The heating elements protect the
stall warning lift transducer vane
and face plate from ice, however, a
buildup of ice on the wing may
change or disrupt the airflow and
prevent the system from accurately
indicating an imminent stall.
Stall Warning Heat. The lift transducer is
equipped with anti-icing capability on both the mounting
plate and the vane. The heat is controlled by a switch
located on the overhead control panel placarded STALL
WARN. The level of heat is minimal for ground operation
but is automatically increased for flight operation through
the left landing gear saety switch. Circuit protection is
provided by a 15-ampere circuit breaker, placarded
STALL WARN, on the overhead circuit breaker panel (fig.
2-56. PITOT AND STATIC SYSTEM.
Description. The pitot and static system supplies
static pressure to two airspeed indicators, two altimeters,
two vertical velocity indicators, and ram air to the
airspeed indicators. This system consists of two pitot
masts (one located on each side of the lower position of
the nose), static air pressure ports in the aircraft's exterior
skin on each side of the aft fuselage, and associated
system plumbing. The pitot head is protected from ice
formation by internal electric heating elements.
Alternate Static Air Source. An alternate static air
line, which terminates just aft of the rear pressure
bulkhead, provides a source of static air for the pilot's
instruments in the event of source failure from the pilot's
static air line. A control on the pilot's subpanel placarded
PILOTS STATIC AIR SOURCE, may be actuated to
select either the NORMAL or ALTERNATE air source by
a two position selector valve. The valve is secured in the
NORMAL position by a spring clip. Refer to Chapter 7 for
airspeed indicator and altimeter calibration information
when using the alternate air source.
Stall Warning System. The stall warning system
consists of a transducer, a lift computer, a warning horn,
and a test switch. Angle of attack is sensed by
aerodynamic pressure on the lift transducer vane located
on the left wing leading edge. When a stall is imminent,
the output of the transducer activates a stall warning horn.
The system has preflight test capability through the use of
a switch placarded STALL WARN TEST OFFLDG GEAR
WARN TEST on the right subpanel. Holding this switch in
the STALL WARN TEST position actuates the warning
horn by moving the transducer vane. The circuit is
protected by a 5-ampere circuit breaker, placarded STALL
WARN, on the overhead circuit breaker panel.
2-57. BRAKE DEICE SYSTEM.
Description. A heated-air brake deice system
may be used on the ground or in flight with gear retracted
or extended. When activated, hot air is diffused by
means of a manifold assembly over the brake discs in
each wheel. Manual and automatic controls are provided.
There are two primary occasions which require brake
deicing. The first is when an aircraft has been parked in a
freezing atmosphere allowing the brake systems to
become contaminated by freezing rain, snow, or ice, and
the aircraft must be moved or taxied. The second
occasion is during flight through icing conditions with wet
brake assemblies presumed to be frozen, which must be
thawed prior to landing to avoid possible tire damage and
loss of directional control. Hot air for the brake deice
system comes from the compressor stage of both engines
obtained by means of a solenoid valve attached to the
bleed air system which serves both the surface deice
system and the pneumatic systems operation.
Operation. A switch on the overhead control
panel, placarded BRAKE DEICE, controls the solenoid
valve by routing power through a control module box
under the aisleway floorboards. When the switch is on,
power from a 5-ampere circuit breaker on the overhead
circuit breaker panel is applied to the control module. A
10-minute timer limits operation and avoids excessive
wheel well temperatures when the landing gear is
retracted. The control module also contains a circuit to
the green BRAKE DEICE ON annunciator light, and has a
resetting circuit interlocked with the gear uplock switch.
When the system is activated, the BRAKE DEICE ON
light should be monitored and the control switch selected
OFF after the light extinguishes otherwise, on the next
gear extension the system will restart without pilot action.
The control switch should also be selected OFF, if deice
operation fails to self-terminate after 10 minutes. If the
automatic timer has terminated brake deicer operation
after the last retraction of the landing gear, the landing
gear must be extended in order to obtain further operation
of the system.
momentarily illuminate during simultaneous operation of
the surface deice and brake deice systems at low N1
speeds. If the lights immediately extinguish. this may be