SOLENTA AVIATION
AIRCRAFT TECHNICAL NOTES
CESSNA 208B
GRAND CARAVAN
TABLE OF CONTENTS
Cessna C208B Grand Caravan
General Airplane
Cockpit And Cabin
Warning Systems
Turbine Engine Theory
PT6A-114A Powerplant
Propeller
Fuel System
Electrical System
Air-conditioning And Ventilation
Oxygen
Ice And Rain Protection
Instruments
Avionics
Limitations
Section 1
Section 2
Section 3
Section 4
Section 5
Section 6
Section 7
Section 8
Section 9
Section 10
Section 11
Section 12
Section 13
Section 14
© 2005 Solenta Aviation (Pty) Ltd- All Rights Reserved
Issued 2006
i
GENERAL AIRPLANE
Table of Contents
Introduction ................................................. 1-1
Fuselage...................................................... 1-2
Wings .......................................................... 1-2
Empennage................................................. 1-3
Cargo Pod ................................................... 1-3
Dimensions ................................................. 1-4
Landing Gear ...............................................1-5
Brake System...............................................1-5
Minimum Turning Distance ..........................1-6
Flight Controls..............................................1-7
Control Locks ...............................................1-8
Wing Flap System........................................1-9
Introduction
The Cessna C208B Grand Caravan is an un-pressurised, all-metal, high wing, single-engine
aeroplane equipped with a fixed tricycle landing gear. A composite cargo pod is optional equipment
Figure 1-1 Cessna 208B Grand Caravan
Fuselage
The fuselage length for the C208B is 41 feet 7 inches (12.67 m), compared to 37 feet 7 inches (11.46 m)
for the shortbody C208. The construction of the fuselage is a conventional formed sheet metal bulkhead,
stringer and skin referred to as semi-monocoque. Stringers are long metal strips which run the length of
the fuselage and are fastened to the rings and bulkheads. The principle of the monocoque design is that
the skin is designed to carry the loads and stresses. The circular shape is being maintained by the use
of frames and stringers. In addition, the stringers are designed to stiffen the skin. Longerons take up the
end loadings due to bending.
Major components of the structure are:
• The front and rear carry-through spar and bulkhead. The front carry-through spar and bulkhead is
an integral fail-safe structure with forgings at the top for attaching the front wing spar and forgings at
the bottom for the attaching the wing strut.
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Cessna 208B Grand Caravan
•
•
The rear carry-through spar and landing gear bulkhead. The rear carry-through spar and landing
gear bulkhead is an integral fail-safe structure with forgings at the top for attaching the rear wing
spar and forgings at the bottom for attaching the main landing gear trunnions.
The forward door post. The forward door post provides the load path for transferring the loads from
the engine mount directly to the primary structure.
Fail-safe construction assures that the structure is designed and built in such a way that should any
structural component fail, the remaining structure is capable of carrying certified limit flight loads.
Wings
The wing span is 52 feet 1 inch (15.88 m). The externally braced wings are constructed of a front and
rear spars, formed sheet metal ribs, doublers and stringers. Ribs maintain the aerodynamic shape of the
wing and transfer loads to the skin and spars. The entire structure is covered with aluminium skin.
The primary wing spars, wing carry-through spars in the fuselage, wing struts and attaching structure
are of fail-safe construction for limit flight loads. The front spar is equipped with wing-to-fuselage and
wing-to-strut attach fittings. The rear spar is equipped with wing-to-fuselage attach fittings.
Figure 1-2 C208B Wing Construction
Both wings contain an integral fuel tank. These tanks are formed by the front and rear spars, upper and
lower skins and inboard and outboard closeout ribs. The inboard tank ends are located 18 inches from
the wing root to form a “dry bay” for keeping fuel away from the cabin in an accident. The wing is
designed to accept impact outboard of the fuel tank while minimising damage to the fuel tank area.
Empennage
The height is 15 feet 5 inches (4.70 m). The empennage consists of a conventional vertical stabiliser,
rudder, horizontal stabiliser and elevator. The vertical stabiliser consists of a forward and aft spar, sheet
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metal ribs and reinforcements, four skin panels, formed leading edge skins and a dorsal fin. The
horizontal stabiliser is constructed of a forward and aft spar, ribs and stiffeners four upper and four lower
skin panels and two left and right wrap around skin panels which also form the leading edges.
Figure 1-3 C208B Stabilizer Vortex Generators
A problem of marginal nose-down elevator power was observed in transitional out-of-trim flight
evaluations. This was alleviated by a single row of vortex generators fitted on top of the horizontal
stabiliser just forward of the elevators. They enhance nose down elevator and trim authority. The
generators are intended to prevent flow separation over the elevators at slow speeds.
Cargo Pod
The pod attaches to the bottom of the fuselage with screws and can be removed. The pod is
fabricated with a Nomex inner housing, a layer of Kevlar and a fibreglass outer layer.
Figure 1-4 C208B Cargo Pod
The pod has a load-carrying capacity of 1090 pounds (494 kg’s). It has 4 separate compartments
divided by aluminium bulkheads. Each compartment has a maximum floor loading of 30 lbs/sq.ft. The
maximum weight for each compartment is as follows:
• Forward compartment
230 lbs. (104 kg’s)
• Centre compartment fwd
310 lbs. (140 kg’s)
• Centre compartment aft
270 lbs. (122 kg’s)
• Aft compartment
280 lbs. (127 kg’s)
Each compartment has an individual loading door. Each door is secured in the closed position by 2
handles which latch the doors when rotated 90° to the horizontal position.
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Dimensions
Figure 1-5 C208B 3-View
Length ...............................41’ 7” .........12.67 m
Wingspan ..........................52’ 1” .........15.88 m
Height ............................... 15’ 5”..................4.70 m
Wheelbase Length......... 13’ 3½”..................4.05 m
Hartzell Propeller Ground Clearance
Standard Nose-gear Fork........... 14 ¼” ..36 cm
Nose Strut Fully Compressed ...... 5 ½” ..14 cm
Extended Nose-gear Fork .......... 17 ¾” ..45 cm
Nose Strut Fully Compressed ......8 7/8” ..22 cm
McCauley Propeller Ground Clearance
Standard Nose-gear Fork ...........11 ¼” .........29 cm
Nose Strut Fully Compressed.......2 ½” ...........6 cm
Extended Nose-gear Fork ..........14 ¾” .........37 cm
Nose Strut Fully Compressed...... 5 7/8” .........15 cm
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Landing Gear
The landing gear is a fixed-gear, tricycle type with steerable nose wheel and two main wheels. To
ensure the stability of the aircraft on the ground while loading, a tail-stand may be temporarily attached
to the rear tie-down point.
Figure 1-6 C208B Landing Gear And Tail-stand
Main gear shock absorption is provided by tubular spring-steel main landing gear struts and an
interconnecting spring-steel tube between the two main landing gear struts. The interconnecting tube
reduces bending in the fuselage landing gear bulkheads. The main landing gear has a “tear away”
feature which reduces damage to the fuselage structure in the event of an accident.
Nose gear shock absorption is provided by an oil-filled oleo shock strut, combined with a drag link spring
providing vertical and aft displacement restraint. Fail-safe features of the drag link include absorption of
some energy even with complete oil loss and retains the nose gear in the event of torque link failure. To
improve operation from unpaved surfaces, the standard nose gear fork and tyres can be replaced by a
3" extended nose gear fork and oversized tyres.
Figure 1-7 C208B Standard And Extended Nose Gear Forks, Nose Gear Over Travel Indicator
Nose gear steering is accomplished by using the rudder pedals. A spring-loaded steering bungee, which
is connected to the nose gear and to the rudder bars, will turn the nose gear through an angle of 15°
each side of centre. By applying differential braking the degree of turn may be increased up to 56° each
side of centre. If the nose wheel is turned beyond its limits during towing, a frangible stop will fracture
and the over travel indicator block hanging on a small cable, will fall into view alongside the nose strut.
Brake System
A single-disc, hydraulically-actuated brake is fitted on the inboard side of each main landing gear wheel.
Each brake is connected by a hydraulic line to a master cylinder attached to each pilot's rudder pedal. A
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single park brake handle is located on the main instrument panel, below the pilot’s control column. To
set the park brake, apply brake pressure and pull the handle aft. To release the parking brake, push the
handle fully in.
Figure 1-8 C208B Park Brake Handle And Hydraulic Reservoir
A brake fluid reservoir located just forward of the firewall on the left side of the engine compartment
provides additional brake fluid for the brake master cylinders. The fluid in the reservoir should be
maintained between the MIN and MAX level prior to each flight. The brake system uses MIL H-5606
hydraulic fluid. This fluid is a mineral base type and is a flammable petroleum product and coloured red.
Systems using mineral base fluids incorporate synthetic rubber seals.
Later model aeroplanes (208B01030 and above) have metallic type brakes. When conditions permit,
hard braking is beneficial in that the resulting higher brake temperatures tend to maintain proper brake
glazing and will prolong the expected brake life. The habitual use of light and conservative brake
application is detrimental to metallic brakes.
Minimum Turning Distance
Figure 1-9 C208B Minimum Turning Radius
Minimum Turn Wheel Travel ..........28’ 1” .... 8.56 m
Minimum Turn Wingtip Travel ........65’ 5” .. 19.94 m
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Flight Controls
The flight controls consist of a conventional aileron, elevator and rudder control surfaces and a pair of
spoilers mounted above the outboard ends of the flaps. The flight control surfaces are manually
operated through mechanical linkages. All control cables are of stainless steel construction.
• Ailerons
Round-nose ailerons are of conventional formed sheet metal ribs and smooth aluminium skin
construction. Aileron trimming is achieved by a trimmable servo tab attached to the right aileron and
connected mechanically to a knob located on the control pedestal. The servo tab on the left aileron
provides reduced manoeuvring control wheel forces. Fences inboard on the ailerons on the shortbody
models C208 and C208A, enhance lateral stability. These fences are also included as options on the
longbody C208B where they are called “payload extenders” and allow an increase of 300lbs in the
MTOW.
Figure 1-10 C208B Wing Fence, Aileron and Trim Control
• Spoilers
Due to the long wing and greater fuel load further out in the wings, spoilers are fitted to improve lateral
control at low speeds by disrupting lift over the appropriate flap. The effect is more enhanced in a flap
down configuration. This feature, in turn, permits shorter length ailerons to accommodate the extremely
long-span wing flaps needed for meeting the FAA’s maximum stall speed limit of 61 knots.
The spoilers are interconnected to the aileron system through a push-rod mounted on an arm on the
aileron. Movement for the first 5° of aileron travel is negligible. Once the aileron has been deflected
upward past the 5° point, spoiler travel is proportional to aileron travel. The spoiler is fully retracted when
the aileron is deflected downward. Maximum deflection angle is 40°.
Figure 1-11 C208B Spoiler
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• Elevators
Construction of the elevator consists of a forward and aft spar, sheet metal ribs, upper and lower skin
panels and wrap-around skin panels for the leading and trailing edges. Both elevator tip leading edge
extensions provide aerodynamic balance and incorporate balance weights.
Elevator trimming is accomplished through two elevator trim tabs attached to the trailing edge of each
elevator by full length piano-type hinges. To minimise the effects of the strong slipstream, each trim tab
is located as far outboard on the elevator as practical. Dual pushrods from each actuator transmit
actuator movement to dual horns on each elevator trim tab to provide tab movement. The elevator trim
tabs are connected to a vertically mounted trim wheel on the control pedestal.
Trimming is manually or electrically. The left
half of the switch provides power to engage the
trim servo clutch. The right half controls the
direction of motion of the trim servomotor.
Operation of the electric trim will disconnect the
autopilot.
