( one or two, on and off), -- Fuel Flow (FMV and HPSOV opening). The EEC energizes the SAV and the position of the SAV is confirmed open at the bottom of the ENGINE page thanks to the speed sensor feedback. The N2 begins to increase, when the engine reaches mínimum speed (18% N2), the EEC activates one igniter and controls the appropriate fuel flow to the burner. On the ENGINE page the spark igniter system (A or B) controlled by the EEC comes into view. On the E/WD, the FF increases, via FMV in the Integrated Fuel Pump and Control (IFPC). The EEC monitors the EGT and N2 accordmg to their schedules to provide the correct FF. When N2 reaches 51% N2, the automatic start sequence ends, EEC command the SAV to close and igniter to OFF, the engine continues accelerate and stabilizes at idle speed. The usual parameters are: -- N1 = 19%, -- N2 =58%, -- EGT = 440 C, -- FF = 227 kg/h. SCL /JGB / REV.00 / Jun--2016

If the second engine has to be started, the ENG MODE rotaty selector should stay on the IGN/START position. This will avoid activating the continuous ignition on the running engine if the selector is cycled to NORM and again to IGN/START. When both engines are running, the selector is set back to NORM, the WHEEL page will appear instead of the ENGINE page if at least one engine running. Automatic start abort TI1e EEC has the authority to abort a start only on the ground. The EEC will abort the start, dry motor the engine for 30 seconds and attempt a single start for the following reasons: -- no light up (EGT low and constant), -- no N2 acceleration (hung start), -- EGT reaches starting limit (impending hot start). Note: The maximum EGT during start sequence is 700  C. The EEC will aborta start, dry motor the engine for 30 seconds and not attempt a restart for the following conditions: -- Failure of automatic restart, -- N1 locked rotor, -- EEC unable to comand both igniters, -- Loss of EGT indication (T5 sensors failed), -- EEC unable to control fuel flow. The EEC will also abort start, will not dry motor the engine and will not attempt a restart if the starter duty cycle is excceded. Manual start abort: The automatic start sequence can be manually aborted by selection of the ENG MASTER lever to OFF position. Thus leads to: -- SAV closure, -- Igniter(s) off, -- FMV. LP and HP fuel shut--off val ves closure. Note: EEC does not dty motor the engine when an automatic start is manually aborted.

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AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Automatic Starting Page: 13



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

MANUAL START Description A manual engine start procedure is included in the EEC engine starting logic. In the manual start mode, engine starting control is under limited authority of the EEC. The SAV, fuel, and ignition are controlled from the cockpit via the EEC. Bleed air source being available, a manual start sequence is commanded by first setting the rotary selector to the IGN/START position to power and signal the EEC. The ENGINE page appears on the SD page of the ECAM and displays the IGN indication, SAV position and bleed pressure during this sequence. At the same time, the APU bleed demand will increase and the pack valves will close. The next action is to engage the ENG MAN START push--button to the ON position . This will lead the EEC to open the SAV. When N2 is above the mínimum speed (on--ground approximately 18% N2), the ENG MASTER lever is set to the ON position . The EEC commands fuel flow and both igniters simultaneously. The EEC monitors the EGT and N2 according to their schedules to provide the correct fuel flow but EGT limit protection is inactive, when N2 reaches 51% N2, the manual start sequence automatically ends when the EEC controls the SAV to close and the igniters to OFF. The engine continues to accelerate and stabilizes at idle speed. Manual start abort When a manual engine start has been initiated on ground or in flight, it shall be interrupted by either: -- de--selecting the ENG MAN START p/b before the ENG MASTER lever is commanded ON, or -- selecting ENG MASTER lever back to OFF position after it has already been selected ON. Interruption of a manual start shall result in the following EEC commands: -- SAV closure, -- igniters off, -- FMV and HP fuel shut--off valve closure. SCL /JGB / REV.00 / Jun--2016

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AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Manual Starting Page: 15



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

CONTINUOUS IGNITION General Continuous ignition is manually selected or automaticaliy controlled by FADEC. During continuous ignition both igniters are active.


Manual command: Once the engine is nmning and above idle, the pilot can manually command continuous ignition at any time by moving the rotary selector to the IGN/START position. Following a ground start, the rotary selector must be moved back to NORM before continuous ignition can be manually selected by moving it back to IGN/ START position. Continuous ignition shall remain commanded by the EEC until the rotary selector is moved back to NORM. ln the event that data position of the rotary selector sent by Engine Interface Unit (EIU) to EEC is not available or invalid, the EEC shall use the last valid value of the rotary selector position if the aircraft is on ground until a valid configuration is received again. Automatic command: The EEC automatically commands continuous ignition at the following conditions: -- lf an engine flameout is detected in flight, or during takeoff, igniters are kept on for a mínimum of 30 seconds after the engine has recovered from the flameout and reached idle, -- If a surge is detected in flight or during takeoff, igniters are powered until 30 seconds after the surge recovers, -- If the EEC detects a quick relight (Master Lever cycled from ON to OFF and back to ON in flight), -- If TCM Cutback is commanded. Automatic continuous ignition shall be inhibited if the burner pressure (PB) is above 150 psi (the nominal deteriorated igniter quench point) to preserve igniter life.


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AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Continuous Ignition Page: 17


POWER PLANT START AND IGNITION ENGINE CRANK General Cranking fuction is used to motor the engine on the ground for a short time with the use of the starter. There are two cranking modes: -- dry cranking, -- wet cranking. Dry Cranking The dry cranking procedure is used to motor the engine to remove unbumed fuel from the combustion chamber or cool down the engine or for some fuel or oil leak tests. The EEC shall enter the engine dry crank sequence when all of the following conditions are true: -- the engine is not mnning and, -- the aircraft is on ground and, -- the rotary selector is set to CRANK. This will power up the EEC and isolate both ignition systems. The ENGINE page appears automatically on the ECAM SD. When the ENG MAN START P/B is set to ON, the EEC commands the SAV to open. The dry motoring can be interrupted at any time by pushing the ENG MAN START p/b to OFF or positioning the ENG MODE rotary selector to NORM position. The usual starter duty cycle is 3 starter crank cycles or 4 minutes maximum of continuous cranking. A 30 minutes cool down period is necessary for additional use.


AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

WARNING:

Wet Cranking The wet cranking procedure is used to motor the engine for specific fue! or oil leak tests. The fuel flow is commanded but both ignition systems are isolated. The fue! goes through the IFPC to the actuator fuel pressure lines, the engine fuel manifolds (primary fuel lines only), and nozzles. Fuel is then sprayed in the combustion chamber. The first steps of the wet crank sequence are the same as the ones for the dry crank: -- the engine is not running, -- the aircraft is on ground, -- the rotary selector is set to CRANK (EEC powered, both ignition systems isolated, ENGINE page appears). -- the ENG MAN START P/B is set to ON. (SAV opening). When N2 speed stabilizes, the ENG MASTER lever is set to the ON position to command the fue! flow. After 15 seconds, the ENG MASTER lever is set to the OFF position to cut the fuel supply. The SAV command is maintained 30 seconds to blow all the fuel from the engine. The wet motoring ends by pushing the ENG MAN START pushbutton to OFF or/and positioning the ENG MODE rotary selector to NORM position.

THE EEC IS ABLE TO INITIATE A START SEQUENCE IMMEDIATELY FOLLOWING A DRY MOTORING SEQUENCE BY SETTING THE ENG MODE ROTARY SELECTOR TO IGN/START POSITION AND THE ENG MASTER CONTROL LEVER TO ON POSITION.


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AIRBUS A-- 320 PW 1100 NEO 74 / 80--00




Engine Cranking Page: 19



AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

EEC PROTECTIONS Auto Restart The EEC will abort the automatic start, dry motor the engine for 30 seconds and attempt a single auto --restart for the following reasons: * no light up (EGT low and constant) * No N2 acceleration (Hung Start) * EGT reaches starting limit (impending hot start or surge)

Starter Time Exceeded lf during a start or a crank sequence, the EEC identifies an excessive starter duty, it will generate the ECAM alert ”ENG x START FAULT (STARTER TIME EXCEEDED)” and abort the automatic sequence.


No Lightup If during an automatic start, the EEC identifies a low EGT: -- It shuts down the fuel supply and the selected igniter, -- It generates the ECAM alert ”ENG x IGN A(B) FAULT”, -- lt maintains the SAV open to clear fuel vapors and cool for 30 seconds, -- Then it controls simultaneously the fuel flow and both igniters, --When N2 reaches the starter cutout speed (or the light up is confirmed) -- it switches the igniters off and controls the SAV closure 1 seconds after (or l seconds after the starter duty cycle is exceeded). The engine continues to accelerate and stabilizes at idle speed. If this auto--restart attempt fails. the start is aborted and the EEC will generate the ECAM alerts ”ENG x START FAULT (IGNITION FAULT)” and ”ENG x. IGN A+B FAULT”. Impending Hot Start lf during an automatic start, the EEC identifies an impending hot start, it maintains the SAV open, the selected igniter and controls a fuel are cutoff for 2 seconds and back on for 12 seconds via the Fue! Metering Valve (FMV) fora maximum of 28 seconds to lower EGT below the limit. The EEC will generate the ECAM alert ”ENG x START FAULT (HOT START)” . If the fault disappears, the statting sequence goes on normally up to the engine stabilizes at idle speed. If the fault is still present, the EEC shuts down the fuel supply and the igniter, perfonns a dry motor for 30 seconds and attempts a single auto--restart. lf this auto--restart attempt fails. the start is aborted and the EEC will generate the ECAM alert ”ENG x START FAULT (EGT OVERLIMIT)”


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AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

NO LIGHT UP IMPENDING HOT START


AUTO RESTART


EEC Protections Page: 21

AIRBUS A-- 320 PW 1100 NEO 74 / 80--00

NOTES:





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AIRBUS A-- 320 PW 1100 NEO

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THRUST REVERSER



POWER PLANT THRUST REVERSER


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PW 1100 NEO

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EXHAUST SYSTEM General The Exhaust System is made up of nacelle components that form a flow path directing the air from the engine core and the engine fan. The shape of the nacelle is optimized to minimize drag and to maximize the thrust from the engine. The Exhaust System is made up of two subsystems: * Thrust Reverser, and * Turbine Exhaust. Thrust Reverser System The Thrust Reverser System protects the engine core, forms a path for fan bypass air and deploys to slow the aircraft upon landing.


