Never cycle the system rapidly, this
may cause the ice to accumulate
outside the contour of the inflated
boots and prevent ice removal.
2-52. FORWARD GUARDRAIL DUAL DATA LINK
ANTENNA RADOME ANTI-ICE.
The forward guardrail dual data link (GDDL) antenna
radome anti-ice system utilizes engine bleed air to
prevent the formation of ice on the radome. The system
is controlled by a switch, placarded RADOME - ON,
located on the overhead control panel (fig. 2-15). The
circuit is protected by a 7.5-ampere circuit breaker,
placarded RADOME ANTI-ICE, located on the overhead
circuit breaker panel (fig. 2-9).
2-53. PROPELLER ELECTRIC DEICE SYSTEM.
Description. The propeller electric deice system
includes electrically heated deice boots, slip rings and
brush block assemblies, a timer for automatic operation,
ammeter, circuit breakers for left and right propeller and
control circuit protection, and two switches located on
the overhead control panel (fig. 2-15), for automatic or
manual control of the system.
Automatic Operation. The two position switch
located on the overhead control panel, placarded PROP
AUTO - ON, is provided to activate the automatic
system. When the switch is placed to the ON position,
the timer diverts power through the brush block and slip
Subsequently, the timer then diverts power to all heating
elements on the other propeller for the same length of
time. This cycle will continue as long as the switch is in
the ON position. The system utilizes a metal foil type
single heating element, energized by DC voltage. The
timer switches every 90 seconds, resulting in a complete
cycle in approximately 3 minutes.
Manual Operation. The manual propeller deice
system is provided as a backup to the automatic system.
overhead control panel, placarded PROP - MANUAL -
ON, controls the manual override relay. When the
switch is held in the ON position, the automatic timer is
overridden, and power is supplied to the heating
elements of both propellers simultaneously. The switch
is of the momentary type and must be held in position for
approximately 90 seconds to dislodge ice from the
propeller surface. Repeat this procedure as required to
avoid significant buildup of ice, which will result in a loss
of performance, vibration, and impingement of ice upon
the fuselage. The propeller deice ammeter will not
indicate a load while the propeller deice system is being
utilized in the manual mode. However, each aircraft
loadmeter will indicate an approximate 10% increase in
load while the manual propeller deice system is
2-54. PITOT HEAT SYSTEM.
Pitot heat should not be used for
more than 15 minutes while the
Overheating may damage the heating
Heating elements are installed in both pitot masts,
located on the nose. Each heating element is controlled
by an individual switch, placarded PITOT, LEFT, RIGHT
and ON, located on the overhead control panel
(fig. 2-15). Circuit protection is provided by the two 7.5-
ampere circuit breakers, placarded PITOT HEAT, on the
overhead circuit breaker panel (fig. 2-9). The true
airspeed temperature probe heat control circuit is also
protected by these circuit breakers. If either left or right
pitot heat is on, the true airspeed temperature probe
heat will be on.
2-55. STALL WARNING HEAT SYSTEM.
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
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 landing gear
safety switch. Circuit protection is provided by a 15-
ampere circuit breaker, placarded STALL WARN, on the
overhead circuit breaker panel (fig. 2-9).
2-56. STALL WARNING SYSTEM.
The stall warning system consists of a transducer, a
lift computer, 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
(fig. 2-1 and 2-2). 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 OFF LDG GEAR
WARN TEST on the copilot's subpanel (fig. 2-8).