Figure 1-12 C208B Elevator Trim Controls
• Rudder
The rudder is constructed of a forward and aft spar, formed sheet metal ribs
and reinforcements and a wrap-around skin panel. The top of the rudder
incorporates a leading edge extension which contains a balance weight.
Figure 1-13 C208B Rudder Trim Control
Trimming is not accomplished through a trim tab, but rather internally through
the nose wheel steering bungee mechanically connected to the rudder control
system. The trimming system acts against the steering bungee to displace the
pedals and move the rudder. The rudder trim setting is controlled by a
horizontal trim control wheel mounted on the control pedestal.
Control Locks
A manual control lock inserted into the pilot’s control column secures the ailerons in the neutral position
and the elevators in slightly trailing-edge-down position. Ensure that the warning flag is over the sidewall
switch panel.
Figure 1-14 C208B Control Lock And Rudder Lock
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A rudder lock locks the rudder in the neutral position. Earlier serial models are equipped with a rudder
lock which is operated by a spring-loaded, pull-type T-handle located on the bottom of the instrument
panel to the right of the control pedestal. An interlock between the rudder lock and the fuel condition
lever prevents locking the rudder when the fuel condition lever is in any position other than CUTOFF.
The handle is released from the locked position by rotating it 90° and allowing it to retract forward to the
unlocked position.
Later serial models incorporate an external rudder gust lock located on the left side of the tail cone. A
fail-safe connection automatically disengages the lock when the elevator is deflected upward about onefourth of its travel from neutral. Because of the fail-safe system, the elevator lock should always be
engaged prior to engaging the rudder lock when securing the aircraft after shutdown. The rudder lock
should be disengaged before towing, starting the engine or moving the aircraft on the ground in any
manner.
Wing Flap System
The wing flaps are large span, single-slotted, semi-fowler type flaps. The flaps span 70% of the wing,
increasing the area by 15%. The outboard portions have rubber leading edge vortex generators, and a
trailing edge lip to prevent flap vibration.
The flaps are driven by an electric motor. The wing flaps are extended or retracted by positioning the
WING FLAP SELECTOR LEVER on the control pedestal to the desired position. A slotted panel
provides mechanical stops at 10° and 20° positions. A white tipped pointer provides flap position
indication. The system is protected by the FLAP MOTOR circuit breaker on the circuit breaker panel.
Figure 1-15 C208B Flaps And Flap Controls
• Standby Flap System
A standby system can be used to operate the flaps if the primary system has a malfunction. It consists
of the following: a standby flap motor, a guarded standby flap motor switch and standby flap motor
up/down switch. The guarded NORM position permits operation of the flaps using the selector on the
control pedestal. The STBY position disables the dynamic breaking of the primary flap motor when the
standby flap motor system is operated. The UP/DOWN switch has a UP, centre-off and DOWN position.
To operate the flaps with the standby system:
• Select the flap lever to the desired position.
• Lift the guard and place the standby flap motor switch in the STBY position.
• Actuate the standby flap motor UP/DOWN switch momentarily to UP or DOWN as desired.
Observe the flap position indicator to obtain the desired position. As the standby flap system has no limit
switches, actuation of the UP/DOWN switch should be terminated before the flaps reach full up or down
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travel, otherwise damage to the standby flap motor
mounts may result.
Use of the standby flap system should be avoided
with the autopilot engaged since it will cause the trim
to run in opposite direction to the autopilot inputs.
Figure 1-16 C208B Standby Flap Controls
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COCKPIT AND CABIN
Table of Contents
Instrument Panel ......................................... 2-1
Control Pedestal.......................................... 2-2
Left Sidewall Switch & C/B Panel ............... 2-2
Overhead Panel .......................................... 2-3
Pilot Seats ................................................... 2-3
Aircraft Doors .............................................. 2-3
Crew Entry Doors.........................................2-4
Passenger Door ...........................................2-4
Cargo Door ..................................................2-5
Interior Lighting ............................................2-6
Exterior Lighting ...........................................2-7
Instrument Panel
The flight instruments layout is designed around the basic “T” configuration. Immediately to the left of
the flight instruments are the clock, propeller anti-ice ammeter, suction gauge, volt/ammeter,
volt/ammeter selector switch, propeller overspeed governor test switch, left air vent pull knob, left air
vent outlet and microphone and headset jacks.
Figure 2-1 C208B Instrument Panel
The lower left side of the instrument panel contains a switch panel for the switches necessary to
operate the aircraft systems. These include toggle switches for the exterior and interior lights, de-ice
and anti-ice systems.
Below the flight instrument panel are the parking brake, lighting switch panel and inertial separator
control.
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Above the flight instrument panel are the annunciator panel, annunciator panel day/night switch,
annunciator test switch and fire detector test switch.
Avionics equipment is placed vertically in dual stacks approximately in the centre and just to the right
of the instrument panel.
Located above the avionics stacks in the top centre of the instrument panel are the engine
instruments consisting of the torque indicator, propeller RPM indicator, ITT indicator, NG% RPM
indicator, oil pressure/oil temperature gauge, fuel flow indicator and left and right fuel quantity
indicators.
Directly to the right of the right flight instrument panel are the hour meter, right fresh air vent pull knob,
right fresh air vent outlet and microphone and headset jacks and map compartment.
Located below the avionics stacks are the cabin heat switch and control panel. Air-conditioning
controls, if installed, are also located here.
Control Pedestal
A control pedestal extending from the centre of the instrument
panel to the floor contains the emergency power lever, power
lever, propeller control lever, fuel conditioning lever and wing flap
selector and position indicator. A quadrant friction lock is located
on the right side of the pedestal.
The elevator trim control and position indicator is located on the
left side of the pedestal. The aileron and rudder trim controls and
position indicators, the fuel shutoff valve control and cabin heat
firewall shutoff valve control and a microphone are located on the
lower end of the pedestal.
Figure 2-2 C208B Control Pedestal
Left Sidewall Switch And Circuit Breaker
Panel
Control switches for the engine, electrical systems and
avionics systems are located on a separate panel
mounted on the left cabin sidewall adjacent to the pilot.
The circuit breakers are located on the side of the panel,
running down to the cockpit floor.
Figure 2-3 C208B Left Sidewall Panel
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Overhead Panel
The left and right fuel tank selectors, standby flap motor
switches, left and right vent air control knobs, and left and
right pilots’ vent air outlets are located on the overhead
panel.
Provisions for an optional oxygen shutoff valve and
pressure gauge are also provided.
Figure 2-4 C208B Overhead Panel
Pilot Seats
The six-way adjustable pilot’s seats may
be moved forward or aft, adjusted for
height and the seat angle changed.
Position the seat by pulling on the small Thandle under the centre of the seat bottom
and slide the seat into position; then
release the handle and check that the seat
is locked into place.
Raise or lower the seat by rotating a large
crank under the front right corner of the
seat. Seat back angle is adjusted by
rotating a small crank under the front left
corner of the seat. The seat bottom angle
will change as the seat back angle
changes. As a safety feature, the seat is
designed to deform with a high g-load, to
minimise injury to the pilot.
Later model aircraft seats are equipped
with armrests which can be moved to the
side and raised to a position beside the
seat back for stowage.
Figure 2-5 C208B Pilot Seat
Both pilots’ seats are equipped with a five-point restraint system which combines the function of
conventional type seat belts, a crotch strap and an inertia reel equipped double-strap shoulder
harness in a single assembly.
Aircraft Doors
Entry to and exit from the aircraft is accomplished through an entry door on each side of the cockpit and
on the Passenger version only, through a two-piece, airstair door on the right side. A cargo door on the
left side can also be used for cabin entry.
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x Crew Entry Doors
The left crew entry door incorporates a conventional interior and exterior door handle, a lock override
knob a key-operated door lock and an openable storm window. The right crew entry door incorporates a
conventional interior and exterior door handle and manually operated inside door lock.
To open either door from the outside, rotate the handle down and forward to the OPEN position. To
close the door from the inside, place the handle in the CLOSE position and pull the door shut; then
rotate the handle forward to the LATCHED position. When the handle is rotated to the LATCHED
position, an over-centre action will hold it in position. The lock override knob on the inside of the left door
provides a means of overriding the outside door lock from inside the aeroplane. To operate the override,
pull the knob and rotate it in the placarded direction to unlock or lock the door.
Figure 2-6 C208B Crew Entry Doors
The door is held open by clipping a small chain, attached to the bottom of the door, to a small ring on the
cowling. As this chain cannot be reached by the pilots from their seats, ensure that the chain is
disconnected before entering the aircraft for flight. To hold the door open while the engine cowlings are
open, an attachment arm from the lower cowling can be extended out to the bottom of the door.
To allow easy access to the cockpit, folding ladders are fitted to the floor beside each door.
Great caution must be taken when using the right-hand crew door due to the close proximity of the
engine exhaust. Even after the engine has been shut down, the exhaust will remain hot for some time,
and serious burns are possible if the exhaust is inadvertently touched. The right-hand crew ladder will
also be a burning hazard if it remains extended with the engine running.
x Passenger Door
The passenger door consists of an upper and lower section. When opened, the upper section swings
upward and the lower section drops down providing integral steps. The upper door section incorporates
a conventional exterior door handle with a separate key operated lock, a push button exterior door
handle release and an interior door handle which snaps into a locking receptacle. If the upper door is
not properly latched, a red light labelled DOOR WARNING will illuminate on the annunciator panel.
The lower door features a flush handle which is accessible from either inside or outside. This handle
is designed so that when the upper door is closed, the handle cannot be rotated to the open position.
The lower door also contains integral door support cables and a door lowering device.
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Figure 2-7 C208B Passenger Door
To enter the aircraft through the passenger entry door, depress the exterior pushbutton door release,
rotate the exterior door handle on the upper door section clockwise to the open position and raise the
door to the over-centre position. Release the lower section by pulling up on the inside door handle
and rotating the handle to the open position. Lower the door section until it is supported by the integral
support cables.
Do not use the outside key lock to lock the door prior to flight since the door could not be opened from
the inside if it were needed in an emergency evacuation.
x Cargo Door
A two-piece cargo door is located on the left side of the fuselage. The door is divided in an upper and a
lower section. When opened, the upper section swings upwards and the lower section swings forward.
The upper section of the door incorporates a conventional exterior door handle with a separate key
operated lock and on the Passenger version only, a pushbutton exterior emergency door handle release
and an interior door handle which snaps into a locking receptacle. If the upper door is not properly
latched, a red light labelled DOOR WARNING, will illuminate on the annunciator panel.
Figure 2-8 C208B Cargo Door
The lower door features a flush handle which is accessible from either inside or outside. This handle
is designed so that when the upper door is closed, the handle cannot be rotated to the open position.
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In an emergency do not attempt to exit the Super Cargomaster version through the cargo doors
without outside assistance. Since the inside of the upper door has no handle, exit from the aeroplane
through these doors without outside assistance is not possible.
Do not use the outside key lock to lock the door prior to flight since the door could not be opened from
the inside if it were needed in an emergency evacuation.
Interior Lighting
Figure 2-9 C208B Interior Lighting Controls
Instrument Lighting
Instrument and control panel lighting is provided by integral, flood and post lights. Four concentric-type
dual lighting control knobs are grouped together on the lower part of the instrument panel. Controls vary
the intensity of the instrument panel lighting, the left sidewall switch and circuit breaker lighting, the
pedestal lighting and the overhead panel lighting.
Cabin Light
Four cabin lights are installed in the interior of the airplane. These lights are controlled either by a twoposition toggle switch, labelled CABIN, on the lighting control panel or rocker-type switches located on
the inside sidewall panel just forward of the cargo door and the passenger entry door. This light circuit
does not require power to be applied to the main electrical system buses for operation. The courtesy
lights circuit may be equipped with a solid-state timer which allows the lights to remain illuminated for
period of 30 minutes.
