levers also control propeller reverse pitch. Distinct
movement (pulling up and then aft on the power
lever) by the pilot is required for reverse thrust.
Placarding beside the lever travel slots reads
POWER. Upper lever travel range is designated
INCR (increase), supplemented by an arrow point-
ing forward. Lower travel range is marked IDLE,
LIFT and REVERSE. A placard below the lever
slots reads: CAUTION REVERSE ONLY WITH
2-23. CONDITION LEVERS.
Two condition levers are located on the control
pedestal (fig. 2-7). Each lever starts and stops the
fuel supply, and controls the idle speed for its
engine. The levers have three placarded positions:
FUEL CUTOFF, LO IDLE, and HIGH IDLE. In
the FUEL CUTOFF position, the condition lever
controls the cutoff function of its engine-mounted
fuel control unit. From LO IDLE to HIGH IDLE,
they control the governors of the fuel control units
to establish minimum fuel flow levels. LO IDLE
position sets the fuel flow rate to attain 52 to 55%
(at sea level) minimum N1 and HIGH IDLE posi-
tion sets the rate to attain 70% minimum N1. The
power lever for the corresponding engine can select
N1 from the respective idle setting to maximum
power. An increase in low idle N1 will be experi-
enced at high field elevation.
2-24. FRICTION LOCK KNOBS.
Four friction lock knobs (fig. 2-7) are located on
the control pedestal to adjust friction drag. One
knob is below the propeller levers, one below the
condition levers, and two under the power levers.
When a knob is rotated clockwise, friction restraint
is increased opposing movement of the affected
lever as set by the pilot. Counterclockwise rotation
of a knob will decrease friction drag thus permitting
free and easy lever movement. Two FRICTION
LOCK placards are located on the pedestal adjacent
to the knobs.
2-25. ENGINE FIRE DETECTION SYSTEM.
a. Description. A flame surveillance system is
installed on each engine to detect external engine
fire and provide alarm to the pilot. Both nacelles are
monitored, each having a control amplifier and
three detectors. Electrical wiring connects all sensors
and control amplifiers to DC power and to the cock-
pit visual alarm units. In each nacelle, one detector
monitors the forward nacelle, a second monitors the
upper accessory area, and a third the lower accessory
area. Fire emits an infrared radiation that will be
sensed by the detector which monitors the area of
origin. Radiation exposure activates the relay circuit
of a control amplifier which causes signal power to
be sent to cockpit warning systems. An activated
surveillance system will return to the standby state
after the fire is out. The system includes a functional
test switch and has circuit protection through the
FIRE DETR circuit breaker. Warning of internal
nacelle fire is provided as follows: the red MASTER
WARNING lights on the glareshield illuminate
accompanied by the illumination of a red warning
light in the appropriate fire control T-handle plac-
arded No. 1 FIRE PULL or No. 2 FIRE PULL (fig.
2-29). Fire detector circuits are protected by a single
5-ampere circuit breaker, placarded FIRE DETR,
located on the overhead circuit breaker panel (fig.
b. Fire Detection System Test Switch. O n e
rotary switch placarded FIRE PROTECTION TEST
on the copilots subpanel is provided to test the
engine fire detection system. Before checkout, bat-
tery power must be on and the FIRE DETR circuit
breaker must be in. Switch position DETR 1, checks
the area forward of the air intake of each nacelle,
including circuits to the cockpit alarm and indica-
tion devices. Switch position DETR 2, checks the
circuits for the upper accessory compartment of
each nacelle. Switch position DETR 3, checks the
circuits for the lower accessory compartment of each
nacelle. Each numbered switch position will initiate
the cockpit indications previously described.
c. Erroneous Fire Detection System Indica-
tions. During ground test of the engine fire detection
system, an erroneous indication of system fault may
be encountered if an engine cowling is not closed
properly, or if the aircraft is headed toward a strong
external light source. In this circumstance, change
the aircraft heading to enable a valid system check.
2-26. ENGINE FIRE EXTINGUISHER SYSTEM.
a. Description. The fire extinguisher system
utilizes an explosive squib, connected to a valve
which, when opened, allows the distribution of the
pressurized extinguishing agent through a plumbing
network of spray nozzles strategically located in the
fire zones of the engines.
b. Fire Pull Handles . The tire control
T-handles, which are used to arm the extinguisher
system are centrally located on the pilots instru-
ment panel (fig. 2-29), immediately below the
glareshield. These controls receive power from the
hot battery bus. The fire detection system will indi-
cate an engine fire by illuminating the master fault
warning light on the pilots and copilots glareshield
and the respective No. 1 FIRE PULL or No. 2 FIRE