AIRBUS A-- 320

Turbine Exhaust System The Turbine Exhaust System makes the path for the turbine gases exiting the engine core. It gives direction to the turbine gases, which helps to increase thrust and reduce turbulence. The Turbine Exhaust System is a cylindrical barrel and cone that makes a smooth exit for fan air and engine exhaust during engine operation. The system helps mix the fan bypass air with the turbine exhaust air. It reduces exhaust noise and helps increase thrust and performance. The system is installed on the aft flange of the engine. Major components include centerbody and nozzle assemblies. The two assemblies help to control the engine exhaust gas flow and send the gas aft. They also supply an aerodynamically smooth surface for the fan air from the thrust reverser. The system also incorporates drainage provisions to expel any hazardous fluids or vapors. Drainage outlets are found at the nozzle assembly, and at forward and aft centerbodies at 6:00. Turbine Exhaust Centerbody Assembly (Exhaust Plug) The centerbody assembly helps to control the engine exhaust gas flow and helps the nozzle to send gas aft. The assembly is attached to the rear of the Turbine Exhaust Case. The centerbody assembly provides an aerodynamic smooth surface for exhaust airflow and is designed to accelerate exhaust flow to high speeds. SCL /JGB / REV.00 / Jun--2016

The assembly has both forward and aft sections. The forward centerbody assembly is installed with 23 bolts on the Ti flange of the engines TEC. The aft centerbody assembly is attached to the forward centerbody assembly with 13 bolts and nut plates. The forward centerbody drain is at 6:00. Two alignment pins at 12:00 and 5:00 on the forward centerbody flange facilitate proper clocking of the centerbody drain provisions. Turbine Exhaust Nozzle Assembly The Turbine Exhaust Nozzle Assembly provides an efficient exit path for the engine exhaust gases leaving the LPT at a velocity and direction required to produce forward thrust. The exhaust nozzle assembly is a tapering cylindrical barrel installed on the To flange of the Turbine Exhaust Case. The nozzle assembly drain is at 6:00. The forward outer surface of the assembly aligns with the thrust reverser inner duct surface. The inner surface makes the outer contour of the engine exhaust. Fire seal fingers, or “turkey feathers” at the top of the exhaust nozzle prevent flames from exiting or entering the core compartment area in the unlikely event of a fire. Two cross flow blockers, one on each side of the Aft Pylon Fairing (APF), reduce the amount of hot core compartment air entering the fairing. The top of the exhaust nozzle has a pylon fire seal. Diverters to the left and right of the pylon seal help prevent cross flow of the exhaust under the pylon aft fairing. The nozzle is not in a flammable fluid zone. Eight outer fairing pads on the side provide protection to the exhaust nozzle during the translating sleeve operation. The exhaust nozzle assembly is made of Inconel alloys. The attachment flange has alignment pin holes at 12:00 and 6:00 to position the drain provisions. Note: the exhaust nozzle can be removed with the engine lift bracket installed. Scallops have been designed to clear the lift bracket nuts. A snubber bracket assembly limits deflections of the thrust reverser during flight. The assembly interfaces between the nozzle assembly and the thrust reverser, and includes a coated wear pad to prevent premature wear of the snubber bracket. A total of eight snubber bracket assemblies are located on the nozzle assembly, four to each side. The wear pads are attached to the snubber bracket assembly with four countersunk rivets. Page: 2


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Exhaust Page: 3




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THRUST REVERSER General The Thrust Reverser System provides the aerodynamic braking for the aircraft on the ground. Reverse thrust reduces the distance the aircraft needs to safely and efficiently stop during a landing or aborted take--off. During taxi and flight the reverser provides an efficient flow path that sends air aft for maximized thrust. Thrust reverser cowls are attached to the pylons on the left and right sides of the engine. Some major system components are shown below. * Translating sleeve * Inner Fixed Structure * Thrust Reverser Actuation System TRAS * Door Opening System DOS The thrust reverser is comprised of two halves that are mechanically independent. The halves hinge at the pylon, latching together along the bottom split line. An Inner Fixed Structure (IFS) provides thrust reverser support and protects the engine core cases and externals. The IFS forms the inner surface of the duct for fan bypass air. An outer structure that includes a translating sleeve and blocker doors forms the outer surface of the fan bypass air duct. The outer fan duct translating sleeve is normally stowed, providing uninterrupted fan air flow aft and producing the required thrust from the fan. Upon landing, the thrust reverser is deployed, moving the translating sleeve aft and allowing blocker doors to rotate to a vertical position and block the fan air. This action redirects the fan airflow through the thrust reverser cascades, sending it forward and outward in a controlled pattern that provides reverse thrust to help decelerate the aircraft. The Thrust Reverser Actuation System (TRAS) is composed of two hydraulic linear synchronized actuators per side. The actuators deploy and stow the reverser. Reverser operation is controlled by the EEC. For ease in opening and closing, thrust reverser cowls are equipped with a Door Opening System (DOS).


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Thrust Reverser Page: 5


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TRASLATING SLEEVE Description The translating sleeve is the movable thrust reverser component responsible for deploying the blocker doors and exposing the cascades. The Thrust Reverser System has two translating sleeves which run aft of the engine fan case from the hinge beam to the latch beam. The blocker doors redirect fan duct flow through the thrust reverser cascades. The doors also form part of the outer fan ducts aerodynamic surface, as well as the acoustic lining of the outer fan duct wall. A total of 10 blocker doors are used for a single thrust reverser, 5 per cowl half. The upper and lower doors are unique in size and cannot be installed into any other position. Two access panels are attached to the translating sleeve. They provide access to fasteners that attach the TRAS actuators to the translating sleeve assembly. Operation: Drag links aid in rotating and positioning the blocker doors into the fan duct, redirecting airflow through the thrust reverser cascades. 1. During activation of the thrust reverser, the translating sleeve slides aft. 2. As it slides, the blocker doors start to lift and rotate about their hinges due to the drag links attached to both the door and the Inner Fixed Structure. 3. Once the sleeve is fully deployed, doors will be in their full upright position.





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ACTUATOR ACCESS


TRASLATING SLEEVE RH

TRASLATING SLEEVE LH


Traslating Sleeve Page: 7




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INNER FIXED STRUCTURE (IFS) General The inner surface of the thrust reverser is formed by the IFS, whose primary purpose is to provide thrust reverser hoop continuity and react to surge and burst pressures through the bumpers, latches and hinges. The thrust reverser fixed structure covers the engine core, defines the core ventilation and fire zone, and forms the fan duct inner aerodynamic surface from the fan case exit to the core nozzle. Thermal blankets are attached to the whole inner side of the IFS and provide a barrier against the hot engine compartment. Reverse thrust loads from the drag link fittings are transferred to a one--piece inner V--blade attached to the IFS panel. Inner Fixed Structure components are shown below. * Hinge beam * Latch beam * Fire seals * Pressure relief door * Bumpers * Pre--cooler inlet duct * Active Clearance Control inlet duct Hinge Beams The hinge beams provide a structural connection between the IFS and transcowl through two tracks integrated into the beams.


Latch Beams Latch beams provide a structural connection between the IFS and transcowl. Latches Latches integral to the beams provide structural hoop load transfer to the nacelle, and offer quick access to the engine core. The latches also provide resistance to loads that might otherwise cause the thrust reverser to disengage or open during the flight cycle.


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Inner Fixed Structure Page: 9




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CLOSURE ASSIST ASSEMBLY General The Closure Assist Assembly consists of a turnbuckle that is used to draw the door together before engaging the latches. The assembly is located on the front of the left thrust reverser at 6:00. An eye on one end of the turnbuckle allows it to pivot so the pin engages the opposite cowling. Once engaged, the turnbuckle is turned manually to draw the two doors together. The turnbuckle is stowed after use. Bumpers Bumpers provide a hoop load path to resist the crushing pressure of the fan air stream upon the barrel sections and bifurcations. Three bumpers are located at the radiused corner intersections between the thrust reverser barrel section and the bifurcation area.


Bifurcation Latch System (BLS) The BLS is used to limit the deflection of the fixed structure, and to maintain the structural integrity in the event of a burst air duct. The BLS resembles the other bumpers, but also incorporates a locking feature to keep it engaged. The BLS handle is accessed through the latch access door. The system is a pull open, push--closed design with a baulking feature to visually indicate positive engagement of the latch. The latch is adjustable to ensure proper engagement. When closed the handle remains parallel to the engine centerline. Fire Seals Fire seals provide fire protection and aerodynamic fit. The seals are located circumferentially on the Inner Fixed Structure. Fire protection consists of fireproof barriers and seal arrangements. The fire seals prevent the escape of flames from a fire zone and prevent air or fluids from entering. The seals are made of a soft, compressible fireproof material. The “turkey feather” barrier seal is made of metal. Pressure Relief Door The pressure relief door provides adequate pressure relief when core cavity pressure exceeds the design limit. The pressure relief door is on the left IFS at 9:00. The door must open to provide adequate pressure relief when core cavity pressure goes above the designed cavity pressure. The pressure relief door opens if cavity pressure exceeds the latch release pressure, such as in the event of a burst cooling air duct. SCL /JGB / REV.00 / Jun--2016

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Closure Assist Page: 11




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TORQUE BOX ASSEMBLY Torque Box Assembly The torque box is a curved beam with torsional stiffness that supports the thrust reverser actuators, cascades, hinge beam and latch beam. The torque box runs from the hinge beam to the latch beam. Torque from the actuators is transferred to the beams through the gussets. Loads from the Door Opening System (DOS) and Hold Open Rods (HORs) are transferred to the torque box through a DOS fitting and brackets attached to the bulkhead.

Hold Open Rods (HORs) HORs prop open the thrust reverser to provide a safe environment for performing maintenance work on the engine. HORs are attached to the thrust reverser torque box and the engines fan case. Each thrust reverser has one Hold Open Rod. Hold Open Rods are stowed on the torque box when not in use.


Cascade Array The cascade array turns fan airflow forward and sideways. Each cascade has a forward attachment at the torque box and an aft attachment at the aft cascade ring. The fastener attachment pattern allows for limited interchangeability between cascade locations while ensuring the cascade cannot be installed backwards. Each engine has seven cascade boxes. Reverse thrust is achieved when the transcowl deploys the blocker doors and exposes an array of carbon--fiber composite cascade boxes. The forward turning component of each cascade contributes to reverse thrust. The combination of forward--turning and side--turning angles produces an efflux pattern that prevents these effects: * reverse thrust airflow re--ingestion into the inlet * cross engine inlet re--ingestion * fuselage impingement * impingement on the aircrafts control surfaces and high--lift devices. Aft Cascade Ring (ACR) The Aft Cascade Ring provides aft support for the cascade segments. The ACR beam runs from the hinge beam to the latch beam along the aft edge of the cascades. The ACR is a curved, open--section metal beam that transmits aerodynamic and dynamic loads from the cascades to the hinge and latch beams.


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CASCADE


TORQUE BOX


Cascade Components Page: 13




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THRUST REVERSER ACTUATION SYSTEM (TRAS) General A Hydraulic Control Unit (HCU) provides isolation and directional control of hydraulic fluid used for thrust reverser actuation. The thrust reverser is comprised of two halves that are mechanically decoupled, and which hinge from the pylon. The halves latch together along the bottom split line. Upon deployment, translating sleeves move aft, causing blocker doors to rotate and block the fan air. This action redirects the fan flow through the thrust reverser cascades, sending air forward and outward in a controlled plume and providing reverse thrust. Two hydraulic linear synchronized actuators per side deploy and stow the reverser. The thrust reverser can be used for either engine position as long as the cascade pattern is changed to match airflow control requirements. Each half of the thrust reverser has two actuator units that include an integral locking mechanism and proximity switch, and one actuator with a LVDT for reverser position feedback. The thrust reverser also includes a “third line of defense” provided by two fully independent Tertiary Locks (TL). Reversers can be deployed and stowed manually on the ground by using the Manual Drive Units (MDU) located on the lower actuators. Components for this system are: * Hydraulic Control Unit HCU * Filter module assembly * T--piece manifold * Locking actuator (2) * Locking feedback actuator (2) * Manual Drive Unit (2) MDU * Synchronization flexshaft (2) * Track Lock Valve (2) TLV * Tertiary Lock (2) TL * Wiring harness and plumbing up to pylon interfaces.


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Thrust Reverser Logic Page: 15




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HYDRAULIC CONTROL UNIT (HCU) General The HCU controls the locking, unlocking and translating operation of actuators. The HCU is located in the pylon upper spar. The HCU contains the Isolation Control Valve (ICV) and the Directional Control Valve (DCV), and their respective solenoids. The ICV and DCV are operated from dual channel solenoid pilot valves upon a signal from the EEC. A manual inhibition function provides these capabilities: * dispatches inoperative thrust reversers * prevents inadvertent activation of the ICV * provides additional safety during ground maintenance activities. Deactivation is provided by an inhibit lever and inhibit pin. Proximity sensors detect the levers position.

T--piece Manifold The T--piece manifold provides a path for hydraulic system pressure to be ported from the Hydraulic Control Unit to the TRAS actuators. The manifold is located in the forward section of the pylon. The T--piece manifold has a total of six lines that distribute hydraulic fluid to the thrust reverser actuators. Three of the lines are located on the left and three are on the right.