No Smoking/Fasten Seatbelt Sign
This installation consists of a small lighted panel mounted in the cabin headliner above the right side of
the forward cabin area. The lights are controlled by two toggle-type switches, labelled SEAT BELT and
NO SMOKE. These warning sign lights are protected by a circuit breaker labelled SEAT BELT SIGN.
Map Light
A map light is mounted on the bottom of the pilot's control wheel. Brightness is adjusted by means of a
rheostat control knob on the bottom of the control wheel.
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Passenger Reading Lights
Passenger reading lights may be installed near each of the aft passenger positions. The lights are
located in small convenience panels above the seats. A pushbutton-type ON, OFF switch, mounted in
each panel, controls the lights.
Exterior Lighting
Figure 2-10 C208B Exterior Lighting Controls
All exterior lights are controlled by toggle switches located on the lighting control panel on the left side of
the instrument panel. The switches are ON in the up position and OFF in the down position.
Landing Lights
Two landing lights are mounted in the outboard section of each wing leading edge. They provide
illumination forward and downward during takeoff and landing. The lights are protected by two circuit
breakers labelled LEFT LDG LT and RIGHT LDG LT. The landing lights have a relatively short service
life and should be used for takeoff and landing only.
Taxi/Recognition Lights
Two taxi/recognition lights are mounted inboard of each landing light. These lights are focused to
provide illumination of the area forward of the aircraft during ground operation. The lights are also used
in the traffic pattern or enroute. They are protected by a circuit breaker labelled TAXI LIGHT.
Strobe Lights
A strobe lights with remote power supply is installed on each wingtip. The lights are used to enhance
anti-collision protection and should be turned OFF when taxiing. Do not operate the strobe lights in
conditions of fog, cloud or haze as the light beam can cause vertigo. The lights are protected by a circuit
breaker labelled STROBE LIGHT.
Navigation Lights
Conventional navigation lights are installed on the wingtip and the tail cone stinger. The lights are
protected by a circuit breaker labelled NAV LIGHT.
Flashing Beacon Light
A red flashing light is installed on top of the vertical fin. The light is visible through 360°. The light is
protected by a circuit breaker labelled BEACON LIGHT. Do not operate the flashing beacon when flying
through clouds as the reflected light can cause vertigo.
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Courtesy Lights
A courtesy light is installed under each wing. The lights illuminate an area adjacent to the crew entry
doors. The lights operate in conjunction with the cabin lights and are controlled by the cabin light
switches.
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WARNING SYSTEMS
Table of Contents
Annunciator Panel....................................... 3-1
Fuel Selectors OFF Warning System ......... 3-2
Stall Warning System.................................. 3-3
Overspeed Warning System........................3-3
Engine Fire Detection System .....................3-3
Annunciator Panel
An annunciator panel to advise the pilots of specific conditions in the aircraft systems is mounted below
the glareshield on the captain’s instrument panel. A green coloured lamp indicates a normal or safe
condition in the system. An amber lamp indicates a cautionary condition exists which may or may not
require immediate corrective action. A red lamp indicates a hazardous condition requiring immediate
corrective action.
Figure 3-1 C208B Annunciator Panel
Two annunciator panel function switches, labelled LAMP TEST and DAY/NIGHT are located to the left
of the panel. By pressing LAMP TEST all annunciators are illuminated and both fuel-selector-off warning
horns are activated. With the DAY/NIGHT SWITCH in the DAY position, any annunciator that is
illuminated will be at full intensity. In the NIGHT position it provides intensity down to a preset minimum
dim level for the green lamps and the following amber lamps:
x Left and right fuel low.
x Standby electrical power inoperative.
x Standby electrical power on.
Annunciator
Cause For Illumination
An excessive temperature condition and/or possible engine fire has occurred in
the engine compartment.
Engine oil pressure at or below 40+2 psi.
The generator is not connected to the airplane bus.
The emergency power lever is advanced out of the NORMAL position.
The auxiliary fuel pump is operating.
Fuel pressure in the fuel manifold is below 4.75 psi.
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AIRCRAFT TECHNICAL NOTES
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The starter-generator is operating in the starter mode.
Electrical power is being supplied to the engine ignition system. 
Electrical system bus voltage is below 24.5 volts, and power is being supplied
from the battery.
The vacuum system suction is less than approximately 3.0” Hg.
The fuel level in the reservoir tank is approximately half-full or less.
Fuel quantity in the left fuel tank is approximately 25 USg’s (95 ltr’s/ approximately
146 lb’s) or less. 
Fuel quantity in the right fuel tank is approximately 25 USg’s (95 ltr’s/
approximately 146 lb’s) or less. 
Electrical power is not available from the standby alternator. 
The selected inverter or has not been turned ON (if installed). 
One or both fuel selectors are OFF, the fuel selector warning circuit breaker is not
set or the start control circuit breaker is not set.
The upper passenger door (passenger version only) and/or the upper cargo door
are not latched.
NiCad battery only - the electrolyte temperature is critically high (160ºF/ 71ºC)
NiCad battery only - the electrolyte temperature is high (140ºF/ 60ºC)
Metal chips have been detected in either the reduction gearbox case or the
accessory gearbox case. 
The standby alternator is supplying electrical power to the bus. 
Electrical power is being supplied to the windshield anti-ice power relay (if
installed). 
Pressure in the de-ice boot system has reached approximately 15 psi. 
 Preset Dim
The intensity can be controlled by the ENG INST lighting rheostat. A lamp may be extinguished by
pushing on the face of the light assembly and allowing it to pop out. To reactivate the annunciator, pull
the light assembly out slightly and push it back in.
Fuel Selectors OFF Warning System
The system consists of redundant warning horns, a red annunciator light labelled FUEL SELECT OFF,
actuation switches and electrical hardware. The dual aural warning system is powered through the
START CONT circuit breaker. A non-pullable FUEL SEL WARN circuit breaker is installed in series to
protect the integrity of the starter system. The annunciator is powered from the ANNUN PANEL circuit
breaker. The warning system functions as follows:
A.
BOTH SELECTORS OFF
The red FUEL SELECT OFF annunciator illuminates. A single fuel select off warning horn is
activated.
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B.
DURING ENGINE START WITH EITHER SELECTOR OFF
The red FUEL SELECT OFF annunciator illuminates. Both fuel select off warning horns are
activated.
C.
ONE SELECTOR OFF AND FUEL REMAINING IN THE TANK BEING USED IS LESS THAN
25 GALLONS
The red FUEL SELECT OFF annunciator illuminates. One fuel select off warning horn is
activated.
If the FUEL SEL WARN circuit breaker has popped or the START CONT circuit breaker has been
pulled, the red FUEL SELECT OFF annunciator will illuminate even with both tanks ON.
Stall Warning System
A vane-type stall warning unit is fitted in the leading edge of the left wing. The unit is electrically
connected to the stall warning horn located overhead the pilot. It senses a change in airflow over the
wing and operates the warning horn at airspeeds between 5 and 10 knots above the stall in all
configurations. The system is protected by the STALL WRN circuit breaker. The vane and sensor unit is
equipped with a heating element operated by the STALL HEAT switch and is protected by the STALL
WRN circuit breaker.
Overspeed Warning System
An airspeed pressure switch in the pitot-static system is used to actuate an airspeed warning horn in the
event o excessive airspeed. The horn is located behind the headliner in the area above the pilot and will
sound when the airspeed exceeds VMO (175 KIAS). A warning signal will also be heard in the pilot’s
headsets.
Engine Fire Detection System
The engine fire detection system consists of a heat sensor in the engine compartment, a warning light
on the annunciator panel and a warning horn.
The heat sensor consists of 3 flexible, closed loops. Each section of the loop is made up of a single wire
surrounded by a continuous string of ceramic beads contained in an inconel tube. The outer shell is
connected to ground at the firewall and the wire inside is connected to the control box. The beads have
a characteristic that causes them to lower their electrical resistance when the sensing element reaches
its preset temperature value.
The core material in both elements prevents electrical current from flowing at normal temperatures.
When the elements are exposed to increasing temperatures, a current flows between the signal wire
and ground. The control box detects a change in resistance in the loops and will illuminate the ENGINE
FIRE light and trigger the warning horn.
The fire warning is initiated when temperatures in the engine compartment exceed 425°F (218°C) at the
firewall, 625°F (329°C) around the exhaust, or 450°F (232°C) at the rear engine compartment. A test
switch, labelled FIRE DETECT TEST, is located adjacent to the annunciator panel. When depressed,
the ENGINE FIRE annunciator will illuminate and the fire warning horn will sound, indicating that the fire
warning circuitry is operational. The system is protected by a “pull-off” type circuit breaker labelled FIRE
DET, positioned third row and second from the left.
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TURBINE ENGINE THEORY
Table of Contents
Introduction ................................................. 4-1
Air Inlet Duct................................................ 4-2
Compressor................................................. 4-3
Centrifugal Compressor .............................. 4-3
Axial-flow Compressor ................................ 4-3
Compressor Stall......................................... 4-4
Combustion Chamber..................................4-4
Turbine.........................................................4-4
Exhaust ........................................................4-5
Turbo-propeller Engine Power Conversion .4-5
Fixed-shaft Engine .......................................4-5
Free-turbine Engine .....................................4-5
Introduction
Turbine engines are made up of the following sections:
x Air inlet
x Compressor
x Combustion section
x Turbine
x Exhaust
x Accessory section
x Support systems (starter, lubrication, fuel and auxiliary components such as anti-icing, airconditioning and pressurisation)
Figure 4-1 PT6A-114A Overview
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Air Inlet Duct
The purpose of the air inlet duct is to channel the incoming air to the compressor with the least
possible loss of energy. In addition to the design of the inlet, three other variables determine how
much air will pass through the compressor:
x Ambient air density
x Airspeed of the aircraft
x Rotational speed of the compressor
Compressor
Two main functions of the compressor are to supply compressed air to the burners, and to supply bleed
air to fulfil various engine and airframe system requirements. To supply compressed air to the burners,
the compressor increases the pressure of the mass of air that is channelled through the air inlet ducts
and routes it to the burners in the amount and pressure required.
Bleed air might be tapped off any of the various compressor stages. With each successive stage the
pressure and temperature of the air increases. The specific stage for bleed air extraction is determined
by the temperature and pressure required for the job.
x Centrifugal Compressor
The compressor accelerates the airflow by slinging it outward from the centre of the compressor wheel.
The primary components of the centrifugal compressor are the impeller (rotor), a diffuser (stator) and a
compressor manifold. The impeller catches the incoming air and accelerates it outward toward the
diffuser. The purpose of the diffuser vanes is to direct the air from the impeller to the manifold at the
gentlest angle possible to retain the maximum amount of energy. Another purpose of the diffuser is to
provide the combustion chamber with air at the correct velocity and pressure for maximum efficiency.
The air is allowed to expand into a divergent duct. As the air spreads out, the velocity drops and the
static pressure increases. Multiple stages are connected by a system of convergent and divergent ducts
to compress and direct the airflow to succeeding stages at the proper angle.
Figure 4-2 Exploded View Of Typical Single-stage Centrifugal Compressor
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Advantages of centrifugal compressors:
x High pressure rise per stage. Up to 10:1 in a single stage.
x Good efficiency (compression) over a wide rotational speed range, idle to full power.
x Simplicity of manufacture and low cost.
x Low weight.
x Low starting power requirements.
x More resistant to foreign object damage.
Disadvantages of centrifugal compressors:
x Large frontal area for a given airflow.
x More than two stages is not practical due to energy loss in the airflow when making the turn from
one impeller to the next.
x Axial-flow Compressor
The airflow follows a path parallel to the axis of rotation, hence the name. Airflow is compressed by
accelerating it across multiple compressor stages. Airflow passes through a convergent duct to force it
into an even smaller area. The size of the compressor blades decreases from the first to the last to
follow the narrowing design of the convergent duct.
