SECTION 5 - FULL AUTHORITY DIGITAL ENGINE CONTROL
(ATA Chapters 73 and 76)
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Full Authority Digital Engine Control
Powerplant Electrical Connections
The diagram opposite details the electrical wiring diagram
for units located in and around the fan case of the engine.
The units identified are discussed in this section of the
course notes.
Units relative to this section
Units not relative to this section
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Full Authority Digital Engine Control
3030VC
T/R 3rd lock
suitcase
discon
4069KS
TAI HP
Switch
3804VC
Pylon/EEC
Channel A
3028VC
T/R DCU
Suitcase
Dsicon.
J23
4070KS
4072KS
PCU
OPU
A
4005KS
Start Air
Valve
4166KS
T/R Ground
Safety
Switch
J22
A
J3
(Channel B)
J4
(Channel A)
B
A
A
A
Fancase/core
Discon.
3009VC
B
JA8
3026VC
Junct.box
Suitcase
Discon.
4060KS2
Oil Press.
Xmitr
4018KS1
TFuel
4117KS
Isol.Valve
Press.SW
JA6
4000KS
3024VC
Junct.box
Suitcase
Discon.
JA9
JA10
3801VC
PylonEEC
Channel B
3006VC
N1/N2 (T) 3
Discon.
3029VC
T/R DCU
suitcase
Discon.
B
J1
A
A
3002VC
P/T20
Nosecowl
Discon.
A
B
4018KS2
TFuel
EEC
JA7
A
4071KS
FMU
JA5
JA3
JA4
A
A
4060KS1
Oil Press.
Xmitr
A
3010VC
Fancase/core
Discon.
JA1
JA2
4041KS
Main Oil
Filter - Diff
Press.Switch
3001VC
P/T20
Nosecowl
Discon.
A
J2
J3
4073KS
AOHE
LVDT
B
4010KS
Fuel Flow
xmtr
J4
4073KS
AOHE
Torque
Motor
A
4042KS1
Oil Temp
T/Couple
A
A
A
4043KS
Fuel
Filter
DPSwitch
4050KS
Oil Qty
Xmitr
B
4076KS
Scav Oil
Filter Switch
4075KS
Fuel Low
Press.
Switch
4042KS2
Oil Temp
T/Couple
FADEC Electrical Connections (1)
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Section Objectives
On completion of this section, you should be able to:•
Define the FADEC system.
•
Identify the main electronic units in the FADEC
system.
•
Differentiate between the different methods of
overspeed protection.
•
State the components and areas of control that
form the FADEC system.
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FADEC System
Full Authority Digital Engine Control
Rotor Integrity Protection
Introduction
• Overspeed Protection Unit (OPU)
A Full Authority Digital Engine Control system (FADEC)
controls the RB211-Trent engine.
• Turbine Overspeed Protection (LPTOS)
The FADEC system is made of sub-systems working
together to form a closed loop control system, maintaining
efficient engine operation at a selected condition ranging
from engine start, through the take-off flight/landing
operation envelope to engine shut-down.
Note Shutdown is effected by the Pressure Raising and
Shut-Off Valve (PRSOV) torque motor (see Fuel System),
which is hardwired directly from the cockpit, giving the pilot
the ability to operate the valve at any time. This function
has priority over any automatic PRSOV command.
The sub-systems are:
Power Generation and Conditioning Control
• Dedicated Alternator
• Power Control Unit
Engine Control
• Electronic Engine Control Unit (EEC)
• Discrete Sensors
• Torque Motors
• Solenoids
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Power Generation and Conditioning Control
Dedicated Alternator
The Dedicated Alternator supplies primary power to the
FADEC system and provides a speed reference signal of
the HP shaft speed (N3).
The unit is mounted on the external gearbox and driven by
direct drive from the HP shaft (N3).
The Alternator consists of two separate three phase stator
windings and two separate single phase stator windings.
The associated rotor magnets are connected to a common
cantilever shaft. (The shaft does not require bearings).
The three phase circuits provide power to the EEC in the
speed range 8% to 125% N3. One of the phase windings
in each three phase circuit provides the EEC with
referencing to the HP shaft rotational speed.
The two separate single windings provide power to the
overspeed protection unit.