Operation: The isolation valve isolates the entire downstream system from the aircraft hydraulic pressure supply, unless an arming signal is received. Hydraulic fluid to power the TRAS is supplied from the aircraft system. The directional control valve uses a spool valve to regulate the direction of pressure application to the actuators in order to achieve the unlock, deploy, stow and relock sequence. DCV position signal comes from proximity sensors. Filter Module Assembly The filter module assembly protects downstream TRAS components from debris that may be present in the aircraft hydraulic system. The assembly is mounted in the pylon above the lower aft pylon fairing. The 15 micron filtration system is installed on the inlet side of the thrust reverser supply. A differential pressure indicator, or “red button” pops out to visually indicate when the filter is clogged. An automatic shutoff valve allows removal and installation of the filter without significant loss of hydraulic fluid from the system. In this instance the system will still function, but will be slow to deploy or stow.


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Hydraulic Control Unit Page: 17




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LOCKING ACTUATOR General The actuators deploy and stow the reverser sleeves and also lock them in the stowed position. The actuators are mounted on the thrust reverser torque box at approximately 5:00 and 7:00. The lower actuators each incorporate an integral, hydraulically released mechanical locking element. A manual lock release provision facilitates manual translation of the reverser via manual drive units located on each side. Lock release and manual drive access is gained by opening the fan cowl doors. Both the lower and upper actuator in each half of the thrust reverser incorporate a hydraulically released mechanical element. A proximity sensor switch is included in the locking mechanism of each actuator to provide the electrical feedback of lock engagement status to the EEC. An inhibit handle is provided to lock out the actuator during maintenance. A gimbal allows for thrust and torsional motion of the actuator during operation.


Operation: The actuators are operated hydraulically using worm wheels and flex drive shafts. When the command is given by the crew through the movement of the throttle levers, and when additional conditions are met, the reverser deploys and stows.


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Locking Actuator Page: 19




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LOCKING FEEDBACK ACTUATOR General The actuators are used to deploy and stow the reverser sleeves, lock them in the stowed position, and provide feedback of sleeve position.The actuators are mounted on the thrust reverser torque box at approximately 1:00 and 11:00. The actuators each incorporate and integral, hydraulically released mechanical locking element. Electrical feedback of thrust reverser position is accomplished by means of an LVDT integrated into each upper actuator. An inhibition handle is provided to lock out the actuator during maintenance. A gimbal allows for thrust and torsional motion of the actuator during operation.


Operation: The actuators are operated hydraulically using worm wheels and flex drive shafts. When the command is given by the crew through the movement of the throttle levers, and when additional conditions are met, the reverser deploys and stows. A single channel LVDT device on each upper actuator can provide true dual channel thrust reverser position feedback to the EEC via the mechanical coupling between the upper and lower actuators.


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Locking Feedback Actuator Page: 21




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MANUAL DRIVE UNIT General The manual drive unit allows manual translation of the thrust reverser during maintenance. The MDU is mounted onto both locking actuators. The MDU is only used when aircraft is in maintenance configuration. Maximum operational speed of the MDU is 600 rpm.


Operation: A 3/8 square drive tool input enables manual rotation of the TRAS synchronization system and translates the sleeve as required. A clutch torque limiter in the MDU that provides over torque protection is installed to prevent damage to TRAS components during TRAS manual operation. A baulking feature on the fan cowl pushes the lock lever to the locked position in the event the MDU lock lever is left unlocked.


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Manual Drive Unit Page: 23




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FLEX SHAFT AND LOCKS Synchronization Flex Shafts Flex shafts provide TRAS synchronization by connecting the upper locking feedback actuator to the lower locking actuator on each thrust reverser half. The synchronization shafts are located inside the deploy tubes which are mounted to the thrust reverser torque box. Each cowl half has one flex shaft. The deploy tube provides a path for the TRAS hydraulic system to pressurize during deploy and stow sequence. The flex shaft square drive interfaces with the actuator drive system and is secured by the deploy tube to the actuator. Track Lock Valve (TLV) The Track Lock Valve controls the locking and unlocking of the tertiary locks. The TLV is mounted to a bracket that is attached to the torque box of each thrust reverser door at approximately 6:00. The TLV is independently controlled by a 115 VAC aircraft signal, which is external to the TRAS and provides fully independent control of the tertiary locks. The TLV is considered part of the “third line of defense” to prevent inadvertent deployment of the thrust reverser during flight.

Operation: 1. Once the Track Lock Valve is commanded open by the aircraft, the TLV solenoid energizes, and the blade on the Tertiary Lock translates to the unlocked position. 2. Proximity sensor targets move to the “far” position, allowing them to send redundant signals to the aircraft that indicate the tertiary lock is unlocked. 3. The thrust reverser deploys. 4. When the Tertiary Locks receive supply pressure ported from the energized TLV, the piston retracts the blade, allowing the Tertiary Lock to unlock. 5. Once the transcowl has translated outside of the stowed locking region, the solenoid stays energized, allowing the Tertiary Lock to remain unlocked until the TRAS is returned to the stowed position.


Operation: When energized, the TLV solenoid allows fluid to be ported to the tertiary locks, unlocking them for deployment of the thrust reverser. Tertiary Lock If the mechanical locking system fails to retain translating sleeves, the tertiary lock prevents the transcowl from deploying past the stowed position. The Tertiary Lock is mounted to the aft section of the latch beam. Each Tertiary Lock incorporates two proximity switches which provide lock/unlock status to the EEC. The Tertiary Lock is biased to the locked position by four springs between the blade and the piston.


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Flex Shaft Page: 25




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DOOR OPENING SYSTEM General The Door Opening System (also known as thrust reverser opening mechanism) opens the thrust reverser cowl assembly and provides maintenance access to the engine core. The DOS is located in the fan compartment. Actuators are attached to the engine fan case on one end, and the thrust reverser torque box on the other end. The system consists of an electro--hydraulic power pack, a reservoir, two reverser cowl C--duct actuators, and two operating switches. Power for the system comes from the aircraft. A manually operated hydraulic pump is coupled to the actuator. When the hydraulic fluid is pressurized, the pressure within the actuator holds the thrust reverser in the open position so a Hold Open Rod can be connected.


Operation: The C--duct configuration provides maintenance and service access to the engine when the halves are opened by the DOS. This is actuated via command switches located on the fan compartment. Hold Open Rods are provided to allow one opening angle of approximately 45. Operating switches mounted on the left and right hand sides of the fan case allow for easy opening of the C--ducts. Pressing and holding the switch activates the hydro--electric power pack. The pack pressurizes the oil stored in the reservoir and supplied to the left or right actuator, depending on the switch being used. The actuator will lift the door and lock once the door is in its proper opening position. Once in position, Hold Open Rods can be installed to keep the doors safely open for maintenance. In the event power is not available, a quick--disconnect fitting on the actuators allows the doors to be opened manually with a single action hand pump.


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Door Opening System Page: 27




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THRUST REVERSER OPERATION General Each system is pressurized by dedicated hydraulic power source: * Green Hydraulic System for engine 1 * Yellow Hydraulic System for engine 2 Each system is made of one HCU including an Isolation Control Valve (ICV) and a Directional Control Valve (DCV), two worm drive actuators per side, locking and monitoring devices. To avoid inadvertent deployment, the system operates under multiple and independent commands and its comprises several lines of defense, primary locks in each actuator and one tertiary lock at the bottom of each traslating sleeve. Deploy Sequence The EEC confirms the engine is running. The thrust reverser are stowed, locked and not inhibited. In this conditions: * the ICV, DCV, Track Lock Valves are deenergized to prevent pressurization. * the 6 proximity sensors indicate locked. * the ICV pressure switch indicates a low pressure. * both LVDTs indicates a stowed condition. * the HCU inhibition lever proximity sensor indicates a non inhibited condition. When the trust reverser lever is set to deploy position, the following sequence occurs: 1-- As soon as the Spoiler Elevator Computers (SECs) receive the signal from the TCU potentiometers (TLA <--3), and from the RA altitud <6 Ft., they control the powering of the TLVs to open. In this position, the TLVs are ready to let the hydraulic pressure release the Track Lock when the ICV will be controlled to open. 2-- When the EIU receives the signal from the TCU switch (< --3.8) and from the LGCUs (aircraft on ground), it controls the closure of internal relays involved in the ICV and DCV powering. 3-- When the EEC receives the signals from the TCU resolvers (TLA < 4.3 ), it closes an internal rtelay to power the ICV to open. The pressure is sent to the actuators rod chambers to perform and overstow and to TLs to release the latches.


4-- When the EEC receives the signal from the TCU resolvers (TLA < 4.3), provided the TLs are confirmed unlocked, it closes and internal relay to power the DCV to open. The pressure is sento to the actuators jack heads to released the actuators internal primary locks and command the traslating sleeves deployment. 5-- Above 85% of travel, the EEC comands the engine to accelerate from reverser idle to max reverser thrust. Maximum allowable thrust is defined as afunctionof sleeve travel and TLA. At 95% of travel, the actuatorsengage the irintegral snubbing devices, thus decreasing their extension speed before the full opening. The TLV, ICV, and DCV remain supplied to maintain the traslating sleeve fully deployed by hydraulic pressure.

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Deployed Page: 29




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Stow Sequence When the thrust reverser lever is set to stow position, the following sequence occurs: 1-- When the EEC receives the signal from the TCU resolvers (TLA > --4.8 ), it deenergizes the DCV. The pressure is sent only to the actuators rod chambers to stow the traslating sleeves until the actuators internal primary locks are reengaged. 2-- 15 seconds after the SECs receive the signals from the TCU potentiometers (TLA>--2 ), they deenergize the TLVs to reengage the TLs. 3-- 15 seconds after the stow sequence is completed, the EEC deenergizes the ICV. Then the EIU opens its internal relays to isolate the ICV and DCV powering. Ground Assisted Stow Sequence (GASS) The EEC shall initiate a thrust reverser GASS operation on ground only in order to lock the thrust reverser system in the following two cases: * at least one primary lock is detected unlock after the normal stow sequence is completed (operational case). * if at least one primary lock is detected unlock after the engine start (maintenance case). The GASS shall be initiated by energizing the ICv for 5 seconds when all the following condition are fulfilled: -- the aircraft is on ground -- the throttle is in forward thrust region and less than CL position. -- no stow sequence is being commanded. -- within 15 seconds after engine transition to idle following an engine start. -- one or two primary locks of any translating sleeve are seen unlocked. -- the sleeve position (left and right) are less than 5% of travel. -- the thrust reverser is not inhibited. -- 28 VDC power is available.


