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the corresponding NO. 1 or 2 ENG CHIP DET caution
capsule will illuminate. Also, the associated ENGINE
CHIP DETECTOR or ENGINE TRANSMISSION CHIP
DETECTOR magnetic indicator on the MAINTENANCE
PANEL (fig. 2-9-2) will latch. Refer to Chapter 9 for emergency
procedures.
2-3-26. Engine Chip Detector Fuzz Burn-Off.
Helicopters equipped with the chip detector fuzz burn -off
system in the engine are identified by a module labeled
PWR MDL CHIP BURN-OFF located below the MAINTENANCE
PANEL. The chip detector fuzz burn-off system
employs an automatically operated fuzz burn-off
electrical circuit with the ability to eliminate nuisance chip
lights caused by minute ferrous metallic fuzz or ferrous
metallic particles on the engine accessory gear box
(AGB) chip detectors. The response time of the fuzz
burn-off circuit is more rapid than that of the helicopter
warning system; thus a successful fuzz burn-off will be
accomplished before any caution capsule on the master
caution panel illuminates. Should the particle or particles
not burn-off, the NO. 1 or NO. 2 ENG CHIP DET caution
capsule will illuminate. Also, the corresponding ENGINE
CHIP DETECTOR or ENGINE TRANSMISSION CHIP
DETECTOR magnetic indicator on the MAINTENANCE
PANEL will latch. Power for the PWR MDL CHIP BURNOFF
is supplied by the No. 1 DC bus through the HYDRAULICS
MAINT PNL circuit breaker on the No. 1 PDP.
2-3-27. Engine Interstage Air Bleed.
NOTE
Bleed band oscillations at low torque settings
(approximately 30% torque per engine), indicated
by fluctuating RRPM and torque, can
occur and are not cause for engine rejection.
To aid compressor rotor acceleration and prevent compressor
stall, an interstage air bleed system is provided
on each engine. A series of vent holes through the compressor
housing at the sixth stage vane area allows pressurized
air to bleed from the compressor area. This enables
the compressor rotor to quickly attain preselected
RPM. The pneumatic interstage air bleed actuator controls
operation of the air bleed by tightening or loosening
a metal band over the vent holes. Should the bleed band
malfunction and remain open, there would be a noticeable
loss in power. 712 The interstage air bleed system
operates automatically when the ENG COND levers or
the engine beep trim switches are used to govern RPM.
714A The interstage air bleed system operates automatically
through the FADEC system.
2-3-28. Engine Drain Valves.
Pressure-operated engine drain valves are in the bottom
of each engine combustion housing. The valves automatically
drain unburned fuel from the combustion chamber
following an aborted start or whenever the engine is
shut down. One valve is at the forward end of the comTM
1-1520-240-10
2-3-7
bustion camber and the other is at the aft end to ensure
complete drainage.
2-3-29. FADEC Description.
714A Each engine is controlled by its own Full Authority
Digital Electronic control System (FADEC) which provides
the following features:
a. Automatic start scheduling.
b. 1 and 2 engine load sharing.
c. Power turbine speed governing.
d. Transient load anticipation (using rotor speed
and collective pitch rates).
e. Transient torque smoothing (using N2 rates).
f. Contingency power capability to meet aircraft demand.
g. Acceleration and deceleration control.
h. Engine temperature limiting throughout the operating
range.
i. Surge avoidance.
j. Compressor bleed band scheduling..
k. Fuel flow limiting.
l. Engine fail detection.
m. Power assurance test.
n. Engine history/fault recording.
o. Engine-to-engine communication (via data
bus).
p. Automatic switchover to reversionary backup in
the event of a FADEC primary system failure.
The FADEC provides automatic engine start, simultaneously
sequencing ignition, start fuel, and stabilized
operation at idle. A data link between 1 and 2 engine
FADEC systems transmits signals to achieve load sharing.
It also provides control of N1 speed and NR (N2)
output shaft speed to maintain the rotor system at a near
constant RRPM throughout all flight power demand conditions.
FADEC provides smooth acceleration and overtemperature
protection when ECLs (both together) are
moved from GROUND to FLIGHT. Overtemperature
protection is provided (through the DECU temperature
limiting function) by control system thermocouple interface
at the power turbine inlet. The control system
compares PTIT temperature signals with reference limits
to calculate and provide appropriate N1 acceleration.
　
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