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Power Generation and Conditioning Control
Power Control Unit (PCU)
The primary source of EEC power is provided by the
Dedicated Alternator. A standby electrical supply is
provided by two aircraft 115V AC supplies.
The aircraft 115V AC supply provides power:
• During engine start upto the Dedicated Alternator
coming on line at a minimum HP shaft speed
• System failure
• Ground test.
The EEC power supply is fed into the Power Control Unit
(PCU). The Dedicated Alternator power or aircraft power
as appropriate is conditioned by the PCU.
The PCU is a dual channel unit, it rectifies, filters and
regulates power supply to the respective channel of the
Electronic Engine Control Unit to 22V DC.
The EEC/PCU also controls and monitors the switching of
Aircraft power to the following:
• Igniter circuits (115V AC supplied via the EIVMU)
• P20/T20 Probe Heater (115V AC supplied directly from
the Aircraft electrical network
• Cabin Air Bleed HPV switching (20V DC derived from
Aircraft 115V AC supplied via the EIVMU)
• Thrust Reverser Directional Control Valve (28V DC
supplied via the EIVMU)
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Power Generation and Conditioning Control
Power Control Unit (PCU)
The Power Control Unit (PCU) is located on the left side of
the fan case inside the EEC suitcase.
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Full Authority Digital Engine Control
Engine Control
Electronic Engine Control Unit (EEC)
The Electronic Engine Control (EEC) unit is the heart of
the FADEC system. It is located on the fancase and is
shielded and grounded as protection against (EMI)
Electromagnetic Interference.
• Heat Management control
The EEC unit is a dual channel digital unit. The channels
are identified as channel A and channel B with a
communication link between each.
• P20/T20 probe heater control
Each channel uses the high integrity computer (HIC)
concept to perform software instructions and utilises dual
interfaces to provide a high degree of fault tolerance.
In normal operation only one channel controls, in the event
of certain failures control is transferred to the alternative
channel.
• IP Turbine case cooling control
• Aircraft cabin air HP valve control
• The FMU in the event of LP shaft breakage
• Thrust reverser control
• Maintenance function management
• Process various parameters for data transmission to the
Aircraft
• Limit protection for N1, N2, N3 and EGT
The EEC also transmits engine performance data and
system test data to the aircraft which is used in flight deck
display, thrust management and condition monitoring
systems.
The Fadec system provides all the necessary Engine
control functions and operate in association with the
appropriate Aircraft subsystems to perform the following:
• Engine starting (Ground & Flight)
• Relighting following flame-out detection
• Control of fuel flow
• VIGV/VSV control
• Handling bleed valves control
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Full Authority Digital Engine Control
Engine Control
Electronic Engine Control Unit (EEC)
The EEC is a single unit, mounted using anti-vibration
mountings. It is located on the left hand side of the engine
fan case with the PCU and OPU in a protective box
(suitcase) with a hinged lid.
The two channels of the EEC are almost the same, each
channel contains circuit boards or cards which match the
cards in the other channel. Each card has a specific
function related to it.
EEC input data is received from the aircraft and engine
sensors. All input data including signals from pressure
transducers is checked and as necessary, converted to
digital format. The inputs are also subject to radio
frequency interference (RFI) filtering and lightning strike
voltage protection.
The cards within an individual channel are inter-connected
by a mother board, each mother board is connected by
electronic wiring to the engine and the aircraft.
A wiring loom within the EEC connects Channel A to
Channel B. There is also a pressure module which
contains pressure transducers and transducer interface
circuits. Each transducer is connected to pressure signal
pipes from the engine.
On the side of the EEC unit is a connector for a Data Entry
Plug (DEP) which is attached when the EEC is installed on
an engine. The EEC computer reads the data and applies
it to adjust the engine operation characteristics.
An external test socket is incorporated in the EEC to allow
external test signals to communicate with the EEC.
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E.E.C. - Integrity
The dual channels, both interfacing with the aircraft
provide a high level of fault tolerance.
This is achieved by using duplicated engine hardware, and
EEC dual channel software.
Each channel of the EEC contains a High Integrity
Computer (HIC). This consists of a control computer and a
monitor computer.
Both control and monitor computers access memory,
internal data and input data from the airframe and engine
sensors. Both computers process the data independently
and should produce identical output. Output from both
computers is fed into a comparator (parity memory) which
is also within the HIC, any discrepancy results in a request
for a channel change and a reset of the faulty channel.