Figure 4-3 Typical 4-Stage Axial-flow Compressor
The main elements of an axial-flow compressor are rotors and stators. Rotor blades are fixed on a
rotating spindle and force air rearward in the same manner as a fan. The rotor blades are preset at a
specific angle and are contoured similar to small propeller blades. Stators, which are stationary bladetype airfoils, are located between each compressor stage. The stators act as a diffuser in each stage
converting part of the energy air into pressure by slightly slowing down the air. They also direct the
airflow into succeeding compressor stages at the correct angle.
The first set of stators, located before the compressor stage, is referred to as inlet guide vanes. A similar
set of guide vanes, called straightening vanes, might be located at the compressor exit to stop the
exiting air mass from rotating as it moves forward into the combustion chamber.
Advantages of axial flow compressors:
x High peak efficiencies (compressor pressures ratio) created by its straight through design allowing
higher ram efficiency.
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x
x
Higher peak efficiencies (pressure) attainable by addition of compression stages if desired.
Small frontal area for a given airflow resulting in low drag.
Disadvantages of axial flow compressors:
x Difficulty and high cost of manufacture.
x Relatively high weight.
x High starting power requirements.
x Low pressure rise per stage.
x Good efficiency only over a narrow rotational speed range.
x Very susceptible to foreign object damage.
x Compressor Stall
The angle at which the incoming air meets the compressor blades varies as a result of the air’s velocity
and the compressor’s rotational speed. The blade’s combined influence forms a vector that is called the
angle of attack. An imbalance between the two vector quantities might stall the airflow.
By definition, a compressor stall occurs when the air mass travelling through the compressor slows
down and stops. In extreme cases the airflow might even reverse direction. An engine compressor stall
may be noted by a single or multiple loud “popping” noise from the engine compartment. The situation
may be resolved by reducing the power to a point where the “popping” discontinues and slowly
advancing the throttle to the required setting. The use of bleed air may also help eliminate engine
compressor stalls if this situation is encountered.
Compressor blades stall for numerous reasons:
x Anything that alters the proper operation of the compressor blades such as blade failure, foreign
object damage, etc.
x A fuel mixture that is too lean.
x Abrupt pitching movement.
x Excess fuel flow.
x Engine operating speeds well above or below normal recommended operating speed.
Combustion Chamber
The combustion chamber houses the fuel nozzles and igniters. Fuel is mixed with the compressed air in
the combustion chamber. The temperature of the air as it leaves the compressor is ±315°C. After being
processed by the combustion chamber, the air temperature is ±870°C.
To accomplish its task of efficiently burning the fuel/air mixture, the combustion chamber must:
x Mix the fuel and air in the proper manner for the ambient conditions to assure the best possible
combustion.
x Cool the hot gasses to a temperature within the normal operating range of the turbine wheel.
x Channel the exhaust gases in such a manner that it maximises turbine wheel effectiveness.
The combustor lining is designed for proper flow of air into and out of the combustor which provides for
proper flame patterns. No flame is allowed to touch the combustor lining. In addition the combustor lining
assures that only a small amount of the total airflow (approximately 25%) is mixed with fuel and burned.
The rest of the airflow provides cooling and dilution of the products of combustion.
Turbine
A turbine section is designed to extract as much energy as possible from the high velocity airflow issuing
from the combustor. The sole purpose is to convert the kinetic energy of combustion chamber exhaust
gasses into mechanical energy to operate the compressor and other accessories. Approximately 60 - 80
percent of the total energy of the exhaust gasses go to that purpose. The turbine section is a divergent
duct. Each stage is larger in diameter than the preceding stage to compensate for loss of energy
through each stage. This allows an equal sharing of the load between stages.
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Stators or turbine guide vanes are located in front of and between each turbine stage to smooth the
airflow and properly direct it to the next turbine wheel. High energy airflow strikes the turbine blades at
an angle. As the airflow does not easily change direction, it imparts a pressure upon the turbine blades
which causes a turning force upon the wheel. The turbine stages are mounted on the same shaft with
the compressor stages.
Exhaust
The exhaust section, which is located directly behind the turbine section has two primary purposes:
produce exhaust gas high exit velocity and minimise exhaust gas turbulence. The purpose of the
exhaust cone is to channel all exhaust gasses as they are discharged from the turbine blades and
combine them into a single, cohesive gas stream. Fixed struts also serve as a stabilising influence on
the gas and minimise the swirling motion. If this were not done, exhaust gasses would exit the engine at
an approximate 45° angle resulting in unwanted drag.
Turbo-propeller Engine Power Conversion
There are two main power output turbine designs: the fixed shaft design and the free turbine engine.
x Fixed-shaft Engine
One central shaft mounts the compressor and the turbine. The shaft extends forward into the gearbox
where high RPM is converted to low RPM suitable to drive the propeller. All rotating components within
a fixed shaft engine rotate together. Advantages of a fixed shaft system include constant engine RPM,
controlled descent capability due to prevention of windmilling overspeed, and rapid reverse thrust
availability.
x Free-turbine Engine
An airflow couple is utilised between the gas generator and power turbine shaft. The gas generator
turbine powers the compressor. The power turbine drives into the reduction gearbox in the same
manner as the fixed turbine. The gearbox converts high RPM low torque input into usable low RPM high
torque power. Because a free turbine system allows the propeller to operate independent of the
compressor, there are several advantages:
x Better control of propeller speed.
x The propeller can be held at very low rpm during taxiing, with low noise and low blade erosion.
x The engine is easier to start, especially in cold weather.
x The propeller and the gearbox do not directly transmit vibrations into the gas generator.
x A propeller brake can be installed during turnarounds.
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PT6A-114A POWERPLANT
Table of Contents
PT6A-114A Powerplant............................... 5-1
Free-turbine Reverse-flow Operation.......... 5-2
Engine Description And Operation.............. 5-2
Inertial Separator......................................... 5-5
Compressor Bleed Valve ............................ 5-5
Ignition System............................................ 5-6
Starting System........................................... 5-6
Engine Start Considerations ....................... 5-7
Exhaust ....................................................... 5-8
Engine Accessory Section ...........................5-8
Magnetic Chip Detectors .............................5-9
Engine Fuel System.....................................5-9
Lubrication System ....................................5-13
Engine Controls .........................................5-15
Engine Instruments ....................................5-16
Engine Limitations......................................5-18
Starter Limitations ......................................5-19
PT6A-114A Powerplant
The Pratt & Whitney Canada PT6A-114A is a free-turbine engine with two independent turbines. The
engine utilises two independent turbine sections: one driving the compressor in the gas generator
section, and the second driving the propeller shaft through a reduction gearbox.
The compressor section utilises 3 axial stages and 1 centrifugal stage to provide compressed air for the
combustion chamber.
Figure 5-1 PT6A-114A Engine, Internal Arrangement
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The gas generator speed is 37,500 rpm = 100% NG. Maximum permissible speed is 38,100 RPM =
101.6% NG. The power turbine speed (NF) is 33,000 RPM at a propeller shaft speed of 1900 RPM.
The engine is flat rated at 675 SHP (1865 foot-pounds of torque at 1900 RPM varying linearly to 1970
foot-pounds of torque at 1800 RPM. Below 1800 RPM, the maximum torque remains constant at 1970
foot-pounds). Torque is the measured amount of shaft horsepower delivered to the propeller. The torque
bending force is caused by the resistance of the air pushing against the advancing blade. At a lower
RPM the blade angle is increased to take a bigger “bite” of air therefore increasing the resistance and
hence the torque.
The engine is supported in a steel 9-element space frame. This frame attaches to the firewall at 5 points.
To minimise directional and longitudinal deviations with changes in thrust, the engine is canted at 3°
nose-down and 5° to the right.
Figure 5-2 C208B Engine Offset Down And Right
Free-turbine Reverse-flow Operation
The free-turbine design of the PT6A engine refers to the fact that the turbine sections rotate freely,
having no physical connection between them. The compressor turbine drives the engine compressor
and engine accessories. Dual power turbines drive the power section and propeller through the
planetary reduction gearbox. The compressor and power turbines are mounted on separate shafts
and are driven in opposite directions by the gas flow across them.
Figure 5-3 Free-turbine Reverse-flow Turboprop Engine
Reverse flow refers to the direction of airflow through the engine. Inlet air enters the compressor at
the aft end of the engine, moves forward through the compressor section and the turbines, and is
exhausted at the front of the engine.
Engine Description And Operation
The engine air inlet is located at the front of the engine nacelle to the left of the propeller spinner. Ram
air entering the inlet flows through the ducting and an inertial separator before entering through a
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circular plenum chamber where it is directed to the compressor guide vanes. The compressor air inlet
incorporates a mesh screen which will prevent entry of large particles, but does not filter the air.
Inlet air is then ducted through an annular plenum chamber, formed by the compressor inlet case,
where it is directed to the compressor. The compressor consists of three axial stages combined with a
single centrifugal stage and assembled as an integral unit.
A row of stator vanes, located between each stage of the compressor, diffuse the air, raise static air
pressure and direct the air to the next stage of compression. The compressed air passes through
diffuser tubes which turn the air through 90º in direction, and converts the velocity to static pressure.
The diffused air then passes through straightening vanes to the annulus surrounding the combustion
chamber liner assembly.
The combustion chamber liner consists of two annular sections bolted together at the front, domeshaped end. The outer wrapper incorporates an integral large exit duct. The liner assembly has
perforations of various sizes that allow entry of compressor delivery air. The flow of air changes
direction 180º as it enters and mixes with fuel. The fuel/air mixture is ignited and the resultant
expanding gases are directed to the turbines. The location of the inner liner eliminates the need for a
long shaft between the compressor and compressor turbine, reducing the overall length and weight of
the engine.
Figure 5-5 PT6A Combustion Chamber
Fuel is injected into the combustion liner through 14 simplex nozzles arranged in two sets of seven for
ease of starting. Each nozzle is supplied by a dual manifold consisting of primary and secondary
transfer tubes and adapters. The fuel/air mixture is ignited by two spark igniters which protrude into
the liner. The resultant gases expand from the liner, reverse direction in the exit duct zone and pass
through the compressor turbine inlet guide vanes to the compressor turbine. The guide vanes ensure
that the expanding gases strike the turbine blades at the correct angle, with minimum loss of energy.
The still expanding gases are then directed forward to the power turbine section.
The power turbine drives the propeller shaft via a reduction gearbox.
The compressor and power turbines are located in the approximate centre of the engine, with their
respective shafts extending in opposite directions. This feature provides for simplified installation and
maintenance inspection procedures. The exhaust gas from the power turbine is collected and ducted
into a single exhaust duct assembly and directed to atmosphere on the right side of the aircraft.
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Figure 5-4 PT6A-114A Engine Cross Section
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AIRCRAFT TECHNICAL NOTES
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Inter turbine temperature, T5, is monitored by an integral bus bar, probe and harness assembly
installed between the compressor and power turbines, with the probes projecting into the gas path. A
terminal block mounted on the gas generator case provides a connection point to the cockpit
instrumentation.
The engine oil supply is contained in an integral oil tank, which forms the rear section of the
compressor inlet case. The tank has a capacity of 2.5 US gallons, 9.5 litres, and is provided with
manual filler access and a direct oil level sight gauge.
Fuel supplied to the engine is pressurised by an engine-driven fuel pump, and its flow to the fuel
manifold is controlled by the FCU.
The power turbine drives a propeller through a two-stage planetary reduction gearbox, located at the
front of the engine. The gearbox contains an integral torque meter device, which is instrumented to
provide an accurate indication of engine power. A magnetic chip detector is installed at the bottom of
the gearbox.
Figure 5-6 C208B Engine System Annunciators
Inertial Separator
An inertial separator system in the air inlet duct prevents moisture particles from entering the
compressor air inlet plenum when in bypass mode. It consists of two moveable vanes and a fixed airfoil
which, during normal operations route the air through a gentle turn into the compressor air inlet plenum.