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Stowed Page: 31


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NOTES:





Page: 32


ABBREVIATIONS



POWER PLANT ABBREVIATIONS

SCL / JGB / REV.00 / Jun -- 2016

Page: 1


POWER PLANT ABBREVIATIONS ABBREVIATIONS D

A ACC ADIRU ADSOV AGB AOHE AOP APF APU ATS

Active Clearance Control Air Data Inertial Reference Unit Active Damper Shut Off Valve Angle Gearbox Air/Oil Heat Exchanger Auxiliary Oil Pressure sensor Aft Pylon Fairing Auxiliary Power Unit Air/Turbine Starter

B BAHE BAM BAPS BAV BFE BLS BOAS

Buffer Air Heat Exchanger Buffer Air Manifold Buffer Air Pressure Sensor Buffer Air Valve Buyer Furnished Equipment Bifurcation Latch System Blade Outer Air Seals

C



CAA CAN CAS CIC CJC CSD

Closure Assist Assembly Controller Area Network Crew Alert System Compressor Intermediate Case Cold Junction Compensation Constant Speed Drive

SCL / JGB / REV.00 / Jun -- 2016

DCV DOS DPHM DSU

Directional Control Valve Door Opening System Diagnostic, Prognostic and Health Management system Data Storage Unit

E EBU ECAI ECAM ECS EDP EEC EGT EHSV EIU

Engine Buildup Unit Engine Cowl Anti--Ice Electronic Centralized Aircraft Monitor Environmental Control System Engine Driven Pump (hydraulic) Electronic Engine Control Exhaust Gas Temperature Electro Hydraulic Servo Valve Engine Indicating Unit

F FADEC FDGS FDV FEGV FFDP FFM FIC FM FMU FOHE FOHEBV FPRV FRTT

Full Authority Digital Engine Control Fan Drive Gear System Flow Divider Valve Fan Exit Guide Vane Fuel Filter Differential Pressure Sensor Fuel Flow Meter Fan Intermediate Case Fuel Manifold Fuel Metering Unit Fuel/Oil Heat Exchanger Fuel/Oil Heat Exchanger Bypass Valve Fuel Pressure Relief Valve Fuel Return--to--Tank

Page: 2


POWER PLANT ABBREVIATIONS G GSE

Ground Support Equipment

H HCU HOR HP HPC HPSOV HPT

Hydraulic Control Unit Hold Open Rod Hydraulic Pump High Pressure Compressor High Pressure Shutoff Valve High Pressure Turbine

LPT LRU LSOP LVDT

Low Pressure Turbine Line Replaceable Unit Lubrication and Scavenge Oil Pump Linear Variable Differential Transformer

M MDU MGB MOP MOT

Manual Drive Unit Main Gearbox Mail Oil Pressure sensor Main Oil Temperature sensor

N

I IBR ICV IDG IDGFOHE IDGOOHE IFPC IFS IPCKV IVB

Integrally Bladed Rotor Isolation Control Valve Integrated Drive Generator IDG Fuel/Oil Heat Exchanger IDG Oil/Oil Heat Exchanger Integrated Fuel Pump and Control Inner Fixed Structure Inlet Pressure Check Valve Inner V Blade

J JOSV

Journal Oil Shuttle Valve

K KVA For Training Purposes Only


Kilo Volt Ampere

L LAP LH LOP LPBSOV LPC

Latch Access Panel Left Hand Low Oil Pressure switch Low Pressure Buffer Shutoff Valve Low Pressure Compressor

SCL / JGB / REV.00 / Jun -- 2016

N1 N2 Nf NAI NAI PRSOV NEO

Low Pressure Turbine rotor speed High Pressure Turbine rotor speed Fan rotor speed Nacelle Anti--Ice Nacelle Anti--Ice Pressure Regulating Shutoff Valve New Engine Option

O OCM ODM OFDP OLS OTAD OVB

Oil Control Module Oil Debris Monitor Oil Filter Differential Pressure Sensor Oil Level Sensor Oil Tank Access Door Outer V Blade

P P2.9 P2T2 P2.5/T2.5 Pamb Pb PCE

Pressure at Station 2.9 Pressure and temperature at Station 2 Pressure and temperature at Station 2.5 Ambient pressure sensor Burner pressure Pre--Cooler Exhaust Page: 3


POWER PLANT ABBREVIATIONS PF PHMU PMA PRSOV PS14

Pressure fuel Prognostics and Health Management Unit Permanent Magnet Alternator Pressure Regulating Shutoff Valve Pressure at Station 14 Quick Attach/Detach Quick Release

R RH rpm RTD RTTV

Right Hand revolutions per minute Resistance Temperature Device Return--To--Tank valve

S SAV SVA

To TRAS

Outer T--flange Thrust Reverser Actuation System

U UTAS

United Technologies Aerospace Systems

V

Q QAD QR


VDC VIB VIGV VORV VSV

Voltage Direct Current Vibration Variable Inlet Guide Vane Variable Oil Reduction Valve Variable Stator Vane

W WC WF

Wiring for core harness Wiring for fan case harness

Starter Air Valve Stator Vane Actuator


T T2 T3 T5 Tamb TAI TCA TCM TEC TF Ti TIC TL TLV

Fan inlet temperature (at Station 2) Compressor exit Temperature (at Station 3) Turbine exit temperature (at Station 5) Ambient temperature Thermal Anti--Ice Turbine Cooling Air Thrust Control Malfunction Turbine Exhaust Case Fuel temperature Inner T--flange Turbine Intermediate Case Tertiary Lock Track Lock Valve

SCL / JGB / REV.00 / Jun -- 2016

Page: 4


FIRE PROTECTION

A320 NEO ATA 26


ATA 26 FIRE PROTECTION

SCL / REV.00 / Jun. 2016

Page: 1


FIRE PROTECTION

A320 NEO ATA 26

FIRE PROTECTION SYSTEM COMPONENT LOCATION SYSTEM OVERVIEW The engine and APU fire protection is done by two sub--systems: the FIRE detection system and the FIRE extinguishing system.


ENGINE AND APU FIRE PROTECTION The engines and the APU have individual fire detection systems. Each system has two identical detection loops (A and B) installed in parallel. Each loop includes 3 detector elements. These detection elements are located around the Accessory Gear Box (AGB), Core engine area and pylon area. The two loops are monitored by a Fire Detection Unit (FDU). FDU 1 monitors the loops on engine 1 and FDU 2 monitors the loops on engine 2. The FDU sends FIRE and FAULT signals to the Flight Warning Computer (FWC) for display on ECAM. The APU has two identical loops (A and B) installed in parallel on the APU compartment. These loops are monitored by FDU APU. The guarded FIRE P/B switches give FIRE indication and are used to isolate the related systems. When the FIRE pushbutton is released out, the engine or APU will shut down. This also arms the extinguishing system. Each engine has 2 fire bottles installed in the pylon. The discharge of each bottle is controlled by a related AGENT P/B SW on the FIRE panel. For the APU, there is only one fire extinguisher bottle, which is installed in the aft fuselage forward of the APU firewall. Its discharge is controlled by one AGENT P/B SW. On the ground, an APU FIRE will cause an automatic shutdown of the APU and an automatic discharge of the bottle. The TEST buttons are used to do tests on the different fire detection and extinguishing systems and make sure they operate correctly.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26

Figure 1 SCL / REV.00 / Jun. 2016

ENGINE AND APU FIRE PROTECTION Page: 3


FIRE PROTECTION

A320 NEO ATA 26

ENGINE AND APU FIRE PROTECTION (NEO SPECIFIC)


For Pratt and Whitney (PW) 1100G engine, the accessory gear box is located in the Core engine area. The PW has 3 fire detectors (pylon, AGB and core). The CFM Leap has 4 detectors (pylon, fan, AGB and core). The fire detection and extinguishing principle is identical on all Single Aisle family.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


ENGINE AND APU FIRE PROTECTION (NEO SPECIFIC) Page: 5


FIRE PROTECTION

A320 NEO ATA 26

AVIONICS SMOKE DETECTION


The A320 family aircraft have a cooling system for the avionics equipment. The cooling system is controlled and monitored by the Avionics Equipment Ventilation Controller (AEVC). The circulation of the air through the system is supplied by a blower fan (cool air supply) and an extraction fan (warm air removal). The extraction airflow is downstream of the avionics equipment. The avionics SMOKE detector, which is installed in the extraction duct, is used for the detection of smoke from the computers and control boxes. The detector is monitored by the AEVC. The smoke detector directly sends the signal to FWC for the AVIONICS SMOKE warning in the cockpit.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


AVIONICS SMOKE DETECTION Page: 7


FIRE PROTECTION

A320 NEO ATA 26

CARGO SMOKE DETECTION The cargo compartment smoke detection system is monitored by the Smoke Detection Function (SDF) integrated in the Cabin lntercommunication Data System (CIDS). The CIDS--SDF receives signals from the cargo detectors and sends SMOKE or FAULT warnings to the Flight Warning Computer (FWC) to give an alert to the flight crew.

CARGO FIRE EXTINGUISHING


The cargo compartment fire--extinguishing agent is discharged into the FWD compartment through one nozzle (2 nozzles for A321) or into the AFT compartment through two nozzles (3 nozzles for A321). The standard system includes one extinguishing bottle in the FWD cargo compartment (or in the AFT bulk cargo compartment, RH side, forA318 and A319). For the A320, the extinguishing bottle can also be in the AFT cargo compartment. An optional system includes two bottles. The second bottle can be used for large range operations.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


CARGO SMOKE DETECTION AND CARGO FIRE EXTINGUISHING Page: 9


FIRE PROTECTION

A320 NEO ATA 26

LAVATORY SMOKE DETECTION AND EXTINGUISHING


The lavatory smoke detection system is monitored by the SDF integrated in the CIDS. The CIDS--SDF receives signals from the lavatory detectors and sends SMOKE or FAULT warnings to the FWC to give an alert to the flight crew. The protection of each lavatory waste-- bin is done by an automatic fire extinguishing system. A small pressurized extinguisher will automatically discharge if there is a fire. The fusible material in the discharge tube melts at high temperature and the pressurized agent is discharged into the waste bin. LAV SMOKE warnings are also sent to the CIDS to give an alert to the cabin crew.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


LAVATORY SMOKE DETECTION AND EXTINGUISHING Page: 11


FIRE PROTECTION

A320 NEO ATA 26

ENGINE FIRE DETECTION


Each fire detection loop contains 3 detector elements connected in parallel, The PW fire detectors are located: S one around the AGB, S one in the Core compartment (270 to 330 degrees) between the fuel nozzles and the aft circumferential ventilation outlet, S one protecting the pylon above the combustion chamber.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


ENGINE FIRE DETECTION Page: 13


FIRE PROTECTION

A320 NEO ATA 26

ENGINE FIRE EXTINGUISHING


The engine fire extinguishing bottles are in the pylon. There are access panels on the two sides of the pylon.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


ENGINE FIRE EXTINGUISHING Page: 15


FIRE PROTECTION

A320 NEO ATA 26

APU FIRE DETECTION AND EXTINGUISHING


Each APU fire detection loop is a single detector element installed around the interior of the APU compartment. The APU fire extinguishing bottle is in the aft fuselage forward of the APU firewall. There is an access panel on the lower fuselage.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


APU FIRE DETECTION AND EXTINGUISHING Page: 17


FIRE PROTECTION

A320 NEO ATA 26

AVIONICS SMOKE DETECTION


There is only one avionics smoke detector, which is in the avionics compartment in the ventilation extraction duct.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


AVIONICS SMOKE DETECTION Page: 19


FIRE PROTECTION

A320 NEO ATA 26

CARGO COMPARTMENT FIRE DETECTION AND EXTINGUISHING


Each cargo compartment has 2 smoke detectors in each cavity. The smoke detectors are installed in recessed panels in the compartment ceiling.

SCL / REV.00 / Jun. 2016

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FIRE PROTECTION

A320 NEO


ATA 26


CARGO COMPARTMENT FIRE DETECTION AND EXTINGUISHING Page: 21


FIRE PROTECTION

A320 NEO ATA 26

LAVATORY SMOKE DETECTION AND EXTINGUISHING


Each lavatory has only one smoke detector, installed in the air extraction duct in the lavatory ceiling. A fire extinguisher is located above the waste bin in each lavatory service cabinet.

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FIRE PROTECTION

A320 NEO


ATA 26


LAVATORY SMOKE DETECTION AND EXTINGUISHING Page: 23


FIRE PROTECTION

A320 NEO ATA 26


This Page Intentionally Left Blank

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO ATA 28


ATA 28 FUEL

SCL / REV.00 / Jun. 2016

Page: 25


FUEL

A320 NEO ATA 28

FUEL SYSTEM SYSTEM INTRODUCTION


A319/A320 The new A319/A320 Fuel System is a combination of the common Wing Structure of the SA Aircraft manufactured (A318/A319/320) and the Fuel System Components installed in A321 since entry into service and will be the manufacturing standard with the beginning of delivery of all NEO Single Aisle Aircraft, except the A318. The benefit to combine both layouts, on the new A319/A320 Fuel System, is to achieve the following; S Weight reduction, S Better protection against UERF, S Cost improvements, S Commonality in between all SA Aircraft variants. Two fuel pumps are installed in each wing tank. One fuel pump is installed for the APU. Fuel is supplied to the engines from the wing tanks only. As the fuel level in the wing decreases, the center tank fuel is transferred to the wing tanks until the center tank is empty. Fuel transfer from the center tank to the wing tanks is controlled by transfer valves. When the transfer valves are opened, they supply pressure to two jet pumps in the center tank and transfer the fuel from the center tank to the wings. Two engine LP valves are installed to supply or cut off fuel to the engines. The LP valve is closed when the related engine is shutdown or when the engine fire pushbutton is released. A crossfeed valve is installed to connect or isolate the left and right hand sides. It enables engine to be fed from any available fuel pump. On the ground, the crossfeed valve enables fuel to be transferred from tank to tank. The valve is closed for normal operation. The fuel system also feeds the APU directly from the left hand side. The APU LP valve is installed to supply or cut off fuel to the APU. It closes when the APU is shut down or when the APU FIRE pushbutton is released out. SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


SYSTEM INTRODUCTION A319/A320 Page: 27


FUEL

A320 NEO ATA 28

SYSTEM INTRODUCTION (CONTINUED)


A321 The A321 fuel tanks are integrated into the center fuselage area and the wings. Like the A318/A319/A320, the center tank is part of the center wing box but unlike the A318/A319/A320, the wing tanks are not divided. The tanks are simply called left and right wing tanks.

