TM 1-1510-224-10
switch closed for fire alarm continuity test functions. As
the temperature around the sensing cable increases, the
gases within the tube to expand. When the pressure
from the expanding gases reaches a preset point, the
contacts of the responder alarm switch close, activating
the respective fire warning system.
b. Warning System. The fire warning system
consists of two lenses, placarded #1 FIRE PULL and #2
FIRE PULL, located in the T handles below the
glareshield, two
MASTER WARNING annunciators
located in opposite sides of the glareshield, and a
responder
unit
with
a
sensor
in
each
engine
compartment. If the sensor develops a leak, the loss of
gas pressure allows the integrity switch to open and
signal a lack of sensor integrity.
c. Testing. Testing system integrity, availability
of power, and the proper operation of the annunciators
(#1 and #2 FIRE PULL and MASTER WARNING) is
accomplished by two switches located on the copilot's
subpanel. The switches are placarded ENG FIRE
TEST, DET - OFF - EXT, LEFT, RIGHT. When either
the LEFT or RIGHT switch is placed in the DET position,
electrical current flows from a 5-ampere circuit breaker,
placarded FIRE DETR, located on the overhead circuit
breaker panel, through the engine fire detector circuitry
to the integrity switch contacts in the respective
responder unit, causing the respective annunciators to
illuminate. If the circuit breaker is open, the system will
not operate during a test, or activate the annunciators if
the detector cable senses an overtemperature condition.
The system may be tested either before, after, or during
flight as desired.
2-26. ENGINE FIRE EXTINGUISHER SYSTEM.
a. Description. The engine fire extinguisher
system (fig. 2-17) consists of a supply cylinder, an
explosive squib, and a valve located in each of the main
gear wheel wells. A gage calibrated in PSI is provided
on each supply cylinder for determining the level of
charge. The extinguishing agent charge level should be
checked during each preflight. When fired, the explosive
squib opens the valve, releasing all of the pressurized
extinguishing agent into a plumbing network. The
plumbing
network
terminates
in
spray
nozzles,
strategically located in the probable fire areas of the
engine compartment, which distribute the extinguishing
agent.
b. Operation. Fire control T handles used to
arm the extinguisher system are centrally located on the
instrument panel, immediately below the glareshield (fig.
2-18). These controls receive power from the hot battery
bus. The fire detector system will indicate an engine fire
by illuminating the MASTER WARNING annunciators on
the glareshield and the respective #1 or #2 FIRE PULL
annunciators in the fire control T handles. Pulling the
fire control T handle will electrically arm the extinguisher
system and close the firewall shutoff valve for that
particular engine. This will cause the annunciator in the
PUSH TO EXTINGUISH switch and the respective #1 or
#2
FUEL
PRESS
annunciator
on
the
warning
annunciator panel to illuminate. Pressing the lens of the
PUSH TO EXTINGUISH fire switch will fire the squib,
expelling all the agent in the cylinder at one time. A
hinged plastic guard covers the PUSH TO EXTINGUISH
fire switch to prevent inadvertently actuating the fire
extinguish switch squib circuit. The respective caution
annunciator, #1 and
#2 EXTGH DISCH on the
caution/advisory annunciator panel and the MASTER
CAUTION annunciator on the glareshield will illuminate
and remain illuminated, regardless of the master switch
position, until the squib is replaced.
c. Testing. The test switches, located on the
copilot's subpanel (fig. 2-8), are placarded ENG FIRE
TEST, DET - OFF - EXT, LEFT and RIGHT, and provide
a test of the fire detection and extinguisher circuitry.
When either of the switches is placed in the EXT
position, the corresponding PUSH TO EXTINGUISH,
SQUIB OK, and EXTGH DISCH annunciators should
illuminate. The system may be tested either before,
after, or during flight as desired.
A gage, calibrated in PSI, is mounted on each
extinguishing agent supply cylinder. The gage indicates
the level of charge and should be checked during
preflight (table 2-1).
2-27. OIL SUPPLY SYSTEM.
CAUTION
Maximum allowable oil consumption
is one quart in 5 hours of engine
operation.
a. The engine oil tank is integral with the air-
inlet casting located forward of the accessory gearbox.
Oil for
Table 2-1. Engine Fire Extinguisher Gage Pressure
TEMP °C
-40
-29
-18
-06
04
16
27
38
48
190
220
250
290
340
390
455
525
605
PSI
to
to
to
to
to
to
to
to
to
240
275
315
365
420
480
550
635
730
BT2D194
2-37