In the BYPASS position the vanes are positioned so that the inlet air is forced to execute a sharp turn in
order to enter the inlet plenum. Any moisture particles are separated from the inlet air and discharged
overboard through the inertial separator outlet in the left side of the cowling.
Figure 5-7 C208B Inertial Separator Airflow
Inertial separator operation is controlled by a T-handle located on the lower instrument panel. The
handle is labelled BYPASS-PULL, NORMAL-PUSH. For details of the inertial separator operation refer
to Section 11 – Ice And Rain Protection Systems.
Compressor Bleed Valve
At low NG RPM, the compressor axial stages produce more compressed air than the centrifugal stage
can use. A pneumatic piston compressor bleed valve compensates for excess air flow at low RPM by
bleeding axial stage air, P2.5, to reduce back pressure on the axial stages. This pressure relief helps
prevent axial stage compressor stall.
At low NG speeds the compressor bleed valve is open. As power is increased the valve begins to
progressively close. Above approximately 90% NG the bleed valve is closed. If the compressor bleed
valve were to stick closed at low NG speeds, compressor stall could result from an attempt to
accelerate the engine to higher power. If the valve were to stick open at high NG speeds, power output
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would be considerably reduced. With the valve open, at a given NG RPM, ITT will increase slightly and
torque will decrease.
Ignition System
Turbine engine ignition systems are designed to initially ignite the fuel/air mixture, which will then sustain
the combustion on its own. The ignition system consists of 2 igniters, an ignition exciter, 2 high tension
leads, an ignition monitor light, an ignition switch and a starter switch.
Engine ignition is provided by 2 igniters in the engine combustion chamber. The igniters are energised
by the ignition exciter mounted on the right hand side of the engine compartment. Electrical energy from
the ignition exciter is transmitted through 2 high-tension leads to the igniters in the engine. A green
annunciator, labelled IGNITION ON, will illuminate when electrical power is being supplied to the
igniters.
Ignition is controlled by an ignition switch and a starter switch. The ignition switch has two positions,
NORMAL and ON.
Figure 5-8 C208B Ignition And Start Switches
NORM Position
Arms the system so that ignition will be obtained when the starter switch is placed in the START
position. It is used for ground starts and air starts with starter assist.
ON Position
Provides continuous ignition regardless the position of the starter switch. The use of ignition for
extended periods of time will reduce ignition system component life. However, the ignition should be
turned ON to provide continuous ignition under the following conditions:
x Air starts without starter assist.
x Operation on water or slush covered runways.
x During flight in heavy rain.
x Inadvertent icing encounters until the inertial bypass separator has been in bypass for 5 minutes.
x Near fuel exhaustion as indicated by the RESERVOIR FUEL LOW light.
Starting System
The main function of the starter switch is control of the starter for rotating the gas generator portion of
the engine during start. It also provides ignition during starting. The starter/generator functions as a
motor for starting. It will motor the gas generator section until a speed of 46% NG. At 46% NG the staring
cycle will automatically be terminated. The starter switch has 3 positions, OFF, START and MOTOR.
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OFF Position
The ignition system is de-energised. This is the normal position at all times except during engine start or
engine clearing.
START Position
Rotates the gas generator section of the engine. Also energises the ignition system provided the ignition
switch is in the NORMAL position. When the engine has started the start switch must be placed in the
OFF position to activate the generator. Starter contactor operation is indicated by the amber STARTER
ENERGIZED annunciator.
MOTOR Position
Motors the engine without having the ignition circuit energised. This position is spring loaded back to the
OFF position. An interlock between the MOTOR position of the starter switch and the ignition switch,
prevents the starter from motoring unless the ignition switch is in the NORMAL position. This prevents
unintentional motoring of the engine with the ignition on.
A minimum battery voltage of 24 volts is not always an indication that the battery is near full charge or in
a good condition. This is especially true for a NiCad battery, which maintains a minimum no-load voltage
of 24 volts even at 50% or less charge condition. Therefore, if the gas generator acceleration is less
than normally observed, return the fuel condition lever to CUTOFF and discontinue the start.
Engine Start Considerations
x
If no ITT rise is observed within 10 seconds after moving the fuel condition lever to LOW IDLE, or
ITT rapidly rises to 1090°C, move the fuel condition lever to CUTOFF and perform the engine
clearing procedure.
x
With a cold engine or after making a battery start, which causes a high initial generator load due to
battery recharging, it may be necessary to advance the power lever slightly ahead of the idle detent
to maintain a minimum idle of 52% NG.
x
Since the generator contactor closes when the starter switch is turned OFF, anticipate the increased
engine load by advancing the power lever to obtain approximately 55% NG before turning the starter
switch OFF. This prevents the initial generator load from decreasing idle rpm below the minimum
52% NG.
x
Under hot OAT and/or high ground elevation conditions idle ITT may exceed maximum idle ITT
limitations of 685°C. Increase NG and/or reduce accessory load to maintain ITT within limits.
x
If, during the start, the starter accelerates the gas generator rapidly above 20% NG, suspect gear
train decouple. Do not continue the start. Rapid acceleration through 35% NG suggests a start on
the secondary nozzles. Anticipate a hot start.
x
After an aborted start for whatever reason, it is essential before the next start attempt to allow
adequate time to drain off unburned fuel. Failure to do so could lead to a hot start, a hot streak
leading to hot section damage, or torching of burning fuel from the engine exhaust on the next
successful ignition.
x
A dry motoring, within starter limitations after confirming that all fuel drainage has stopped, ensures
that no fuel is trapped before the next start.
x
If the STARTER ENERGIZED annunciator fails to go out after engine start, the generator will not
function because the start contactor may be stuck closed. The battery switch should be turned off
and the engine should be shut down if such an indication is observed.
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Exhaust
The primary exhaust is attached to the right side of the engine aft of the propeller reduction gearbox. A
secondary pipe, fitted over the end of the primary exhaust, carries the gasses away from the cowling
into the slipstream. The juncture of the primary exhaust pipe and the secondary exhaust duct is located
directly behind the oil cooler.
Since the secondary exhaust duct is of a larger diameter then the primary exhaust pipe, a venturi effect
is produced by the flow of the exhaust. This venture effect creates a suction behind the oil cooler which
augments the flow of cooling air through the cooler. This additional air flow improves oil cooling during
ground operation of the engine.
Do not leave the power lever in beta mode for more than 30 seconds with a right crosswind as the cargo
pod may be damaged by the hot exhaust gases.
Engine Accessory Section
All engine-driven accessories with the exception of the propeller governor and propeller tachometergenerator are mounted on the accessory gearbox at the rear of the engine.
Oil Pump
The oil pump is located in the lowest part of the oil tank and is driven by the accessory gear shaft.
Fuel Pump
The engine driven fuel pump is driven through a gear shaft and splined coupling. The coupling splines
are lubricated by oil mist from the auxiliary gearbox through a hole in the gear shaft. Fuel from the oilfuel-heater enters the pump through a 74 micron filter. The pressure is boosted and enters the FCU
through a 10 micron pump outlet filter. A bypass valve enables unfiltered fuel to reach the FCU in the
event of a blockage. An internal passage originating at the FCU, returns bypass fuel from the FCU to the
pump inlet downstream of the inlet screen.
NG Tacho Generator
The NG tachometer produces electric current which is used in conjunction with the gas generator RPM
indicator.
Torquemeter
The torquemeter is located within the power reduction gearbox. It is a hydro-mechanical device
connected to the first-stage reduction gear to provide an indication of engine output. It consists of a
cylinder, piston and an oil metering-type plunger. Rotation of the first-stage reduction output drive ring
gear is resisted by a helical spline which imparts an axial movement to the first stage ring gear and the
torquemeter piston. This forces the piston onto the oil valve plunger allowing engine oil to enter the
cylinder. This movement continues until the oil pressure within the torquemeter is equal to torque being
applied to the first-stage ring gear.
A bleed hole is located in the torquemeter cylinder to allow a continuous flow of oil and to bleed pressure
off when power is reduced. When engine oil is supplied to the plunger valve, it acts as a variable inlet
metering orifice. The bleed hole acts a fixed calibrated leak. On acceleration, there is more oil supplied
than bleeding away, so the pressure builds up in the torquemeter cylinder.
Starter/generator
The starter/generator is a 28 volt, 200 ampere engine driven unit. It functions as a motor for starting and,
after engine start, as a generator for the electrical system. When operating as a starter, the speed
sensing switch in the starter/generator will automatically shut down the starter, thereby providing
overspeed protection and automatic shutoff. It is air-cooled by an integral fan and by ram air ducted from
the front of the engine cowling.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
ITT Sensing System
The ITT sensing system is designed to give the pilot an accurate indication of engine operating
temperatures taken between the compressor and power turbines. It consists of 8 individual chromelalumel thermocouple probes connected in parallel. Each probe protrudes through a threaded boss on
the power turbine stator housing into an area adjacent to the leading edge of the of the power turbine
vanes.
Propeller Governor
See Section 6 – Propeller Systems.
Propeller Overspeed Governor
See Section 6 – Propeller Systems.
Engine Reduction Gearing System
The gearbox contains a two stage planetary gear chain, three accessory drives and propeller shaft. The
first-stage reduction gear is contained in the rear case, while the second stage reduction gear,
accessory drives and propeller shaft are contained in the front case.
Torque from the power turbine is transmitted to the first-stage reduction gear, from there to the secondstage reduction gear and then to the propeller shaft. Reduction ratio is from a maximum power turbine
speed of 33,000 RPM down to a propeller speed of 1900 RPM.
Oil Breather Drain Can
Some aircraft have an oil breather can mounted on the right lower engine mount truss. This can
collects any engine oil discharge coming from the accessory pads for the alternator drive pulley,
starter/generator, air conditioner compressor and the propeller shaft seal. It should be drained after
every flight.
The allowable quantity of oil discharge per hour of engine operation is 14 cc for aeroplanes with air
conditioning and 11 cc for aeroplanes without. If the quantity of oil drained from the can is greater than
specified, the source of the leakage should be identified and corrected.
Magnetic Chip Detectors
Two chip detectors are installed on the engine, one on the underside of the reduction gearbox and
one on the underside of the accessory gearbox case. The chip detectors are electrically connected to
an annunciator, labelled CHIP DETECTOR. The annunciator will illuminate when metal chips are
present in one or both of the detectors. By itself this does not demand immediate action but if
accompanied by signs of engine distress (erratic engine operation or fluctuation in engine power gage
indications), engine operation may be continued at the discretion of the pilot consistent with crew
safety.
Engine Fuel System
The engine fuel system for the PT6A-114A consists of the following basic components; oil-to-fuel heat
exchanger, an engine-driven fuel pump, a fuel control unit (FCU), a flow divider and dump valve, a
dual fuel manifold with 14 simplex nozzles, and two fuel drain valves. The system also includes the
fuel topping governor, covered separately in Section 6 – Propeller Systems. The system provides fuel
flow to satisfy the speed and power demands of the engine.
x Oil-Fuel Heater
Fuel is delivered from the reservoir to the oil-to-fuel heater. The oil-to-fuel heater utilises heat from the
engine lubrication system to preheat the fuel. A fuel temperature-sensing oil bypass valve regulates the
fuel temperature by either allowing oil to flow through the heater circuit or bypass it to the oil tank.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Figure 5-9 C208B Simplified Engine Fuel System
x Fuel Pump
Fuel enters the EDP fuel pump through a 74 micron filter. The inlet screen is spring-loaded and should it
become blocked, the increase in differential pressure will overcome the spring and allow unfiltered fuel
to enter the pump chamber.