SCL / REV.00 / Jun. 2016

Page: 28


FUEL

A320 NEO


ATA 28


SYSTEM INTRODUCTION A321 Page: 29


FUEL

A320 NEO ATA 28



A318 The fuel tanks are integrated into the center fuselage area and the wings. The A318 center tank is part of the center wing box. The wing tanks are divided into inner and outer cells. To reduce the structural load on the wings, the fuel in the outer cells is not used until the fuel load in the inner cells decreases to a low level. Two fuel pumps are installed in the center tank, two fuel pumps are installed in each wing tank inner cell and one fuel pump is installed for the APU. Fuel is supplied to the engines from the center tank first. After the center tank is empty, fuel is supplied from the wing inner cells. There is no direct supply from the outer cells to the engines. Two intercell transfer valves in each wing let the fuel transfer from the outer cells to the inner cells when the low level is reached. Two engine Low Pressure (LP) valves are installed to supply or cut off fuel to the engines. The LP valve is closed when the related engine is shut down or when the engine fire pushbutton is released. A crossfeed valve is installed to connect or isolate the left and right hand sides. It enables engine to be fed from any available fuel pump. On the ground, the crossfeed valve enables fuel to be transferred from tank to tank. The valve is closed for normal operation. The fuel system also feeds the APU directly from the left hand side. The APU LP valve is installed to supply or cut off fuel to the APU. It closes when the APU is shut down or when the APU FIRE pushbutton is released out.

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FUEL

A320 NEO


ATA 28


SYSTEM INTRODUCTION A318 Page: 31


FUEL

A320 NEO ATA 28

CONTROL AND INDICATING This section will highlight the control panels and indications for the fuel system. CONTROL PANELS The FUEL control panel is operated from the overhead panel. The A319/A320/A321 FUEL control panel is very similar to the A318, except: S The fuel transfer link between the center and wing tanks is indicated, S CTR TK XFR is indicated instead of CTR TK PUMPS.


The wing tank pumps are controlled manually but the center tank pumps are normally controlled automatically. On the A318 fuel control panel, the MODE SEL P/B SW enables the pilot to select automatic or manual mode for the center tank pumps. The MODE SEL P/B SW on the A319/A320/A321 FUEL control panel enables the pilot to select manual or automatic mode for the CTR TK XFR valves.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


CONTROL AND INDICATING - CONTROL PANELS Page: 33


FUEL

A320 NEO ATA 28

CONTROL AND INDICATING (CONTINUED) ECAM FUEL PAGE The configuration of the fuel system valves and pumps as well as i quantity indications are displayed on the ECAM FUEL system page. The total Fuel On Board (FOB) indication is duplicated on the Engine/Warning Display. Let’s briefly review all the A321 differences using the ECAM FUEL page: S There is no inner and outer cells in the wing tanks, S Fuel is transferred from the center tank to the wing tanks via two jet pumps and transfer valves, S Fuel is always fed to the engines from the wing and not from the center tank.


For the actual A319/A320 configuration there is a. combination of the A318 and the A321 Fuel System: S Still a separation into inner and outer cell for the wing tank S Fuel is transferred from the center tank to the w ing tanks via two jet pumps and transfer valves, S Fuel is always fed to me engines from the wing and not from the center tank.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


CONTROL AND INDICATING ECAM FUEL PAGE Page: 35


FUEL

A320 NEO ATA 28

REFUEL/DEFUEL PANEL


The Refuel/Defuel panel functions are: S Automatic or manual refueling S High level test, S Defueling, S Fuel transfer, S Refueling on batteries.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


REFUEL/DEFUEL PANEL Page: 37


FUEL

A320 NEO ATA 28

REFUEL/DEFUEL COUPLING AND REFUEL VALVE


The Refuel/Defuel coupling is below the RH wing leading edge. The Refuel/Defuel coupling shown is the optional one on the LH side. There is one refuel valve per tank. Each of the three refuel valves has a manual plunger. When pressed, the plunger holds the valve open in case of a valve electrical failure during refueling.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


REFUEL/DEFUEL COUPLING AND REFUEL VALVE Page: 39


FUEL

A320 NEO ATA 28

MAINTENANCE/TEST FACILITIES


The 2--channel Fuel Quantity Indication Computer (FQIC) calculates the fuel mass, controls automatic refueling and monitors the system with different interfaces. The Fuel Level Sensing Control Units (FLSCUs) send fuel level signals to the FQIC and to different aircraft circuits and systems. They are installed in the avionics compartment. The BITE test of the FQIC does a check of the Fuel Quantity lndicating System and the Fuel Level Sensing System (FLSS). The FLSCUs do not interface directly with the Centralized Fault Display System (CFDS).

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


MAINTENANCE/TEST FACILITIES Page: 41


FUEL

A320 NEO ATA 28

OPTIONS


There are some options for the fuel system which can be selected by operators. The Refuel/Defuel panel can also be installed at the wing leading edge. A cockpit Refuel panel can be installed. It will always take priority over the external Refuel/Defuel panel. Auxiliary fuel tanks can be installed in the aircraft. These Additional Center Tanks (ACTs) are installed in the cargo compartments and extend the range of the aircraft.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


OPTIONS Page: 43


FUEL

A320 NEO ATA 28

SAFETY PRECAUTIONS When you do work on the aircraft, make sure that you obey all the Aircraft Maintenance Manual (AMM) safety procedures. This will prevent injury to persons and/or damage to the aircraft. Here is an overview of the main safety precautions related to the fuel system. Aircraft fuel is poisonous. Do not get aircraft fuel in your eyes, mouth, nose, ears or on your skin. Use solvents, cleaning agents, sealants or other special materials only with good airflow in the work area. Use protective clothing to prevent personal contamination and formation of static electricity. Make sure that correct firefighting equipment is available. Make sure that the safety area is clear and clean. Obey the safety precautions in the safety area. Put ”NO SM()KING” warning notices around the work area. Ground (earth) and bond the aircraft. In the work area: S Do not use flames without protection and do not use any material or tools which can cause sparks, S Use only approved electrical / electronic equipment, S Make sure that the work area has sufficient airflow to do the work safely. If not, use a respirator, S Do not pull or move metal objects on the ground, S Immediately flush away or remove fuel leakage.


During refueling, do not transmit with the HF system This can cause fire or injury to personnel.

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FUEL

A320 NEO


ATA 28


SAFETY PRECAUTIONS Page: 45


FUEL

A320 NEO ATA 28

FUEL TANK SAFETY After three fuel tank explosions in recent decades, which caused in 346 fatalities, the U.S. Department of Transportation, Federal Aviation Administration (FAA), introduced new regulations to improve fuel tank safety. These regulations are related to the prevention of ignition sources in fuel tanks of current type certificated aircraft. To obey these regulations, a one--time fuel system safety and design review must be done. CRITICAL DESIGN CONFIGURATION CONTROL LIMITATIONS (CDCCL) The FAA issued Special Federal Aviation Regulation (SFAR) 88, which gives a detailed description of the CDCCL concept. The EASA requested the SFAR 88 (TGL 47) to be added to PART145, PART M and PART 147 to reinforce the application of these regulations. This includes: S A conception part intended to aircraft design features, S A maintenance part.


A CDCCL is a limitation requirement to preserve a critical ignition source prevention feature of the fuel system design that is necessary to prevent the occurrence of an unsafe condition. The function of the CDCCL is to give instructions to prevent critical ignition source feature from alterations, repairs or maintenance actions during configuration change. The aircraft manufacturers must supply a document to their customers to give the list of all the maintenance tasks impacted by the CDCCL. For AIRBUS this document is called the Fuel Airworthiness Limitations and it is added to the Airworthiness Limitation Section part 5. The CDCCL items are listed in the Airworthiness Limitations Form.

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FUEL

A320 NEO


ATA 28


FUEL TANK SAFETY - CRITICAL DESIGN CONFIGURATION CONTROL LIMITATIONS (CDCCL) Page: 47


FUEL

A320 NEO ATA 28

FUEL TANK SAFETY (CONTINUED)


FUEL SYSTEM DESIGN CONFIGURATION The Airbus aircraft fuel systems have, by design, a number of features that are intended to protect the system from inadvertent ignition.

SCL / REV.00 / Jun. 2016

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FUEL

A320 NEO


ATA 28


FUEL TANK SAFETY - FUEL SYSTEM DESIGN CONFIGURATION Page: 49


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G) B2 SCOPE CAUTION:

MODULE TAGGED B2 SCOPE. BE AWARE THAT ONLY AVIONICS/ELECTRICAL TOPICS SHOULD BE LEARNED FOR A T2 COURSE.


PRINCIPLE The temperature of the Integrated Drive Generator (IDG) oil is decreased by fuel through a recirculation system. Some of the fuel that supplies the engines is used to decrease the temperature of the IDG oil. A Fuel Return To Tank valve (FRTT) lets the hot fuel return to the outer cell. The FRTT opens the fuel flow back to the aircraft tank in special engine A configurations (N2, fuel flow.). The return valve mixes the hot fuel with cold fuel from the Low Pressure (LP) fuel pump to keep the temperature of the returned fuel less than 100_C (2l2_F). The Fuel Level Sensing Control Unit (FLSCU) l and the Engine Electronic Control (EEC) 1 control the recirculation system in the left wing. FLSCU 2 and EEC 2 control the right wing system.

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FUEL

A320 NEO


ATA 28


B2 SCOPE AND PRINCIPLE Page: 51


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G) FUEL RETURN The recirculated fuel is sent to the outer cell through a check valve and a pressure--holding valve to not let the fuel get to boiling temperature. The pressure--holding valve keeps a pressure of 15.5 psi in the return line. If the pressure increases, fuel bleeds through the valve into the outer cell. The check valve prevents fuel flow from the wing tank to the engine when the recirculation system is not in operation. WHEN THE OUTER CELL IS FULL, THE FUEL OVERFLOWS INTO THE INNER CELL THROUGH A SPILL PIPE.


NOTE:

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FUEL

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ATA 28


FUEL RETURN Page: 53


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G)


PUMP LOGIC While fuel is supplied from the center tank, the wing tanks will stay full and will possibly overfill because the returned fuel is supplied to the wing tanks. lf this occurs, the center tank transfer valves close when the inner cell gets to the FULL level sensor. The wing tank pumps will supply the fuel to the engine until approximately 500 kg (1100 lbs) of fuel are used and the UNDERFULL sensor is reached. The logic circuit then open the center tank transfer valves again.

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FUEL

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ATA 28


PUMP LOGIC Page: 55


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G) FUEL RETURN TO TANK VALVE CLOSURE


OVERFLOW The FRTT closes if the center tank transfer valves do not obey the logic signals of the full level sensors. This causes the wing tank to overflow through the tank ventilation system into the vent surge tank. The overflow sensor sends an electrical signal to the FLSCU. The FLSCU sends a closure signal to the EEC through the Engine Interface Unit (EIU). The EEC closes the FRTT and stops the fuel supply back to the outer cell.