The channel change is achieved without interruption to
engine control.
The EEC contains multi-programmes of engine model
logic. Aircraft/Engine input data is continuously fed into
these programmes for comparison. Disagreements are
processed to generate corrective output signals which are
fed to engine mounted units which control engine
operation.
Contingency programmes are built into the EEC to cater
for channel failure. In addition the EEC is programmed
with logic to synthesise any primary control signal loss
from available data. The FADEC system provides a high
level of fault tolerance ie. following a signal failure, loss of
redundancy occurs rather than loss of function. This
provides a system which is safe and reliable allows
tolerance of faults for a limitedperiod of time.
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EEC Architecture - Input Types
The EEC inputs can be generally categorised into one of
three types:
•
Duplex - This type of input has a characteristic
of both channels having their own independent
input. The benefit from Duplex inputs is that
single internal or external failures can be
accommodated.
•
Simplex, Cross Wired - This type of input has a
single input to one channel of the EEC which is
then wired to the other channel of the EEC
before the data is processed. In this type of
input, single internal failures can be
accommodated, however, single external
failures are not.
•
Simplex - This type of input has a single input to
one channel of the EEC which processes the
information before making it availble to the other
channel on the Interchannel Highway. With this
input, single internal or external failures are not
accommodated.
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a) Duplex
b) Simplex, Cross-wired
ttr70451
a) Simplex
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EEC Architecture - EEC Input Types
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EEC Architecture - Output Tpyes
All the outputs from the EEC are duplex.
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ttr70452
All EEC Outputs are Duplex
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EEC Architecture - EEC Output Types
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Confirmed and Unconfirmed Fault Messages
The EEC has the capability of storing 64 flight legs of
faults, up to a maximum of 200 faults. A fault shall be
recorded the first time it is detected in a flight leg , even if
it was recorded in the previous leg(s).
However, the sytem requires both the detection and
confirmation of a fault , the check of confirming or not is
made against 3 timing points.
The diagram opposite shows that once a fault has
occurred, there is a momentary delay (between t0 and t1)
in order for the system to detect the fault. This delay is
determined by the rate at which the fault can be read by
the software.
At time t2, the effect occurs and if applicable there will be
a cockpit effect.
Between t2 and t3 there will be an additional delay prior to
the generation of the maintenance message, this delay is
a result of the software's fault intergration time.
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Example a) below shows a fault that occurs for a short
period of time that is less than the delay t1-t0, no
maintenance message will be sent.
Example b) below shows a fault that lasts more than time
t1 but is less than time t2. As with example a), no
maintenance message will be sent.
Example c) below shows a fault that has a time greater
that t2 but less than t3, this time a maintenance message
will be sent at time t2. (However to register the fact that
the fault was < t3, control bit (19) in the EEC is set to 0).
Example d) below shows a fault that now exceeds t3, a
maintenance message will be sent at time t3 and the
control bit (19) will be set to 1.
NOTE
Bit 19 is part of the ARINC 429 STX (Start of
Transmission - label 355/357) word that is sent with the
fault message. If Bit 19 is set to zero it indicates an
unconfirmed fault, 1 indicates the fault has been
confirmed.
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FAULT
OCCURENCE
FAULTDETECTION
AND
ACCOMMODATION
COCKPIT
EFFECT
FAULT
CONFIRMATION
t0
t1
t2
t3
Fault
a)
ok
Fault
b)
ok
Fault
c)
ok
Fault
d)
ttr70453
ok
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Full Authority Digital Engine Control
System Criticality - Fault Detection
This page has the first list of FADEC parameters and their
criticality.
The list define the parameter, the configuration input to the
EEC, how the EEC validates the parameter and what the
parameter is used for.
Validation techniques used by the EEC to verify the
accurracy of each input is a s follows:
•
PARAMETER RANGE - the parameter data
is checked against limits within the EEC and
to be valid must fall with the upper and lower
limits. For example, P20 has a parameter
range check of 1.5 to 23 PSIA.
•
HARDWARE RANGE - this check compares
the measurement of a parameter against the
limits for the hardware.
