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the APU protective shutdowns showing the ones that are inhibited in-flight (essential mode):
* - An asterik indicates these shutdowns are inhibited in the essential mode.
28 VDC Power Loss Fire Indication*
Reverse Flow* No Acceleration (during start)*
Loss of EGT Signal* Overspeed
Low Oil Pressure* Emergency Fire
High Oil Temperature* Internal ECU Failure
Overtemperature* Loss of Overspeed Protection Circuit
No Flame (during start)*
1
AUXILIARY POWER UNIT
May 06/2005 Flight Crew Operating Manual Volume 2
REV 2 CSP 100-6 05-01-07
START AND IGNITION
DESCRIPTION
The starting and ignition systems operate automatically during the APU start sequence.
COMPONENTS AND OPERATION
STARTER
The APU start system consists of a DC starter and a series of relays controlled by the electronic control unit (ECU). The
electrical power source is:
- ECU — battery bus
- Starter — APU battery direct bus
To start the APU, the APU start contactor is closed and the right battery is connected to the APU starter.
The right battery contactor is opened to isolate the right battery from the right essential bus during APU start to prevent a
voltage drop on the essential busses. The left battery maintains power to the left and right essential busses during APU start
It is important to delay starting the APU for five minutes after application of anti-icing fluid near the APU area. This allows
any anti-icing fluid inadvertently directed into the APU inlet to drain out.
Three consecutive APU start attempts may be made at 1 minute intervals, followed by a 20 minute wait. Two more start
attempts may be made. No further attempts may be made for a period of at least 40 minutes.
Six consecutive and successful APU starts may be made at 10 minute intervals with a one-hour waiting period prior
to additional start attempts.
IGNITION SYSTEM
The APU ignition system consists of an ignition unit, igniter plug lead and a single igniter plug. The ignition system is
energized during the APU start cycle and is de-energized when the APU start sequence is complete. To engage the APU,
pull up on the switch, position the switch to RUN for approximately 3 seconds until the APU indications are displayed on
the left MFD. Then move the switch to START and hold until the START icon is displayed on the left MFD, the APU will
then begin the start sequence.
The APU ignition system consists of a single output ignition unit. Should a flameout occur during operation, the ignition
unit automatically fires the igniter through the auto relight function of the ECU. An immediate restart (APU starter re engagement)
can be initiated upon APU rolldown below 10% rpm.
SYSTEM OPERATION
The ECU uses 28 vdc power from the right aircraft battery bus to energize the ignition system during the ignition/start sequence.
During the start sequence, the R BATT switch is illuminated OFF on the ELECTRICAL panel and the flow line
from R BATT to the R ESS BUS on the Electrical synoptic page is depicted in white. The ignition system is energized at
5% APU rpm and stays energized until 95% rpm.
REV 2
AUXILIARY POWER UNIT
Volume 2 Flight Crew Operating Manual Sep 13/2004
05-01-08 CSP 100-6 REV 1
PNEUMATIC SUPPLY
DESCRIPTION
APU compressor bleed air is used to pressurize the right side of the bleed air manifold to provide pneumatics for engine
starting and air conditioning. Reverse airflow protection is provided by a one way check valve.
COMPONENTS AND OPERATION
The APU pneumatic control system consists of a bleed valve and a surge control valve.
BLEED VALVE MODULATION/BLEED AIR LOADS
The pneumatic load on the APU is controlled by the ECU as a function of demands sent from the APU bleed valve. When
the bleed valve is opened, a portion of the compressor discharge air is diverted from the plenum chamber to the right side
of the bleed air manifold. The bleed valve opens when the flight deck APU bleed load switch is placed in the open position.
The APU engine reaction to bleed loads is different from shaft loading. Bleed air extraction from the APU diverts pneumatic
energy from the turbine and some of the cooling air from the combustion system. As with the shaft load, the rpm
drops, and with the corresponding ECU command fuel flow increase, recovers to full governed speed.
Air diverted into the aircraft bleed manifold is no longer available to drive the turbine or cool the combustion system. The
loss of pneumatic energy is replaced with thermal energy (fuel) with significant temperature increase.
The following variables influence exhaust gas temperature (EGT) under bleed air load conditions:
- Aircraft demand for bleed air
- Ambient air temperature and density
　
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