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FUEL

A320 NEO


ATA 28


FUEL RETURN TO TANK VALVE CLOSURE - OVERFLOW Page: 57


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G) FUEL RETURN TO TANK VALVE CLOSURE (continued)


OUTER CELL HIGH TEMPERATURE The FRTT closes if the fuel temperature is too high in the outer cell, i.e. 52.5_C (126.5_F). Because the returned fuel from the engine is hot, the FLSCU prevents an overtemperature in the wing tanks. The FLSCU sends at closure signal to the EEC through the EIU. The EEC closes the FRTT and stops the fuel supply back to the outer cell.

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FUEL

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ATA 28


FUEL RETURN TO TANK VALVE CLOSURE - OUTER CELL HIGH TEMPERATURE Page: 59


FUEL

A320 NEO ATA 28



INNER CELL HIGH TEMPERATURE The FRTT closes if the fuel temperature in the inner cell is too high, i.e. 55_C (131_F). Thus a large volume of high--temperature fuel will not go into the inner cell if the intercell valve opens. This also keeps the fuel temperature at an acceptable level if a tank rupture occurs.

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FUEL

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ATA 28


FUEL RETURN TO TANK VALVE CLOSURE - INNER CELL HIGH TEMPERATURE Page: 61


FUEL

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PUMP PRESSURE LOSS The FRTT closes if a fuel pump Low Pressure (LP) is sensed by all pump pressure switches of one wing for the related engine when the crossfeed valve is closed, or if a fuel pump LP is sensed by all pump pressure switches of the two wings when the crossfeed valve is open. This is to stop the return fuel How during engine gravity feeding. LP is sensed by the pump LP switch and a signal is sent to the FLSCU.

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FUEL

A320 NEO


ATA 28


FUEL RETURN TO TANK VALVE CLOSURE - PUMP PRESSURE LOSS Page: 63


FUEL

A320 NEO ATA 28

FUEL IDG COOLING SYSTEM PRESENTATION (A319/A320 PW1100G) FUEL RETURN TO TANK VALVE CLOSURE (continued) LOW LEVEL The FRTT closes when the fuel level in the inner cell decreases to the INNER LOW LEVEL sensor at 280 kg (620 lbs). WHEN THE FRTT CLOSES, THIS DECREASES THE QUANTITY OF FUEL THAT CANNOT BE USED.


NOTE:

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FUEL

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ATA 28


FUEL RETURN TO TANK VALVE CLOSURE - LOW LEVEL Page: 65


FUEL

A320 NEO ATA 28



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ICE AND RAIN

A320 NEO ATA 30


ATA 30 ICE AND RAIN

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ICE AND RAIN

A320 NEO ATA 30

SYSTEM PRESENTATION (PW1100G) USERS The Nacelle Anti--Ice (NAI) System is designed to prevent ice formation on the engine inlet which could affect the engine operation. The engine air intake is heated during icing conditions using its related bleed air. The hot air is then discharged overboard. SOURCE Hot air for the Nacelle anti--ice system is supplied by a dedicated HP Compressor (HPC) bleed: S on the CFM--LEAP, 7th stage, S on the PW1000G, 6th stage. VALVE The NAI System is controlled and monitored by the (Propulsion Control System (PCS) (Engine Electronic Controller (EEC) and Engine Interface Unit (EIU)). Each engine NAI System consists of two electrically controlled, pneumatically operated Pressure Regulating and Shut--Off Valves (PRSOV). The EEC energizes the solenoid to CLOSE the PRSOV. Therefore, in case of loss of electrical power supply, the valves will go fully open provided the engine bleed air supply pressure is high enough. In the absence of air pressure, the valve is spring--loaded to the closed position.


CONTROLS When the ENG ANTI ICE P/B SW is selected ON, signals are sent to EEC for controlling the valves and to the EIU to calculate the bleed decrements. ECAM PAGE lf at least one of the two engine air intake anti--ice protection systems is selected ON, a message appears in green on the upper ECAM right MEMO. The EEC monitors the valve position through transducers and processes them to generate necessary indications and warning through the Flight Warning System (FWS). The FAULT indication in the PB S/W is activated by the PCS.

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ATA 30


USERS ECAM PAGE Page: 69


ICE AND RAIN

A320 NEO ATA 30

SYSTEM PRESENTATION (PW1100G) NAI SYSTEM Each engine air intake has its own independent Nacelle Anti--Ice (NAI) protection system. NAI System uses the hot blood air from a dedicated engine bleed port (6th stage High Pressure Compressor (HPC) for PW1100G). This blood air is load to engine air inlet through a food duct which passes along the RH side of the engine core and fan case. Each engine NAI system consists of one command P/B SW but two Pressure Regulating and Shut--Off Valves (PRSOV) for good operability, two Pressure Transducers (PTs), temperature protection and supply ducts. Both PRSOVs are located on the engine core, Right Hand (RH) side.


AIR INLET COWL The air is released into the air intake lip (D--Duct) through a swirl system which mixes the air and injects it in a specific pattern for effective heating. The airflow exits the air intake lip by a single exhaust grid at the bottom of the nacelle outside the fan which has 6 oval holes.

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NAI SYSTEM AND AIR INLET COWL Page: 71


ICE AND RAIN

A320 NEO ATA 30

ENGINE AIR INTAKE ICE PROTECTION SYSTEM DESCRIPTION (PW1100G) PRSOV CONTROL AND OPERATION The NAI system is controlled and monitored by the Propulsion Control System (PC S) (Engine Electronic Controller (EEC) and Engine Interface Unit (EIU)).The EEC controls the PRSOV operation by energizing/de--energizing the solenoids. PRSOV 1 is controlled by EEC Channel A and PRSOV 2 is controlled by Channel B. Each PRSOV pneumatically regulates the downstream air pressure. When the NAI PB S/W is selected to “ON” position, the EEC de--energizes the solenoid valves of PRSOV to OPEN the valves. Only when both the valves are open the bleed air is fed to the engine intake lip. The PRSOV 1 regulates the upstream pressure then in cascade PRSOV A 2 the downstream pressure at different threshold.


MONITORING The EEC does a detailed monitoring of the PRSOVs with two PTs (PT1 & PT2) located downstream each PRSOV. A PT1 is located in between the PRSOVs in the core engine area. It gives the feedback to channel B only and use for trouble shooting. PT2 is located downstream of PRSOV 2 in the fan case. it gives me feedback to both the EEC channels for monitoring function in case of single failure of EEC channel. A dual temperature sensor located in the fan case, provides the EEC (one per channel) with the fan compartment temperature measurement for NAI leakage detection. When the engine is running and a “Hot Air Leakage” event is detected, the EEC energizes PRSOVs solenoids, which provide insulation function.

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PRSOV CONTROL AND OPERATION MONITORING Page: 73


ICE AND RAIN

A320 NEO ATA 30

ENGINE AIR INTAKE ICE PROTECTION SYSTEM DESCRIPTION (PW1100G) ENGINE ANTI ICE P/B SW The P/B SW sends a discrete signal to the EEC to operate the PRSOVs. The P/B SW position and the opposite engine P/B SW position are monitored by the EIU for computing the bleed decrements. The ”FAULT” light is triggered by the EIU based on the input from EEC. It appears when the engine is running and NAI is failed in OPEN or CLOSED. It also appears in case of monitoring fault. PCS (EEC and EIU) The EEC controls the PRSOV to open when the P/B SW is set to ON. The EEC monitors the position of the PRSOV by the two NAI transducers to trigger associated fault messages. The System Data. Acquisition Concentrator Flight Warning System (SDAC/ FWS), Flight Data Interface and Management Unit (FDIMU) and Centralized Fault Display Interface tuna (CFDIU) interfaces with the PCS.


FAILURE CONDITION The fail safe position of the valves in case of EEC dual channel failure is OPEN. In case of a single valve failure, the corresponding valve being failed open, the anti--ice function is still available. The two pressure Transducers (PT1 for core zone and PT2 for fan zone) monitors leak or burst scenarios and a dual fan case thermocouple helps in identifying over temperature conditions due to leaks or burst. The EEC monitors the same and generates warning messages to the FWS. Master Minimum Equipment List (MMEL) IMPACT-- In case of both NAI valve failures, dispatching with one of the two valves locked close will not be possible.

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ICE AND RAIN

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ATA 30


ENGINE ANTI ICE P/B SW FAILURE CONDITION Page: 75


ICE AND RAIN

A320 NEO ATA 30



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PNEUMATIC

A320 NEO ATA 36


ATA 36 PNEUMATIC

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PNEUMATIC

A320 NEO ATA 36

PNEUMATIC SYSTEM PW1100G SYSTEM INTRODUCTION The Airbus Single Aisle family pneumatic system supplies High Pressure (HP) air for: S Air conditioning, S Wing ice protection, S Water Tank pressurization, S Hydraulic reservoir pressurization, S Engine starting, S Fuel tank inerting system. High Pressure air can be supplied from three sources: S The Engine Bleed system, S The APU, S A HP Ground Air source.


The pneumatic system operates electro--pneumatically and is controlled and monitored by 2 Bleed Monitoring Computers (BMC 1 & 2). There is one BMC for each engine bleed system. Both BMCs exchange data. In this NEO configuration, one BMC can control & monitor both sides when the other BMC fails.

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SYSTEM INTRODUCTION Page: 79


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SYSTEM INTRODUCTION (CONTINUED) ENGINE BLEED The Engine Bleed Air is pressure and temperature regulated before it supplies the pneumatic system. Air is bleed from an Intermediate Pressure (IP) stage HP3 or the HP8 stage with the High Pressure Valve (HPV) which is used for the pneumatic regulation. The IP check valve gives protection to the IP stage from reverse flow when the HP valve is open.


NOTE:

THE ENGINE BLEED AIR SYSTEM (EBAS) USES ELECTRO-PNEUMATIC VALVES.

APU BLEED/EXTERNAL AIR The left and right bleed systems are connected by a crossbleed duct. A Crossbleed Valve is used for their interconnection or isolation. The APU is mainly used for bleed air supply on the ground for air conditioning and for engine start. But APU BLEED air can also be used in flight, but limited in altitude. The APU bleed supply is connected to the left side ofthe crossbleed duct. On the ground, an HP Ground cart can be connected to the left side pneumatic system. The Crossbleed valve has to be opened to supply the right side.

The HP bleed is only used when the engines are at low power and for engine efficiency the High Pressure Valve (HPV) is kept closed during cruise. The Pressure Regulating Valve (PRV) regulates the bleed air pressure. The PRV is used as a protective shut off valve when the parameters are abnormal. In case of EBAS electrical failure, the PRV operates in back--up pneumatic mode. An Overpressure Valve (OPV) is installed downstream of the bleed valve to give protection to the system if an overpressure condition occurs. On this PW Engine the OPV is installed in the engine core. The Fan Air Valve (FAV) modulates Fan discharge air through an air--to--air heat exchanger called ”Precooler” to reduce the Bleed temperature . BMCs are Dual Channel computers. Each BMC channel A is a full digital channel embedding all the control and monitoring functions. Channel B is a hardware part and back--up channel able to detect system overtemperature. For the monitoring, the BMC s read pressure transducers (upstream / downstream ofthe PRV), Precooler Differential Pressure and downstream temperature with the Bleed Temperature Sensor (BTS).

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ENGINE BLEED AND APU BLEED EXTERNAL AIR Page: 81


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LEAK DETECTION Leak detection loops are installed along the hot air supply ducts of the pneumatic system. The loops are made of multiple sensing elements connected in series to the BMCs Overheat Detection System (OHDS). If a leak is detected, a signal is sent to the BMC 1 or 2 which automatically isolates the affected area by closing the crossbleed valve and shutting off the engine bleed on the affected side. The leak detection system is organized into three loops. Here are the loops and the protected areas: S Pylon: dual loop from the precooler to the wing leading edge. S Wing; dual loop from wing leading edge, including the wing air inlet supply, and belly fairing (cross bleed duct, pack supply ducts and APU forward supply duct). S APU: single loop at APU aft supply duct (left hand side ofthe fuselage) from APU firewall to wheel well area.