•
CROSS CHECK - data can be crosschecked against different items to confirm
validity. For example, the P20 measurement
is cross-checked against the other channel
in the EEC, it is also cross-checked against
Pt from the ADIRUs.
•
MODEL Check - some of the data input into
the EEC is checked against a mathematical
model within the EEC software that is based
on other engine data.
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Full Authority Digital Engine Control
Sys te m C o n fig u r atio n
Simplex
V alid atio n T e ch n iq u e s
Cros sControl / Monitoring Parameter or Func tion
Simplex
w ired
Duplex
Parameter
Hardw are
Cros s
Model
Range
Range
Chec k
Chec k
Primary Purpos e / Us e
En g i n e C o n tr o l Sen so r s
P0 (s tatic pres s ure)
X
X
X
X
P20 (intake total pres s ure)
X
X
X
X
X
X
X
X
Engine f uel s cheduling (trans ient, limiting, idle)
X
X
X
X
Engine EPR (thrus t) control
X
X
P30 (burner pres s ure)
X
P50 (ex haus t pres s ure)
Engine rating, bleed v alv e and V SV s c heduling
Engine rating and EPR (thrus t) control
T20 thermoc ouple (total temperature)
X
X
X
Engine rating, bleed v alv e and V SV s c heduling
TGT thermoc ouple (turbine temperature)
X
X
X
Engine temperature indic ation
X
X
Engine f uel s cheduling (s tarting, inc lement w eather)
X
Engine f uel / oil thermal control
X
Engine thrus t control, bleed v alv e & V SV s c heduling
X
T30 thermoc ouple (HPC ex it temperature)
Toil thermoc ouple (s c av enge oil temperature)
X
X
TRA res olv er (throttle angle s ens or)
X
X
LPC s peed
X
X
X
Engine N1 (thrus t) c ontrol and f uel f low limiting
IPC s peed
X
X
X
Bleed v alv e and V SV sc heduling and f uel f low
HPC speed
X
X
X
LPT s peed
X
X
limiting
Engine f uel s cheduling (trans ient, limiting, idle) and
bleed v alv e s c heduling
X
Engine turbine ov erspeed protec tion
En g i n e co n tr o l a ctu a to r s
X
FMU metering v alv e - pos ition res olv er
X
FMU HPSOV - s olenoid driv e
X
FMU HPSOV - pos ition mic ros w itc h
X
V SV control v alv e - torque motor driv e
X
V SV ac tuators - pos ition LV DT
X
A OHE (A MV ) - torque motor driv e
X
A OHE (A MV ) - pos ition LV DT
X
Handling bleed v alv e s olenoid driv e
Turbine impingement c ooling s olenoid driv e
Ttr70454.ppt
FMU metering v alv e - torque motor driv e
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X
X
Control of engine f uel f low
X
X
Control of engine f uel f low
X
X
Engine f uel s election/s hut-of f
X
Engine f uel s election/s hut-of f
X
X
X
Engine IPC V SV control
X
X
X
Engine IPC V SV control
Engine f uel / oil thermal control
X
X
Engine f uel / oil thermal control
X
X
X
IPC and HPC handling bleed c ontrol
X
X
X
Turbine impingement c ooling air c ontrol
SYSTEM CRITICALITY - SHEET 1
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System Criticality - continued
This page shows the second list of FADEC parameters
and their criticality.