The pump increases the fuel pressure and delivers the fuel to the FCU via a 10 micron filter in the pump
outlet. A bypass valve in the pump body enables unfiltered, high pressure fuel to enter the FCU in the
event of this filter becoming blocked.
x Fuel Control Unit (FCU)
The FCU consists of a fuel metering section, temperature compensating section, a compressor turbine
(NG) governor, a pneumatic computing section, a manual override system and a power turbine (NF)
governor.
The FCU meters proper fuel amounts for all modes of engine operation. Flow rates are calibrated for
starting, acceleration and maximum power. The FCU compares NG with power lever setting, and
regulates fuel to the engine fuel nozzles. The FCU also senses compressor discharge pressure, P3,
and compares it to NG to establish acceleration or deceleration fuel flow limits. A pre-set minimum flow
orifice guarantees sufficient fuel flow at all operating altitudes to sustain engine operation at minimum
power.
Metering Section
The metering valve input is supplied with fuel at pump pressure (P1). The fuel pressure immediately
after the metering valve is called metered fuel (P2) which flows to the fuel divider. The bypass valve
maintains a constant fuel pressure differential (P1 - P2) across the metering valve. The metering valve
consists of a contoured needle operating in a sleeve and regulates the flow of fuel by varying the orifice
area. The orifice area of the metering valve is changed by the valve movement to meet specific engine
requirements.
Fuel pump pressure (P1) in excess of requirements is returned to the fuel pump. A fuel cut-off valve is
situated downstream of the metering valve. It provides a positive means of shutting off fuel flow to the
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
engine. During normal operation the valve is fully open and offers no restriction of fuel flow to the divider.
The valve is operated by a cut-off lever which is mechanically linked to the FCU.
Figure 5-10 C208B Simplified FCU
Pneumatic Computing Section
The computing section consists of an acceleration bellows and a governing bellows connected to a
common rod. The end of the acceleration bellows is attached to the body casting and provides an
absolute pressure reference. The governor bellows is secured in the body cavity and its function is
similar to that of a diaphragm. Movement of the bellows is transmitted to the metering valve via a torque
tube assembly. This tube is positioned during assembly to provide a force in a direction tending to close
the metering valve, while the bellows act against this force to open the metering valve.
Third stage turbine discharge pressure (P3) is split up into PX and PY pressure. PX and PY vary with
changing engine operating conditions as well as inlet air temperature. PY pressure is applied to the
outside of the governor bellows, while PX pressure is applied to the inside of the bellows and the outside
of the acceleration bellows. Any change in PY will therefore have more effect on the diaphragm than an
equal change in PX pressure due to the difference in effective area. When both pressures increase
simultaneously, as during acceleration, the bellows move downwards and the metering valve moves in
an opening direction. When PY decreases as the desired NG is approached, the bellows will travel to
© Solenta Aviation (Pty) Ltd
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
reduce the opening of the metering valve. When both pressures decrease simultaneously the bellows
will move upwards and reduce the metering valve opening because PY is more effective than PX. This
occurs during deceleration and moves the metering valve to the minimum flow stop.
Compressor Turbine (NG) Governor
The power lever incorporates a speed scheduling cam which depresses an internal rod when the power
lever is advanced. The governor lever is pivoted and one end operates against an orifice to form the
governor valve. The enrichment lever is also pivoted and actuates a fluted pin which operates against
the enrichment hat valve. The speed scheduling cam applies tension to the governor spring which
applies a force to close the governor valve. The enrichment spring between the enrichment and
governor levers provides a force to open the enrichment valve. As the drive shaft rotates, it in turn
rotates a table on which the governor flyweights are mounted. As the NG increases, centrifugal loading
causes the flyweights to move outwards. This in turn moves the NG platform upwards overcoming the
enrichment spring force, closing the enrichment valve and opening the governor valve.
Manual Override System
The retaining plate and cover containing the governor bellows stop are replaced by a shaft and stop
assembly. If operated, it pushes against the end of the governor bellows to open the metering valve and
increase the fuel flow.
Power Turbine (NF) Governor
In the event of a power turbine overspeed condition, a governing orifice in the NF governing section is
opened by flyweight action of the governor. PY pressure is bled off to the atmosphere. When this occurs,
PY pressure acting on the FCU governor bellows decreases and moves the metering valve in a closing
direction, thus reducing fuel flow. This in turn decreases NG speed and consequently NF speed.
x FCU Operation
Engine Starting
The starting cycle is initiated with the Power Lever placed in the IDLE position and the Fuel Condition
Lever in the CUTOFF position. The ignition and starter are switched on, and when the required Ng
speed is attained, the Fuel Condition Lever is placed in the LOW IDLE position. Following ignition, the
engine accelerates to idle speed. During the starting sequence, the metering valve in the FCU is in a low
flow position. The flow divider schedules the metered fuel from the FCU, between the primary and
secondary nozzles. Fuel is delivered to the combustion chamber through 10 primary and 4 secondary
nozzles. Metered fuel is delivered initially by the primary nozzles, with the secondary nozzles cutting in
above a certain value.
As the compressor accelerates, the discharge pressure (P3) increases. This creates an increase in PX
and PY pressure, which is modified P3 pressure. As PY pressure acts on a greater area, the bellows are
forced down causing the main metering valve to move in an opening direction. Excess fuel supplied by
the fuel pump will pass via the bypass valve back to the tank. When Ng approaches idle speed, the
centrifugal force of the NG governor flyweights begin to overcome the governor spring force and open
the governor valve, bleeding off PY pressure. This creates a PX/ PY differential which causes the
metering valve to move in a closing direction until the required idle speed fuel flow is obtained. Any
variation from the selected speed will be sensed by the NG governor flyweights and will result in an
increased or decreased weight force.
Acceleration
As the power lever is advanced above the idling position, the speed scheduling cam is repositioned
moving the cam follower lever to increase governor spring force. The governor spring then overcomes
the flyweights and moves the governor lever, closing the governor valve. PX and PY immediately
increase, causing the metering valve to move in an opening direction.
As NG and consequently NF increase, the propeller governor increases the pitch of the propeller blades
to control NF at the selected speed and applies the increased power as additional thrust. Acceleration is
complete when the centrifugal force of the governor flyweight again overcomes the governor spring and
opens the governor valve.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Governing
Once the acceleration cycle has been established, any variation in engine speed from the selected
speed will be sensed by the NG governor flyweights and will result in increased or decreased weight
force. This change in weight force will cause the governor valve to either open or close which will be
reflected by the change in fuel flow necessary to re-establish the selected speed.
Altitude Compensation
Altitude compensation is automatic since the acceleration bellows assembly in the FCU is evacuated
and affords an absolute pressure reference. Compressor discharge air (P3) is a measurement of engine
speed and density. PX is proportional to P3, so it will decrease with a decrease in air density. This is
sensed by the acceleration bellows which act to reduce fuel flow on acceleration at altitude.
Deceleration
When the Power Lever is retarded, the speed scheduling cam raised, reducing the governor spring force
and allows the governor valve to move in an opening direction. The resulting drop in PY pressure moves
the metering valve in a closing direction until it contacts the WF minimum flow stop. This stop ensures
sufficient metered fuel flow to the engine to prevent flameout. The engine continues to decelerate until
the governor flyweight force decreases to balance the governor spring force at the set position.
Reverse Thrust
Reverse thrust can be obtained at any propeller speed. When the Power Lever is moved to the FULL
REVERSE position it will increase compressor turbine speed (NG) and propeller reverse pitch. The
propeller governor is maintained in an underspeeding condition in the reverse thrust range by controlling
propeller speed with the NF governing section of the propeller control.
If NF exceeds the desired speed the NF governing orifice will open to decrease PY pressure in the
computing section of the FCU and cause a reduction in fuel flow and NF speed, thereby limiting the
propeller speed and maintaining the CSU in an underspeed condition.
Engine Shutdown
The engine is shut down by moving the Fuel Condition Lever to the CUTOFF position. Fuel is returned
to the pump inlet via the bypass passages. Fuel in the primary and secondary manifolds is drained via
the dump valve ports in the flow divider and dump valve. Residual fuel is allowed to drain into the fuel
drain can.
Lubrication System
The PT6A engine lubrication system functions primarily to cool and lubricate engine bearings and
bushings. It also provides oil to the propeller governor and propeller reversing control system. The
main oil tank houses a gear-type engine-driven pressure pump, an oil pressure regulator, a cold
pressure relief valve and an oil filter. The engine oil tank, an integral part of the compressor inlet case,
is located in front of the accessory gearbox.
Oil is drawn from the bottom of the tank through a filter screen where it passes through a pressure relief
valve for regulation of the pressure. The pressure oil is then delivered from the main oil pump to the oil
filter. Pressure oil is then routed to the number 1 bearing and accessory gearbox, bearings number 2, 3,
4, reduction gears, torquemeter and propeller governor. Pressure oil is also routed to the oil-fuel heater
where it then returns to the tank.
After cooling and lubricating all the moving parts, oil is scavenged as follows:
x Oil from the number 1 bearing sump to accessory gearbox by gravity.
x Oil from the number 2 bearing sump is pumped to the accessory gearbox by number 2 scavenge
pump.
x Oil from the number 3 and 4 bearing sump is pumped to the accessory gearbox by the power
turbine bearing scavenge pump.
x Oil from the propeller governor, front thrust bearing, reduction gear accessory drives and
torquemeter is pumped to the tank by the reduction gearbox scavenge pump through a
thermostatically controlled air-oil cooler and then returned to the tank.
© Solenta Aviation (Pty) Ltd
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Figure 5-11 PT6A-114A Lubrication System
© Solenta Aviation (Pty) Ltd
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
x
The rear element of the internal scavenge pump scavenges oil from the accessory gearbox and
routes it through the air-oil cooler where it then returns to the oil tank.
The presence of pressurised air in the bearing cavities is a result of gas path air leaking across carbon
and oil seals. This air pressure assists in oil return to the tank by putting a head of pressure on the
scavenge oil at the bearing sumps. Breather air from the engine bearing compartments and from the
accessory and reduction gearboxes is vented overboard through a centrifugal breather installed in the
accessory gearbox. The bearing compartments are connected to the accessory gearbox by cored
passages and existing scavenge return lines.
The oil tank capacity is 9.5 US quarts and the total system capacity is 14. US quarts (including oil in
filter, cooler and hoses). An oil dipstick indicates US quarts low when the engine is hot. Fill to within 1½
qts of max hot or max cold positions. To obtain an accurate oil level reading, it is recommended the oil
level be checked within 10 minutes after shutdown when oil is hot (MAX HOT marking) or prior to the
first flight of the day (MAX COLD marking). If more than 10 minutes has elapsed and oil is still warm
perform an engine dry motoring run before checking the oil level.
As the oil tank is pressurised at 3 to 6 psi, ensure that the oil dipstick is securely latched down.
Operating the engine with less than the recommended oil level and with the dipstick cap unlatched will
result in excessive oil loss and eventual engine stoppage.
Engine Controls
Power Lever
The power lever is connected through linkage to a cam
assembly mounted in front of the FCU. It controls engine
power, via pneumatic control of the metering valve, from
maximum take off power back through idle to full reverse
thrust. The power lever has MAX, IDLE, BETA and
REVERSE range positions. The lever selects propeller
pitch when in beta range. The beta range enables the
pilot to control propeller blade pitch from idle thrust to
through a zero or no thrust condition to maximum
reverse thrust.
Do not move the power lever into the beta range when
the propeller is feathered as the propeller reversing
linkage will be damaged.
Figure 5-11 Engine Controls
Emergency Power Lever
The emergency power lever is connected through linkage to the manual override lever on the FCU. It
governs fuel supply to the engine should a pneumatic failure in the FCU occur. A pneumatic failure will
result in the fuel flow decreasing to idle (48% NG at sea level and 52% NG at 5000 ft).