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CONTROL & INDICATING CONTROL & INDICATING This section is related to the control panels and indications for the pneumatic system.


CONTROL PANEL Controls for the pneumatic system are part ofthe AIR COND panel and are operated from the overhead panel.

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CONTROL PANEL Page: 85


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CONTROL & INDICATING (CONTINUED)


ECAM INDICATION The pneumatic system indications are displayed on the lower part of the ECAM BLEED page: S HPV, PRV positions with delivered bleed pressure and temperature, S APU bleed and crossbleed status.

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ECAM INDICATION Page: 87


PNEUMATIC

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MAINTENANCE/ TEST FACILITIES Using the Multipurpose Control and Display Unit (MCDU), you can have access to the Centralized Fault Display System (CFDS) fault messages of the PNEUMATIC system. BMC 1 and BMC2 Built--In Test Equipment (BITE) is standard type.

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MAINTENANCE/TEST FACILITIES Page: 89


PNEUMATIC

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SAFETY PRECAUTIONS When you do work on the aircraft, make sure that you obey all the Aircraft Maintenance Manual (AMM) safety procedures. This will prevent injury to people and / or damage to the aircraft. Here is an overview ofthe main safety precautions related to the pneumatic system. Make sure that the pneumatic system is depressurized before you start the work. HP air can cause unwanted pressurization of the aircraft, and injury to personnel. Be careful when you do work on the engine components immediately after the engine shutdown. The engine components can stay hot for one hour.

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PNEUMATIC

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PNEUMATIC SYSTEM COMPONENT LOCATION SYSTEM OVERVIEW The Pneumatic system is used to supply High Pressure (HP) air for air conditioning, pressurization, Fuel Tank Inerting System (FTIS), engine start and anti--icing. HP air can be supplied from the two engines, the APU or an external ground source.


ENGINE BLEED The engine bleed air is pressure regulated and temperature controlled before it supplies the aircraft pneumatic system. Air is bleed from the engine High Pressure Compressor (HPC) stages: HP3 via an Intermediate Pressure Check Valve (IPCV) and HP8 via the HP Valve (HPV). The High Pressure Bleed Valve (HPV) supplies air to the system when the engine is at low power. When the Intermediate Pressure (IP) bleed is sufficient, the HPV closes. The bleed valves are electro--pneumatically controlled. The Pressure Regulating Valve (PRV) installed downstream the IPCV and HPV regulates the bleed pressure. Each Bleed Monitoring Computer (BMC) controls and monitors its engine bleed system and the opposite. An Overpressure Valve (OPV) is installed downstream from the bleed valve as a protection of the system if the pressure is too high. The engine bleed air is temperature regulated. The hot bleed air goes through an air--to--air heat exchanger called Precooler. Fan discharge air, modulated by the Fan Air Valve (FAV), is blown across the precooler to keep the temperature within limits.

APU BLEED/EXTERNAL AIR The left and right bleed systems are connected by a crossbleed duct. A Crossbleed valve is used for their interconnection or isolation. The APU can also be used for bleed air supply. This is usually done on the ground for air conditioning and for engine start. But APU BLEED air can also be used in flight, in relation to the altitude. The altitude can be different for each aircraft. These altitude limits are given by the manufacturer. The APU bleed supply is connected to the left side ofthe crossbleed duct. On the ground, a HP ground power unit can be connected to the left side pneumatic system. The right side can be supplied by opening the crossbleed valve.

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ENGINE BLEED AND APU BLEED/EXTERNAL AIR Page: 93


PNEUMATIC

A320 NEO ATA 36

SYSTEM OVERVIEW (CONTINUED)


LEAK DETECTION Leak detection loops are installed along the hot air supply ducts of the pneumatic system and are connected to the BMCs. The leak detection system is organized into three loops. Here are the loops and the protected areas: S PYLON: the precooler outlet area, S WING: wing leading edge and belly fairing, S APU: APU aft supply duet (left hand side ofthe fuselage) from APU firewall to wheel well area.

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LEAK DETECTION Page: 95


PNEUMATIC

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COMPONENT LOCATION The primary components ofthe pneumatic system are installed on the engines and in the pylons. PRESSURE REGULATION COMPONENTS The pressure regulation components on the engines are the: S Engine HPV, S Engine BLEED PRV, S OPV, S Bleed Monitoring Pressure Sensor (BMPS), S Bleed Pressure Sensor (BPS), S Differential Pressure Sensor (DPS).


To get access, open the right fen cowl and thrust reverser cowl.

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PRESUURE REGULATION COMPONENTS Page: 97


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COMPONENT LOCATION (CONTINUED)


TEMPERATURE REGULATION COMPONENTS The pressure regulation components are in the pylons: S the FAV S the Precooler, S the Bleed Temperature Sensor (BTS).

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COMPONENT LOCATION (CONTINUED)


OTHER COMPONENTS The Crossbleed valve is in the forward section ofthe lower fuselage belly fairing area. The access to the HP ground connector is through a small access door on the lower fuselage belly fairing. The APU bleed valve is on the APU. The APU supply duct is installed along the left hand side of the fuselage to the wheel well area and is connected to the crossbleed duct in the forward belly fairing area.

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OTHER COMPONENTS Page: 101


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) B2 SCOPE CAUTION:

MODULE TAGGED B2 SCOPE BE AWARE THAT ONLY AVIONICS/ELECTRICAL TOPICS SHOULD BE LEARNED F OR A T2 COURSE.


GENERAL The Engine Bleed Air System (EBAS) supplies pressure and temperature regulated airflow from each engine to the air system users. During normal operation, each engine bleed system is isolated from adjacent system by the Crossbleed valve; except during 2nd engine starting using air bled from 1st started engine, Crossbleed. valve opened or under APU Bleed. The pressure regulation system is controlled and monitored by two Bleed Monitoring Computers (BMCs). As compared to A320 CEO, the NEO engine has higher bleed air temperatures during High Pressure (HP) operation, lower air pressure during Intermediate Pressure (IP) Operation, lower fan pressures for cooling air How Supply and limited space for installation due to new pylon configuration. To achieve better performance requirements a new electro--pneumatic bleed air system is designed for A320 NEO.

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B2 SCOPE GENERAL Page: 103


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G)


BMC Normally BMC 1 Channel A does all the control and monitoring of the LH EBAS and BMC 2 Channel A the RH EBAS. Each BMC channel A controls torque--motor and solenoid for the electro--pneumatic valves, monitors sensors. As both BMC interface, each one is capable to control both sides. The channel B is a fully hardware part able to detect the system overtemperature: Electrical Protection Function (EPS). This detection is fully independent from software part. Each BMC reports the failures independently of each other.

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BMC Page: 105


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) HPC HP VALVE (HPV) The engine air bleed pressure is pneumatically regulated by the HP Valve (HPV) when air is supplied by the High Pressure Compressor (HPC) stage or directly by the Pressure Regulating Valve (PRV) when the air is supplied by the Intermediate Pressure (IP) HPC stage. S Intermediate--pressure service port; IP is defined by HP3. S High--pressure service port: HP is defined by HP8. The HPV lets air to be bled from the engine HP stage at lower power settings. It is a pressure regulating and shut--off valve with a butterfly closure element. It regulates the pressure of the bleed air between 15 and 65 psig. With the Solenoid energized, the minimum upstream muscle pressure needed to operate the valve is 15 psig. When the solenoid is not energized, the HPV is commanded to the full closed position. When the solenoid is energized but without pressure in the valve body, the HPV stays closed. The HPV is forced to close when the PRV is closed. The valve has a manual override and test port for pneumatic test in--situ.


IP CHECK VALVE (IPCV) An Intermediate Pressure Check Valve (IPCV) lets air to be bled from the engine IP stage. It is closed when air is bled from HP stage. The purpose of this IPCV is to allow the flow from IP stage and avoid the reverse flow from either the HP port or the pneumatic manifold.

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HPC-- HP (HPV) AND IP CHECK VALVE (IPCV) Page: 107


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) PRESSURE REGULATING VALVE (PRV) The Pressure Regulating Valve (PRV) is a 4 inch diameter butterfly valve, installed downstream of the IPCV and HPV. It regulates the pressure ofthe bleed air at 42 ±2 psig in normal dual bleed operation (50 ±2 psig in single bleed operation). Its setting is modulated by the electric command on the torque--motor. When the torque--motor is de--energized, the PRV is commanded to the full closed position. When the torque--motor is energized but without pressure, the PRV stays closed. With the torque--motor energized, the minimum upstream muscle pressure needed to operate the valve is 15 psig. The PRV operates as a shut off valve when abnormal conditions occur. In case of electrical failure of the EBAS, pressure control is ensured by the PRV in back--up pneumatic control mode. The valve has a manual override and test port for pneumatic test in--situ.


OVERPRESSURE VALVE (OPV) The Overpressure Valve (OPV) downstream ofthe PRV in the engine core, protects the system against damage if overpressure occurs. It operates pneumatically. The OPV, normally in spring--loaded open position will be fully closed if bleed pressure reaches 90 psig. The valve has a manual override and test port for pneumatic test in--situ.

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PRESSURE REGULATING VALVE (PRV) - OVERPRESSURE VALVE (OPV) Page: 109


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) PRESSURE SENSORS BLEED MONITORING PRESSURE SENSOR (BMPS) The Bleed Monitoring Pressure Sensor (BMPS) is used to perform bleed port switching function. It is also used to estimate the position ofthe HPV butterfly and to monitor the HPV and the PRV. BLEED PRESSURE SENSOR (BPS) The Bleed Pressure Sensor (BPS) is installed downstream the PRV. It provides to BMC the actual bleed air pressure delivered through the PRV. This sensor is also used by the BMC for system monitoring (overpressure and low pressure alarms) and to monitor the position ofthe OPV butterfly.


DIFFERENTIAL PRESSURE SENSOR (DPS) The Differential Pressure Sensor (DPS) ensures the reverse flow protection by sensing the differential pressure between Precooler hot side inlet and outlet. It also provides to the BMC an indication of the PRV and OPV position.

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BLEED MONITORING PRESSURE SENSOR (BMPS) DIFFERENTIAL PRESSURE SENSOR (DPS) Page: 111


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) BLEED TEMPERATURE SENSOR (BTS) The dual Bleed Temperature Sensor (BTS) installed downstream the Precooler provides to the BMC the actual EBAS temperature. The BMC uses EBAS temperature to position the Fan Air Valve (FAV). The wiring connected to channel A ofthe BTS is fully segregated from the wiring connected to channel B. Both BMCs interchange temperature measurements and can carry out both sides temperature regulation. This dual sensor is also used by the BMCs for system monitoring (overtemperature and low temperature alarms). CHANNEL B OF ONE BMC IS CONNECTED TO CHANNEL A OFTHE OTHER BMC, SO THAT IN CASE OF LOSS OF TEMPERATURE MONITORING AND CONTROL IN CHANNEL A OF ONE SIDE, THE OPPOSITE CONTROLLER CAN TAKE OVER CONTROL OFTHE WHOLE EBAS.


NOTE:

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BLEED TEMPERATURE SENSOR (BTS) Page: 113


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) TEMPERATURE REGULATION FAN AIR VALVE (FAV) The FAV pneumatically regulates the fan airflow to the Precooler for bleed air temperature regulation. The FAV butterfly valve actuator rod is adjusted by the BMC via a torque motor servo--control depending on BTS input. The BMC set point is 200_C (392_F) in normal operations and 160_C (320_F) in Climb and Hold with 2 bleeds and Wing Anti--Ice (WAI) off. With no electrical power and enough muscle pressure, the FAV valve is fully open. The valve has a test port for pneumatic test in--situ.


PRECOOLER EXCHANGER The Precooler is a stainless steel and nickel alloy air--to--air heat exchanger. It cools down the hot air supplied from the engine HP compressor stage by a heat exchange process with cooling flow taken from the engine fan.