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Full Authority Digital Engine Control
Sy s tem Conf iguration
Simplex
V alidation Tec hniques
Cros s Control / Monitoring Parameter or Func tion
Simplex
w ired
Duplex
Parameter
Hardw are
Cros s
Model
Range
Range
Chec k
Chec k
X
Engi ne sta r t contr ol
Starter air v alv e - s olenoid driv e
X
X
Control of engine s tart v alv e
Starter air v alv e - position mic ros w itc h
X
X
Control of engine s tart v alv e
Igniter s y stem - relay s
X
X
Energis ation of igniters f or engine s tarting
Engi ne oper a ti o n m oni to r s
Dedicated generator pow er s upply
X
X
A irc raf t 115 v olt pow er s upply
X
X
IPT c ooling air temperature (f ront)
X
X
IPT c ooling air temperature (rear)
X
X
Oil press ure trans mitter LV DT
X
X
X
X
HP oil f ilter dif f erential pres s ure sw itc h
Pow er s upply av ailability monitoring
Pow er s upply av ailability monitoring
X
IPT c ooling air ov erheat (internal f ire w arning)
X
IPT c ooling air ov erheat (internal f ire w arning)
X
Engine oil s upply health monitoring
X
Indic ation of impending HP oil f ilter by pas s
Sc av oil f ilter dif f erential pres sure s w itc h
X
X
Indic ation of impending sc av oil f ilter by pass
LP f uel f ilter dif f erential pres s ure s w itc h
X
X
Indic ation of impending LP f uel f ilter bloc kage
Zone 3 temperature
X
X
Indic ation of nacelle temperature
A nti-ic e high pres s ure s w itc h
X
X
Indic ation of high c ow l anti-ic e s y stem press ure
Low f uel pres s ure s w itc h
X
X
Indic ation of low f uel pres sure to HP pump
X
Engine f uel us eage monitoring / totalis ing
Engi ne condi ti on m o ni tor s
X
X
P160 (f an by pas s pres s ure)
X
X
X
X
Fan health monitoring
P25 (IPC ex it pres s ure)
X
X
X
IPC health monitoring
T25 (IPC ex it temperature)
X
X
Fuel f low meter
Ttr70455.ppt
Oil Quantity
Revision 6.0
X
X
IPC health monitoring
X
X
Indic ation of oil quantity
SYSTEM CRITICALITY - SHEET 2
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Full Authority Digital Engine Control
System Criticality - continued
This page shows the third list of FADEC parameters and
their criticality.
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Full Authority Digital Engine Control
Sy s tem Conf iguration
Simplex
V alidation Tec hniques
Cros s -
Parameter
Hardw are
Cros s
Model
Duplex
Range
Range
Chec k
Chec k
Door pos ition RV T
X
X
X
X
T/R door pos ition indic ation and thrus t limiting
Is olation v alv e s olenoid driv e
X
X
X
T/R IV energis ation f or deploy / s tow
Hy draulic pres s ure s w itc h
X
X
Hy draulic pres s ure monitoring
Ground s af ety s w itc h
X
DCV
X
Control / Monitoring Parameter or Func tion
Simplex
w ired
Th r ust r ever ser co ntr ol
X
Ground tes t operation of T/R
X
Direc tional c ontrol v alv e operation
Tertiary loc k s w itc hes
X
X
Tertiary loc k s tatus
Stow s w itc hes (1,2,3 & 4)
X
X
Door loc k s tatus
X
Thrus t rev ers er inhibit
Manual inhibit s w itc h
X
Ai r cr a ft g ener a to r m oni tor s
IDG oil inlet thermoc ouple
X
X
IDG outlet oil thermoc ouple
X
X
IDG f ilter dif f erential pres s ure s w itc h
X
A irc raf t s erv ic es ' generator oil s y s tem monitoring
A irc raf t s erv ic es ' generator oil s y s tem monitoring
X
A irc raf t s erv ic es ' generator oil s y s tem monitoring
A irc raf t interf ac e / c ommunic ations
Ai r cr a ft In ter fa ce
A RINC input bus s es
X
X
X
X
A RINC output bus s es
X
X
X
X
Hardw ired s peed indic ations
X
Igniter pow er
X
P2T2 heater pow er
A irc raf t interf ac e / c ommunic ations
V ibration monitoring trac king f ilters
X
Pow er s upply f or igniters
X
Mas ter lev er (hardw ired dis c rete)
X
Pow er s upply f or P2T2 heater
X
Starting
X
Starting
X
X
A utothrus t
dis c rete)
X
X
N1 rev ers ionary mode
X
A lternate s tart (hardw ired dis c rete)
X
A utothrus t engaged (hardw ired dis c rete)
A utothrus t ins tinc tiv e dis c onnec t (hardw ired
A utothrus t
Manual s elec tion of alternate (N1) c ontrol
X
X
Cabin air bleed s w itc hing (IP8/HP6)
Cabin air bleed v alv e - pos ition mic ros w itc h
X
X
Cabin air bleed s w itc hing (IP8/HP6)
Pres s ure regulating v alv e dis c rete
X
X
Status of airc raf t PRV
Engine running indic ation
X
Ttr70456.ppt
X
Cabin air bleed v alv e - s olenoid driv e
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Engine running indic ation to airf rame
SYSTEM CRITICALITY - SHEET 3
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Full Authority Digital Engine Control
Engine Control
Data Entry Plug (DEP)
The EEC has been designed to control all possible
configurations of the engine regardless of individual
characteristics.
