HEADQUARTERS OH-58A, Техническое руководство, Страница: 68 / 292

TM 55-1520-228-10
2-24 Change 11
2-14. INFRARED SUPPRESSION SYS-
TEM.
The infrared suppression system is specially formulated
to reduce detection of IR-seeking missiles. The system
consists of two exhaust stacks located at the top of the
engine cowling and two shields located on the side en-
gine cowling (figure 2-2). The exhaust stacks have cool-
ing fins and a shelf around them to reduce the infrared
signature of the hot engine exhaust. The shields on the
side engine cowling prevent detection of the hot engine
area from the sides.
2-15. TEMPERATURE MEASUREMENT
SYSTEM.
The temperature measurement system consists of four
chromel-alumel single junction thermocouples in the
gas producer turbine outlet and an associated integral
harness. The voltages of the four thermocouples are
electrically averaged in the assembly and delivered by
the assembly lead to the airframe terminal block for
reference to the engine temperature indicating system.
2-16. COMPRESSOR BLEED AIR SYS-
TEM.
The 5th stage compressor bleed valve permits rapid
engine response. The system consists of a compressor
discharge pressure sensing port on the scroll, tubing
from the sensing port to the bleed valve, a compressor
bleed control valve, and a bleed air manifold on the
compressor case. Elongated slots between every other
vane in the compressor 5th stage bleeds compressor air
into a manifold which is an integral part of the compres-
sor case. The manifold forms the mounting flange for the
compressor bleed control valve when the compressor
case halves are assembled. Compressor discharge air
pressure sensing for bleed control valve operation is
obtained at a sensing port of the compressor scroll. The
bleed control valve is normally open. It is closed by
compressor discharge pressure.
2-17. ENGINE OIL SUPPLY SYSTEM.
The lubricating system is a dry sump type with an exter-
nal reservoir and heat exchanger. A gear type pressure
and scavenge pump assembly is mounted within the
power and accessory gearbox. The oil filter, filter bypass
valve, and pressure regulating valve are in a unit which
is located in the upper right side of the power and acces-
sory gearbox housing and are accessible from the top
of the engine. The oil tank is mounted aft of the engine
rear firewall on top of the intermediate cabin section. A
check valve is located between the housing and the filter
unit. Probe type magnetic chip detectors are installed at
the bottom of the power accessory gearbox and the
engine oil outlet connection. All engine oil system lines
and connections are internal with the exception of pres-
sure and scavenge lines to the front compressor sup-
port, the gas producer turbine support, and the power
turbine support. The oil cooler blower is an integral part
of the tail rotor drive and is located aft of the freewheel-
ing unit and adjacent to the oil tank.
2-17.1
.
ENGINE OIL SUPPLY SYSTEM -
EXTERNAL SCAVENGE OIL FILTER
(After
Compliance with MWO 55-1520-228-50-44).
The in-line external scavenge oil filter is located above
the tail rotor drive shaft, behind the engine oil tank. The
filter assembly consists of a filter element, oil bypass
indicator, and bypass valve. When the filter element
becomes clogged, it will give a warning by extending the
oil bypass (red) indicator. The indicator extends when a
set differential pressure across the filter is exceeded.
When in the reset position, the indicator will be hidden
from view. If the indicator is extended, it can be reset by
pressing in. An extended indicator is no sufficient reason
to ground the helicopter. If filter bypass indicator (red
button) is showing, reset indicator, ground run engine,
and reinspect. If indicator red button is not showing after
ground run, aircraft may be released for operation. If
indicator red button is showing, do not fly aircraft, notify
maintenance.
2-18. IGNITION SYSTEM.
a. The engine ignition system consists of a keylock
ignition switch, a low tension capacitor discharge igni-
tion exciter, a spark igniter lead, and a shunted surface
gap spark igniter. The system derives its input power
from the helicopter 28-volt DC electrical system.
b. The keylock ignition switch (figure 2-13 and fig-
ure 2-14) locks out the starter system and prevents un-
authorized use of the helicopter, thereby preventing
possible injury to personnel and/or damage to the equip-
ment.
2-19. STARTERS SWITCH.
The starter switch (figure 2-3 and figure 2-4), located in
the collective stick switch box, is a push button type
switch. When the switch is pressed, the circuit to the
starter relay actuating coil, the igniter unit, and the fuel
boost pump are energized. The switch is released when
the engine start cycle is completed or abort start proce-
dures is completed. The keylock ignition switch must be
ON to complete the ignition circuit. The circuit is pro-
tected by the START ENG, IGN ENG and FUEL PUMP
circuit breakers.