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TEMPERATURE REGULATION - FAN AIR VALVE AND PRECOOLER EXCHANGER Page: 115


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ENGINE BLEED SYSTEM DESCRIPTION (PW1100G) PROTECTION -- ISOLATION The PRV operates as a shut--off valve. It is commanded to close in the following conditions: S Overtemperature downstream ofthe Precooler (BTS): 257_C (495_F) < T 270”C (518_F) during 55s, 270_C (518_F) < T 290_C (554_F) for 15s, T > 290_C (554_F) for 5s. S Overpressure downstream ofthe PRV > 60± 3 psig at BPS, S Engine fire (consequence of crew action on the ENG FIRE P/B), S Leak detection in pylon/wing/fuselage ducts surrounding areas, S APU bleed valve not closed & APU BLEED P/B selected


Depending on the Crossfeed Bleed Valve (CBV) position, only one PRV (left engine PRV if CBV is closed) or both (if XBleed is open). S Reverse flow detected by DPS, S ENG BLEED P/B selected OFF or ENG not running, S Associated Starter Air Valve (SAV) not closed, S HPV failed open, S Dual BTS channels failed.

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BMC INTERFACES (PW1100G)


BMC The pneumatic system uses 2 identical controllers with a microprocessor and command channel A and a back--up channel B. Each channel is supplied by a different 28V DC bus bar. Both Bleed Monitoring Computers (BMCs) will work as MASTER/SLAVE so long as the ARINC429 cross communication is working properly. lf one ARINC429 bus is lost from one BMC to the other, the BMC receiving no data will take over control and would inform to the opposite BMC.

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EIU The Propulsion Control System (PCS) informs both BMCs via both Engine Interface Units (EIUs) when engines start/run. The Electronic Engine Control (EEC) will need information relative to the Aircraft Environmental Control. System (ECS) from the EIU ARINC data bus as system bleed pressure, bleed and anti--ice configuration. The EIUs receive positions of ENG BLEED P/Bs ON, APU BLEED P/B OFF, Crossbleed valve status.

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DATA LOADING The up and down data loading system is an interface between the onboard computers as BMCs and the ground--base data processing stations. For data loading purposes, the BMC 1 Channel A is connected to Data Loading Routing Box (DLRB). The BMC 2 Channel A will be loaded through BMC 1 Channel A, The BMC 2 will be uploaded through the crosstalk bus from the BMC l once the BMC 1 has been fully uploaded from the data loader.

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ACSC The BMC inform the Air Conditioning System Controller (ACSC) on the precooler outlet temperature for pack flow calculation. The bleed pressure Sensor (BPS) and the wired Crossbleed valve position are used for Pack Inlet Pressure Sensor (PIPS) monitoring. The BMC send a discrete input of its Pressure Regulating Valve (PRV) position. Another discrete signal informs about the precooler delivered bleed pressure. The ACSCs input the BMCs for Pack 1/2 P/B SW position, Pack Inlet Pressure and wing anti--ice valves position.

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DISPLAY The BMCs 1 and 2 transmit ARINC signals to the System Data Acquisition Concentrator (SDAC) for monitoring, fault indication, warning and data recording purposes by the Flight Warning Computer (FWC), Electronic Instrument System (EIS) and Digital Flight Data Recording System (DFDRS), The Centralized Fault Display Interface Unit (CFDIU) is connected to the BITE ofthe BMCs to centralize the pneumatic system data for maintenance via the Multipurpose Control and Display Units (MCDUs), printer and Aircraft Communication Addressing and Reporting System (ACARS).

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APU The APU/Electronic Control Box (ECB) system sends to the Engine Bleed Air System EBAS/BMC the information about APU bleed valve position in order to command the PRV to close when APU BLEED P/B is ON. The EBAS transmits to the ECB information related to the APU Bleed Valve open Command in order to provide APU Bleed valve control in when APU flow is required.

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) ROUTING The leak detection system is used to detect leaks in the vicinity of the packs, wings, pylons and APU hot air ducts. There are two independent loops as redundancy in both pylons and both wing sides. The APU hot air duct is monitored by a single loop. Protected areas with a double loop for: S Engine 1 and engine 2 pylons, S RH wing and pack 2, S LH wing and pack 1 and mid fuselage APU duct. Protected areas with a single loop for: S APU duct. EACH LOOP CONSISTS OF SENSING ELEMENTS CONNECTED IN SERIES. BOTH EXTREMITIES OF THE OVERHEAT DETECTION LOOP ARE CONNECTED TO THE BMC.


NOTE:

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PNEUMATIC LEAK DETECTION SYSTEM D/O (PW1100G) DETECTION LOGIC Both Bleed Monitoring Computers (BMC s) permanently receive signals from the leak detection loops primarily tested at power--up. They exchange data via an ARINC bus for the double loop detection. Each BMC channel A normally controls its side engine bleed air system, so monitors the Overheat Detection System (OHDS). THE WING AND PYLON LOOPS A ARE CONNECTED TO ONE BMC AND WING AND PYLON LOOPS B TO THE OTHER BMC. THE CROSSTALK BUS ALLOWS WING LEAK WARNINGS TO BE ACTIVATED THROUGH AN AND LOGIC. THE APU LOOP IS CONNECTED TO BMC 1 ONLY.


NOTE:

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) WARNING CONSEQUENCES The ENG BLEED FAULT light comes on when a leak is detected by the wing loops A and B or by the pylon loops A and B. The APU BLEED FAULT light comes on when an APU duct leak is detected. When an overheat condition is detected by both loops, the following alerts are generated for the affected zone: S AIR ENG 1(2) LEAK for a leak/overheat detected in the Pylons, S AIR L(R) WING LEAK for a leak/overheat detected in the Wings, S AIR APU LEAK for a leak/overheat detected in the APU line, S AIR APU LEAK [APU LEAK FED BY ENG] for a leak/overheat detected in the APU line and the leak is automatically isolated.


A new warning alert has been introduced on the A320NEO, the AIR BLEED LEAK to isolate a bleed leak in the opposite pylon to the operative bleed with manually open Crossbleed Valve. The failure of a single loop for Pylon or Wing is identified by a MAINTENANCE message displayed on the STATUS SD page. Dual engine loop failure is identified by the AIR ENG 1(2) LEAK DET FAULT and is NO GO. lf one BMC is failed, the other BMC takes over monitoring of the bleed system and triggers the ECAM warnings. The aircraft dispatch is for 10 days with the BMC 1 inoperative for non--ETOPS operations provided that the Engine l Bleed Air System (EBAS l) is considered inoperative and the APU leak detection loop is considered inoperative. LEAK CONSEQUENCE A detected leak will close associated valves, as shown on the table. These valves are automatically controlled to close if they were open. NOTE:

APU AND CROSS BLEED (X--BLEED) VALVES DO NOT CLOSE DURING MAIN ENGINE START (MES).

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PNEUMATIC SYSTEM LINE MAINTENANCE (PW1100G) MEL ITEMS EBAS MEL The aircraft dispatch is for 10 days with the Engine Bleed Supply System inoperative on one side provided that: S The associated bleed is isolated by setting the ENG BLEED P/B SW to OFF, S The X--BLEED valve is manually open to supply both sides, S The speed brakes are operative.

the Crossbleed valve in order to supply both sides from the opposite EBAS which is operative.

For an Extended Range Twin Engined Aircraft Operations (ETOPS) flight, Auxiliary Power Unit (APU) Bleed should be available. One Engine Bleed Air System (EBAS) remaining available, it supplies both sides for Wing Anti--lce (WAI) and air conditioning. However, there is limitation on A320 NEO compared to A320 CEO due to lower capacity ofthe heat exchanger in case of single bleed operations.


NOTE:

ONLY ONE PACK CAN BE SUPPLIED. THEREFORE, THE ASSOCIATED OPERATIONAL PROCEDURE WILL ASK TO SWITCH ONE PACK OFF.

HPV FAILURE Failed closed High Pressure Valve (HPV) can lead to low bleed pressure or low bleed temperature when engine is at low power settings (in idle or in holding conditions). HPV failed in open position, leads to Bleed overpressure or Bleed overtemperature identified by AIR ENG 1(2) BLEED FAULT. In case of failure of one HPV, the aircraft can be dispatched for 10 days with the valve secured closed. The consequence of having the HPV secured closed is that the bleed air from the Intermediate Pressure (IP) port will be insufficient at low engine power settings (taxi, descent, holding). That is the reason why the crew procedure requests to switch off the associated EBAS at low power setting and to open

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PNEUMATIC SYSTEM LINE MAINTENANCE (PW1100G) MEL ITEMS (continued)


BLEED VALVE DEACTIVATION In case of failure, Pressure Relief Valve (PRV) and HPV have to be deactivated CLOSED for dispatch under Minimum Equipment List (MEL). The deactivation procedure is the same for both valves: S make sure pneumatic system in not pressurized, BLEED switches OFF, S deactivate the thrust reverser S open the RH fan and reverser cowls, S move the manual override to the CLOSED position, S secure in CLOSED position with locking pin, S close cowlings, S reactivate the thrust reverser,

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) MEL ITEMS (continued)


WING LEAK DETECTION The WING leak detection is a dual--loop system. To generate a WING LEAK warning, both A and B loops have to detect the overheat. For dispatch, WING leak detection must be operational (at least one loop) on each wing. If a single loop fails, the MAINTENANCE message AIR BLEED will be displayed on the STATUS page associated with a Centralized Fault Display System (CFDS) message L(R) WING LOOP (INOP). The aircraft may be dispatched per MEL with the MAINTENANCE message displayed. For troubleshooting it is important to understand that the WING detection elements monitor much more than just the wings alone. The protected areas are: S wing leading edge (wing anti--ice supply duct), S air conditioning compartment--belly fairing -- (pack supply, crossbleed manifold, APU supply, ground air supply), S APU forward supply duct (from the APU check valve through the wheel well).

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) MAINTENANCE TIPS CFDS CFDS menus for all failure reports and interactive mode displays are generated by the, Bleed Monitoring Computer (BMC) itself. In normal mode, the BITE transmits maintenance messages (Standard A type 1) for detection results on level of: S OverHeat Detection System (OHDS), S Valves, S Precooler, S Sensors, S External communication, S Internal communication, S BMC (Hardware and Software).


The electrical test verifies the EBAS following functions: S Central Processing Unit (CPU) (microprocessor, RAM, ROM), S discrete outputs, S leak detection loops and interfaces, S discrete and analog inputs, S digital Inputs! Outputs, S torque motors, solenoid, S pressure sensors failures, S temperature sensors failures, S valves. The pressure sensor drift test shall detect any pressure drift in Differential Pressure Sensor (DPS) and/or Bleed Pressure Sensor (BPS). Electrical Protection System (EPS) corresponds to the channel B Electrical Protection Function (EPF) test. The reports menu displays the status in real time for all the system.

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) MAINTENANCE TIPS (continued)


TEST SET The Test Set P/N 98L36103002000 is available to assist in troubleshooting the pneumatic system. The test set enables calibrated pressure to be applied to individual valves, components and isolated parts of the system to check for normal operation and sense line integrity (i.e.; PRV HPV, Overpressure Valve (OPV), Fan Air Valve (FAV), Bleed Pressure Regulated Transducer).

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PNEUMATIC LEAK DETECTION SYSTEM SYSTEM D/O (PW1100G) MAINTENANCE TIPS (continued)


ENGINE START WITH GROUND AIR To perform an engine start with ground air, the connection is located on the lower fuselage. The access door is on the belly fairing. During a ground air start, the crossbleed valve must be operated manually. For safety, it is recommended to use the ground air supply to start the first engine. Then disconnect the ground air supply and perform a crossbleed start for the second engine. On the ECAM BLEED page, the GND indication DOES NOT indicate ground air supply connected or available. This indication appears when the aircraft is on the ground to show that the ground air is directly supplied to the LEFT side of the system only. The left bleed system pressure indicator will indicate pressure when the ground air is supplied.

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