To provide interchangeability of the unit, specific engine
information must be made available to the EEC. The
following information is contained within the DEP.
• Engine Serial Number
• Thrust Rating
• EPR/Thrust Trim Relationship
• Exhaust Gas Temperature Trim
• Engine Standard
• Intermix/Retrofit
• Fan Stall Index
• Idle Trim
The plug is a dual channel memory device which is
programmed with relevant engine data which is used by
the EEC to enable correct engine operation control.
Note: The Data plug remains with the engine throughout
its operational life not with the EEC
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Engine Control
DEP Programming Unit
To show the data entry plug programming unit.
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Engine Control
DEP Programming Unit
To show data entry plug programming unit printouts.
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Rotor Integrity Protection
Overspeed Protection Unit (OPU)
The OPU is an independent system to the EEC
It is designed to prevent severe LP and IP shaft overspeed
in the event of severe mal-scheduling VSV's and/or fuel
upward runaway due to fuel metering valve failing open.
In the event of overspeed the unit indirectly signals the
pressure raising and shut-off valve to close.
The unit is physically, electrically and functionally separate
from the EEC It is bolted to the back of the PCU and
mounted alongside the EEC unit on the left hand side of
the fan case inside the suitcase.
The unit is twin channel with specific logic circuit cards
and a control card, using digital technology.
The OPU interfaces with the LP and IP shaft speed
probes, electrical supply from the dedicated alternator two
separate single phase circuits and with the EEC to transfer
BITE data, and to enable EEC selection of any 2 of 3 input
probe signals for N1 and N2 EEC speed reference.
In an overspeed shut-down B.I.T.E. sequence one logic
circuit card takes the role of master and the other of slave
- the generation of matched signals between master and
slave allows the shut-down sequence to proceed.
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Overspeed Protection Unit (OPU)
Purpose of diagram
To show location and construction of OPU
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Turbine Overspeed Protection System
In the unlikely event of a broken shaft, the turbine is
relieved of its load of driving the compressor. This would
cause the turbine to overspeed and in the worst case
some of the blades may be released, unless sufficient
energy is removed from the core engine gas flow.
LP Shaft Breakage
Three LP compressor speed probes send signals of shaft
speed to the overspeed protection unit (OPU). This unit
makes a selection of two satisfactory N1 signals and
transmits them to the Electronic Engine Controller (EEC).
One N1 signal is supplied to each of the logic lanes 'A' and
'B' on the turbine overspeed circuit board in the EEC lane
'A'.
Three LP turbine speed probes send signals directly to the
LP Turbine Overspeed (L.P.T.O.S.) circuit board of the
EEC lane 'A'.
Full Authority Digital Engine Control
Each logic lane compares the LP compressor speed with
its L.P. turbine speed. If the two logic lanes detect a speed
difference between L.P. compressor and L.P. turbine in a
specified time limit it is accepted as a true failure
condition. If an LP shaft failure is accepted as true the
system will signal a closure of the Pressure Raising and
Shut-off Valve (P.R.S.O.V.). The engine is immediately
shut down.
Once the fuel flow has been shut off the P.R.S.O.V. is
latched in the fuel off position. If inadvertent shut down
occurs the pilot has a reset facility in the flight deck. If
there is a failure of a compressor speed signal, which
shows that overspeed of the turbine is not possible, the
related overspeed protection circuits are disarmed.
If one logic lane 'A' or 'B' becomes defective the turbine
overspeed circuit is disarmed. This prevents incorrect
operation of the system.
Each logic lane is supplied with one N1 signal. If one of
these signals is not satisfactory then the applicable logic
lane makes the selection of the alternative signal. When
the L.P. rotor system reaches a speed higher than 1000
R.P.M. the turbine overspeed protection system is armed.
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Turbine Overspeed Protection System
Built In Test Equipment (B.I.T.E.)