The emergency power lever has NORMAL, IDLE and MAX positions. The NORMAL position is used for
all normal engine operation when the FCU is operating normally and power is selected by the power
lever. The range from IDLE to MAX positions governs engine power and is used when a pneumatic
malfunction has occurred in the FCU and the power lever is ineffective. A mechanical stop in the lever
slot requires that the emergency power lever be moved to the left to clear the stop before it can be
moved from the NORMAL position.
The knob has cross hatching and is visible when the lever is in the MAX position. Also, the emergency
power lever is annunciated on the annunciator panel when the lever is out of the normal position. These
precautions are intended to preclude starting of the engine with the emergency power lever in any
position other than NORMAL.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
The emergency power lever and its associated manual override system is considered to be an
emergency system and should only be used in the event of a fuel control pneumatic malfunction
For starting ensure that the emergency power lever is in the normal position otherwise an over
temperature or hot start may result.
When using this lever in the override position engine response may be more rapid than when using the
power lever as the fuel is not metered. Utilise slow and smooth movement of the emergency power lever
to avoid engine surges and/or exceeding ITT, NG and torque limits. The emergency power lever may
have a dead band such that no engine response is observed during the initial forward travel from the
IDLE position. When using the emergency power lever, monitor gas generator RPM when reducing
power near idle, to keep it from decreasing below 65% NG in flight.
Propeller Control Lever
The propeller control lever is connected to the propeller control governor mounted on the front section of
the engine. It controls propeller governor settings from maximum RPM to full feather.
The propeller control lever has MAX, MIN and FEATHER positions. The MAX position is used when
high RPM is desired and governs the propeller speed at 1900 RPM. Propeller control lever settings from
the MAX position to MIN permit the pilot to select desired propeller RPM for the cruise. The FEATHER
position is used for shutdown to stop rotation of the power turbine and the front section of the engine.
Once the gas generator is shut down no lubrication is available and rotation of the forward section is not
desirable. Also, feathering the propeller when the engine is shut down minimises propeller windmilling
during windy conditions. A mechanical stop in the lever slot requires that the propeller control lever be
moved to the left to clear the stop before it can be moved into or out of the FEATHER position.
The propeller lever operates a speeder spring inside the primary governor to reposition the pilot valve,
which results in an increase or decrease of propeller rpm. For propeller feathering, the propeller control
lever lifts the pilot valve to a position that causes complete dumping of high pressure oil, allowing the
counterweights and feathering spring to change the pitch.
Fuel Condition Lever
The fuel condition lever is connected through linkage to a combined lever and stop mechanism on the
FCU. The lever and stop also function as an idle stop for the FCU rod. It controls the minimum RPM of
the gas generator turbine (NG) when the power lever is in the IDLE position.
The fuel condition lever has CUTOFF, LOW IDLE and HIGH IDLE positions. The CUTOFF position cuts
all fuel to the fuel nozzles. LOW IDLE positions the control rod stop to provide an RPM of 52% NG. HIGH
IDLE provides an RPM of 65% NG.
Quadrant Friction Lock
A quadrant friction lock, located on the right side of the pedestal, is provided to minimise creeping of the
engine controls once they have been set. The lock increases the friction on the engine controls when
rotated clockwise.
Engine Instruments
Figure 5-12 C208B Engine Instruments
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Torque Indicator
The torque indicator indicates the torque being produced by the engine. One some Super Cargomaster
aircrafts the torque-indicator system is powered by 28-volt DC power. On the other cargo versions and
passenger versions, the torque indicator is pressure actuated.
Two independent lines enter the back of the torque indicator. One line measures engine torque pressure
and the other the reduction gearbox internal pressure. The torque indicator monitors the engine torque
pressure and converts this pressure into an indication of torque in ft-lbs.
Normal operating range (green arc) is from 0 to 1865 ft-lbs. The alternate power range (striped green
arc) is from 1865 ft-lbs to 1970 ft-lbs. The maximum torque (red line) is 1970 ft-lbs and maximum take
off torque (red wedge) is denoted by "T.O" at 1865 ft-lbs.
Propeller RPM Indicator
The instrument is marked in increments of 50 RPM and indicates propeller speed in revolutions per
minute. The instrument is electrically operated from the propeller tachometer generator which is
mounted on the right side of the front case. Normal operating range (green arc) is from 1600 to 1900
rpm. The maximum (red line) is 1900 rpm.
ITT Indicator
The ITT (interturbine temperature) indicator displays the gas temperature between the compressor and
the power turbines. Normal range (green arc) is 100°C to 740°C. Engine operations which exceed
740°C may reduce engine life. A maximum (red line) at 805°C. Maximum starting temperature (red
triangle) is 1090°C. These limitations are restricted to 2 seconds. Although not indicated, the idle ITT is
restricted to 685°C.
NG% RPM Indicator
The NG% indicator indicates % of gas generator rpm based on a figure of 100% at 37,500 RPM. It is
electrically operated from the gas generator tachometer-generator mounted on the lower right-hand side
portion of the accessory case. The normal operating range (green arc) is from 52% to 101.6% NG. The
maximum (red line) is at 101.6% NG.
Fuel Flow Indicator
The fuel flow indicator indicates fuel flow in lbs/hr based on Jet A fuel. Fuel is measured downstream of
FCU just before it enters the flow divider. When the power is removed from the indicator, the needle will
stow below the zero in the OFF band. The instrument is electrically operated.
Oil Pressure Indicator
A direct pressure oil line from the engine delivers oil at operating pressure to the indicator. Minimum
pressure (red line) is 40 psi. Normal operating range (green arc) is from 85 to 105 psi. The maximum
pressure (red line) is at 105 psi.
Normal oil pressure is 85 to 105 psi at gas generator speeds above 72% with the oil temperature
between 60° and 70°C. Oil pressures below 85 psi are undesirable and should be tolerated only for the
completion of the flight, preferable at a reduced power setting. Oil pressures below 40 psi are unsafe
and require that either the engine be shut down or a landing be made as soon as possible using
minimum power as required to sustain flight.
Oil Temperature Indicator
The instrument is operated by an electrical-resistance type temperature sensor which receives power
from the aeroplane electrical system. Minimum temperature (red line) is -40°C. Normal operating range
(green arc) is from 10° to 99°C. Maximum operating temperature (red line) is 99°C.
© Solenta Aviation (Pty) Ltd
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Engine Limitations
Airplane and engine limits are described in the Limitations section of the AFM. These limitations have
been approved by the FAA, and must be observed when operating the C208B. The following engine
operating limits chart provides important limitations for all operating conditions.
OPERATING CONDITIONS FOR C208B INSTALLATION
MAXIMUM
ITT
°C
NG RPM
% NG
(2)
PROP.
RPM
OIL
PRESS.
PSIG
(3)
OIL
TEMP.
°C
(7)
1865
(1)
1865
1970 (4)
1865
1970 (4)
805
(10)
101.6
1900
85 TO 105
10 TO 99
765
101.6
1900
85 TO 105
0 TO 99
740
101.6
1900
85 TO 105
0 TO 99
----
----
685
52
----
40
-40 TO 99
675
1865
805
101.6
1825
85 TO 105
0 TO 99
----
2400
(6)
850
102.6
2090
----
0 TO 99
----
----
1090
----
----
----
675
1865
805
101.6
1900
85 TO 105
POWER
SETTING
sHP
(9)
TAKEOFF
675
MAXIMUM
CLIMB
MAXIMUM
CRUISE
IDLE
MAXIMUM
REVERSE
(5)
TRANSIENT
(11)
STARTING
(11)
MAXIMUM
CONTINUOUS
(8)
675
675
TORQUE
FT-LBS
(Minimum)
(Minimum)
-40
(Minimum)
10 TO 99
1. Maximum allowable torque decreases with altitude and temperature, as per the Engine
Torque For Takeoff figure of AFM Section 5 Figure 5-8.
2. For Every 10ºC below -30ºC OAT, reduce the maximum allowable NG by 2.2%
3. Normal Oil Pressure is 85 – 135 psi at NG speeds above 72% with oil temperature
between 60º – 70ºC. Oil pressures below 85 psi are undesirable, and should be tolerated
only for the completion of the flight, preferably at a reduced power setting. Oil pressures
below normal should be reported as an engine discrepancy and should be corrected
before the next flight. Oil pressures below 40 psi are unsafe, and require that either the
engine be shut down, or a landing be made as soon as possible using the minimum
power possible to sustain flight.
4. Propeller RPM must be set so as not to exceed 675 sHP with torque above 1865 ft-lbs.
Full 675 sHP rating is available only at RPM setting of 1800 or greater.
5. Reverse power operation limited to 1 minute.
6. Time limited to 20 seconds.
7. For increased oil service life, an oil temperature between 74ºC – 80ºC is recommended.
A minimum oil temperature of 55ºC is recommended for fuel heater operation at takeoff.
8. Use of this rating is intended for abnormal situations (e.g. extreme icing, windshear, etc).
9. The maximum allowable sHP is 675. Less than 675 sHP is available under certain
temperature and altitude conditions, as reflected in the takeoff, climb and cruise
performance charts.
10. When the ITT exceed 765ºC, this power setting is time limited to 5 minutes.
11. Time limited to 2 seconds.
During engine start, temperature is the most critical limit. The ITT starting limit of 1090ºC is limited to
2 seconds. During any start, if the indicator needle approaches this limit, the start should be aborted
before the needle approaches the red triangle.
© Solenta Aviation (Pty) Ltd
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Starter Limitations
Engine starters are time-limited during the starting cycle to prevent the possibility of starter damage
due to overheating. The time limitations on the starter vary depending on whether the start is being
attempted using aircraft battery power, or external ground power.
Aircraft Battery Start Cycle ................................ 30 seconds ON - 60 seconds OFF,
30 seconds ON - 60 seconds OFF,
30 seconds ON - 30 MINUTES OFF.
External Power Start Cycle ............................... 20 seconds ON - 120 seconds OFF,
20 seconds ON - 120 seconds OFF,
20 seconds ON - 60 MINUTES OFF.
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PROPELLER SYSTEMS
Table of Contents
Propeller System......................................... 6-1
Blade Angle................................................. 6-2
Propeller Governor...................................... 6-2
Propeller Governor Operation ..................... 6-2
Low Pitch Stop ............................................ 6-3
Beta And Reverse Control .......................... 6-4
Beta And Reverse Control Operation ..........6-4
Propeller Overspeed Governor....................6-5
Power Turbine (NF) Overspeed Governor ...6-5
Power Lever.................................................6-7
Propeller Control Lever................................6-7
Feathering....................................................6-7
Propeller System
The Cessna 208B is equipped with a constant speed, full-feathering, reversible, single-action, governor
controlled, three bladed propeller, mounted on the output shaft of the reduction gearbox. Since the
engine is a free turbine, with no mechanical connection between compressor and power turbines, the
propeller can rotate freely on the power shaft when the engine is shut down. Propeller bungees and
engine intake plugs are provided to prevent windmilling at zero oil pressure when the airplane is
parked. The propeller is either a McCauley aluminium propeller or a Hartzell composite propeller.
Figure 6-1 McCauley Prop
Figure 6-2 Hartzell Prop
Propeller pitch and speed are controlled by engine oil pressure supplied to the propeller dome through
engine-driven propeller governors. A governor oil pump boosts oil pressure delivered by the engine oil
system to a pressure high enough to control movement of the propeller blades. When oil pressure is
present in the propeller dome, propeller pitch (blade angle) is controlled normally by the propeller
governor or by the beta valve, depending upon the propeller’s mode of operation. As oil pressure
increases, the propeller moves toward low pitch (high RPM). Loss of oil pressure will cause centrifugal
counterweights and feathering springs to move propeller blades toward high pitch (low RPM). As oil
pressure decreases during engine shutdown, the propeller automatically moves toward feather.