The EEC does a test of the LP turbine overspeed function
during every ground start - pre light up. The B.I.T.E.
provides the necessary turbine speed difference to the
turbine overspeed protection circuits to momentarily shut
off the fuel during start sequences. Movement of the
P.R.S.O.V. in the F.M.U. to the closed position is
monitored by the EEC Almost immediately the EEC
cancels the B.I.T.E. test signal to cause the P.R.S.O.V. to
become open again. Therefore the engine start sequence
is not stopped by the test. Defects found during the
B.I.T.E. test are stored in the EEC The defects are
subsequently transmitted to the Central Maintenance
Computer (C.M.C.).
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P20/T20 Probe
The P20/T20 probe is mounted inside the air intake cowl
at 15o to right of top dead centre when viewed from rear.
The probe measures both engine intake pressure and
temperature.
Temperature is measured by two independent platinum
resistance elements. A small amount of air passes over
the elements, whilst the rest of the air passes straight
through the probe.
The pressure signal offtake is just above where the main
airstream flows through the probe. A pipe passes through
the body to the pressure connector on the base plate.
An electrical de-icing heater element is configured around
the probe powered by aircraft 115V 400HZ supply.
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Maintenance Practices
Full Authority Digital Engine Control
.
Connect air tube connector and torque load to 240
lbf/in (2,71 mdaN) and safety with OMat 238
lockwire.
.
Connect the electrical connectors IAW standard
practices 70-50-02.
.
Refit access panel and torque load screws to 135
lbf/in (1,52 mdaN).
P20/T20 Probe Removal/Installation
.
Get access to the P20/T20 probe access panel on
top of the air intake cowl.
.
Remove access panel.
.
Disconnect air tube connector and fit blanking caps.
.
Disconnect electrical connectors and fit blanking
caps.
.
Disconnect the bonding strap.
.
Remove the outer row of nuts.
WARNING
If you remove the inner row of nuts the probe will fall into
the engine air intake. This can cause injury and/or
damage.
.
Remove the probe assembly and mounting plate.
Installation of Replacement Probe:
.
Remove the mounting plate and packer from the
removed probe and fit to replacement probe.
.
Torque load the nuts to 40 lbf/in (0,45 mdaN) and
safety with OMat 238 lockwire.
.
Attach the mounting plate to the stud ring assembly
and torque load the nuts to 100 lbf/in (1,13 mdaN).
.
Clean mating faces of bonding strap and
attachment point and fit bonding strip. Torque load
nuts to 55 lbf/in (0,62 mdaN).
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Dedicated Alternator Stator
Removal/Installation
Removal Procedure
CAUTION
Be careful when you remove the alternator stator. The
magnets on the alternator rotor will apply a magnetic pull
to the stator. Damage to the stator and the rotor can occur
if you do not hold the stator carefully.
.
Disconnect electrical connectors and fit blanks.
.
Hold the alternator stator and remove the bolts.
.
Carefully pull the alternator stator forward in a
straight line until clear of alternator rotor.
.
Remove and discard the seal ring and fit blanking
cover to stator.
Installation Procedure
.
Fit new seal ring and install alternator stator.
.
Torque load the bolts to 100 lbf/in. (1,13 mdaN).
.
Make sure the electrical connectors are clean
before connecting.
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Data Entry Plug
Removal/Installation
Removal procedure:
.
Put 1 3/32 inch allen key into the end of the Hi-Lok
pin shank.
.
Hold the Hi-Lok pin with the allen key and remove
the Hi-Lok collar with pliers and remove the washer.
.
Remove Hi-Lok pin to release lanyard.
.
Discard the Hi-Lok pin and the collar.
.
Disconnect the data entry plug.
.
Put cap on the data entry plug and EEC receptacle.
Carry out FADEC system test (MM Task
73-21-00-740-801) and EEC configuration test (MM Task
73-21-00-710-802).
Installation Procedure:
Note Check the data on the DEP identification label
agrees with the engine records and engine data plate.
.
Make sure all connections are clean .
.
Connect DEP.
.
Put lanyard on the new Hi-Lok pin and put Hi-Lok
pin in its position on the electronics protection box.
.
Put washer and Hi-Lok collar on the Hi-Lok pin and
tighten with fingers.
.
Using 7/16 inch allen key in the end of Hi-Lok pin,
turn the Hi-Lok collar in clockwise direction until the
hexagonal end breaks off.
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End of section
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