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The minimum low pitch propeller position is determined by a mechanically-actuated hydraulic stop,
referred to as the primary low pitch stop. The power lever controls beta and reverse blade angles by
adjusting the low pitch stop position in beta and reverse ranges.
Two governors (a propeller governor and a propeller overspeed governor) control propeller RPM. The
primary governor controls the propeller through its normal governing range. The propeller control lever
selects propeller RPM by adjusting the governor condition. Should the primary governor malfunction,
the propeller overspeed governor prevents propeller speed from exceeding approximately 2000 RPM.
The power turbine (NF) overspeed governor acts as a backup governor limiting propeller speed to
106% of that selected by the propeller lever. In the reverse range, the power turbine (NF) overspeed
governor is reset, limiting propeller RPM to approximately 95% of the governor setting. The power
turbine (NF) overspeed governor limits propeller RPM by reducing fuel flow to the engine.
Blade Angle
Blade angle is the angle between the chord of the
propeller and the propeller’s plane of rotation. Because of
the normal twist of the propeller, blade angle is different
near the hub than it is near the tip. Blade angle for the
C208B is measured at the chord, 42 inches from the
propeller’s hub for a Hartzell Propeller, and at 30 inches
for a McCauley propeller.
Propeller Governor
Figure 6-3 Blade Angles
Also known as the Primary Governor on some other PT6 installations.
The propeller governor is located in the 12 o'clock position on the front case of the reduction gearbox. It
combines the functions of a normal propeller governor (CSU), a reversing (beta) valve and a power
turbine (NF) governor into a single unit.
Under normal flight conditions, the governor acts as a CSU by maintaining propeller speed selected by
the pilot. If the Power Lever is pushed forward, the FCU simply schedules an increase in fuel flow. Thus,
effectively more energy is available to turn the turbine. The turbine is forced to absorb the extra energy
that is transmitted to the propeller in the form of torque. If there were no governor, the propeller speed
would increase. Instead the blade angle increases to maintain a constant propeller RPM and the
propeller is allowed to take a larger bite of air, hence the increase in torque.
Therefore, the governor varies the propeller blade pitch angle to match the load to the engine torque in
response to changing flight conditions. A spring-loaded pilot valve is installed in a driveshaft centrebore
within the propeller governor. Ports in the driveshaft and the position of the pilot valve in the shaft,
control the direction of oil flow within the housing. The rotating shaft with the rotating flyweights
determine the position of the pilot valve while opposing spring load on the valve is varied by the speed
adjusting lever at the head of the governor. The speed adjusting lever is connected to the Propeller
Control Lever in the cockpit. The propeller governor can maintain any selected propeller speed from
approximately 1600 RPM to 1900 RPM.
Propeller Governor Operation
The primary governor modulates oil pressure in the propeller dome to change blade angle to maintain
a constant propeller speed. As oil pressure in the dome changes, propeller blade angles change to
maintain the propeller speed the pilot has selected.
For example, suppose an airplane is in normal cruising flight with the propellers set at 1750 RPM. If
the pilot begins a descent without changing power, the airspeed will increase. This decreases the
angle of attack of the propeller blades, causing less drag on the blades, thus causing the RPM to
increase. The governor will sense this overspeed condition and increase blade angle to a higher pitch.
The higher pitch increases the blade’s angle of attack, slowing it back to an on-speed 1750 RPM.
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Cessna 208B Grand Caravan
Figure 6-4 Propeller Governor Operation
Likewise if an airplane changes from cruise to climb attitude without a power change, the propeller
RPM tends to decrease. The governor responds to this under-speed condition by decreasing blade
angle to a lower pitch, and the RPM returns to its original value. Thus the governor is able to keep the
variable pitch propeller at a constant speed.
Power changes, as well as airspeed changes, cause the propeller to momentarily experience
overspeed and under-speed conditions, but again the governor reacts to maintain the on-speed
condition. Due to the smooth action of the governor, the pilot will notice few, if any, of these minor
adjustments.
There are times, however, when the primary governor is incapable of maintaining the selected RPM.
For example, imagine an airplane approaching to land with its propellers set at 1900 RPM. As power
and airspeed are both reduced, under-speed conditions exist which cause the governor to decrease
blade angle as it attempts to restore the on-speed condition. If blade angle were allowed to decrease
to its full reverse limit, aircraft control would be dramatically reduced as the propeller blades moved
into a high drag disking position.
To prevent this undesirable situation, a device is provided to stop the governor from selecting blade
angles that are too low for safety. As blade angle is decreased by the governor, eventually the low
pitch stop is reached. Blade angle then becomes fixed, preventing its continued movement toward a
lower pitch. At the low pitch stop, the governor is prevented from restoring the on-speed condition,
and propeller RPM decreases below the selected governor RPM setting. Once the low pitch stop is
reached, blade angle cannot decrease further until the pilot selects beta or reverse.
Low Pitch Stop
With a non-reversing propeller the low pitch stop is simply at the low pitch limit of travel, determined
by the propeller’s construction. But with a reversing propeller, extreme travel in the low pitch direction
is past 0º, into reverse or negative blade angles. Consequently, the C208B’s propeller system has
been designed to allow the low pitch stop to be repositioned when reversing is desired.
The low pitch stop is created by mechanical linkage which senses blade angle. At the low pitch stop
the linkage causes a valve to close, stopping the flow of oil into the propeller dome. Since more oil
causes low pitch and reversing, blocking off oil flow creates a low pitch stop. The low pitch stop valve,
commonly referred to as the “beta” valve, is spring-loaded to provide redundancy in the event of
mechanical loss of beta valve control.
Low pitch stop operation is determined by a mechanically monitored hydraulic stop. The propeller
servo piston is connected by four spring-loaded sliding rods to the slip ring mounted behind the
propeller. A carbon brush block riding on the slip ring transfers the movement of the slip ring through
the propeller reversing lever to the beta valve of the governor. The initial forward motion of the beta
valve blocks off the flow of oil to the propeller. Further motion forward dumps the oil from the propeller
into the reduction gearbox sump. A mechanical stop limits the forward motion of the beta valve.
Rearward motion of the beta valve does not affect normal propeller control. When the propeller is
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
rotating at a speed lower than that selected on the governor, the governor pump provides oil pressure
to the servo piston and decreases the pitch of the propeller blades until the feedback of motion from
the slip ring pulls the beta valve into a position blocking the supply of oil to the propeller, thus
preventing further pitch changes.
Beta and Reverse Control
The position of the low pitch stop is controlled from the cockpit by the power lever. Whenever the
power lever is at IDLE or above, this stop is set at 15.6º on McCauley propellers, and 9.0º on Hartzell
propellers. Bringing the power lever aft of idle progressively repositions the low pitch stop to lower
blade angles.
The geometry of the power lever linkage through the cam box is such that power lever movement
from idle to full forward thrust has no effect on the beta valve’s position. When the power lever is
moved from idle to the reverse range, it repositions the beta valve to direct governor pressure to the
propeller piston, decreasing blade angle through zero into a negative range. The travel of the
propeller servo piston is fed back to the beta valve to null its position and, in effect, to provide infinite
negative blade angles all the way to maximum reverse. The opposite will occur when the power lever
is moved from full reverse to any forward position up to idle, thus providing the pilot with manual blade
angle control for ground handling.
The region of propeller travel between idle and ground fine is referred to as “beta”. In this range NG
remains at idle. To enter the beta range, the pistol-trigger catch on the power lever must be lifted, and
the lever brought aft past the idle stop. The aft stop of the beta range is called ground fine. With aft
movement of the power lever, blade angle moves progressively from idle to ground fine.
The region between ground fine and maximum reverse is referred to as the reverse range. In this
range, NG progressively increases while the blade angle decreases. To enter the reverse range the
power lever is moved aft of the beta range. Once the reverse range has been entered further aft
movement of the power lever will cause the blade angle to progressively decrease from ground fine to
maximum reverse.
Beta and Reverse Control Operation
Figure 6-5 Beta Range and Reverse Diagram
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
When the propeller blade angle approaches the low pitch stop setting, the four flanges extending from
the dome make contact with four beta nuts. As propeller pitch angle continues to decrease each
flange on the propeller dome pushes each beta nut and attached rod forward. As the rod moves
forward it pulls the feedback ring forward. In turn a beta valve inside the governor is pulled into the oil
cut-off position. The linkage is set to cut off oil supply to the dome when blade angle reaches 15.6º on
McCauley propellers, and 9.0º on Hartzell propellers. This provides the governor with a hydraulic low
pitch stop of for flight operations.
If the low pitch stop were fixed at the propeller could not enter the beta and reverse range; however,
the low pitch stop can be reset to allow the aircraft to enter into the beta and reverse ranges while the
aircraft is on the ground.
When the power lever is lifted back over the idle detent into the beta range, it is pulling back on the
top of the reverse lever. As the reverse lever moves back, the beta valve is pushed back, reestablishing oil flow to the propeller dome. This allows propeller blade angle to go below the low pitch
stop. As the propeller blades go below the low pitch stop, the propeller dome and feedback ring
continue forward, eventually pulling the beta valve back into the oil cut-off position.
Power Turbine (NF) Overspeed Governor
Also known as the Fuel Topping Governor on some other PT6 installations.
The primary propeller governor contains a power turbine (NF) overspeed governor which prevents
power turbine overspeed if a propeller malfunctions. An overspeed could occur, for example, if a
propeller blade were to stick in a fixed position during normal primary governor operation. In addition,
during reverse thrust operation, the fuel topping governor is set below the speed selected by the
primary governor to permit indirect control of propeller speed by the FCU servo system.
The speed at which the power turbine (NF) overspeed governor operation occurs is determined by the
speed selected with the propeller levers, and by the position of the reset lever. In the flight range the
reset lever is set to regulate power turbine, N2, speed at approximately 106% of the propeller lever
setting. In the ground range the lever is reset to 95% of the propeller lever position.
In the event of a turbine overspeed, an air bleed orifice in the propeller governor is opened by flyweight
action to bleed off compressor discharge pressure (PY) through the governor and computing section of
the FCU. Compressor discharge pressure, acting on the FCU governor bellows, decreases and moves
the metering valve in a closing direction. Fuel flow will be reduced, and engine power will decrease.
When this occurs, propeller RPM normally remains constant, but it may decrease if propeller blades
are frozen in a fixed position.
Propeller Overspeed Governor
Also known as the Overspeed Governor or Hydraulic Governor on some other PT6 installations.
Since the PT6’s propeller is driven by a free turbine (independent of the engine’s compressor),
overspeed can rapidly occur if the primary governor fails. The overspeed governor provides protection
against excessive propeller speed in the event of primary governor malfunction.
The propeller overspeed governor is installed in parallel with the propeller governor and is mounted at
the 10 o’clock position on the front case of the reduction gearbox. It is incorporated to control any
propeller overspeed condition by immediately bypassing pressure oil from the propeller servo to the
reduction gearbox sump.
When an engine overspeed occurs, the increased centrifugal force sensed by the flyweights overcomes
the spring tension, lifts the pilot valve and bypasses propeller pitch change mechanism oil back to the
reduction gearbox. This permits the combined forces of the counterweights and return springs to move
the blades toward a coarse pitch position, absorbing the engine power and preventing further engine
overspeed.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Figure 6-6 Overspeed Governor Operation
A solenoid valve, which resets the governor to a value below its
normal overspeed setting provides ground testing. The overspeed
governor test switch is located on the left side of the instrument panel.
When depressed, a solenoid is actuated on the propeller overspeed
governor. Speeder spring pressure is reduced and flyweights move
out at a lower speed to simulate an overspeed condition. Propeller
RPM is restricted when the power lever is advanced. During the test
the propeller RPM should not exceed 1750 ± 60 RPM.
Figure 6-7 Overspeed Governor Test Switch
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Power Lever
The power lever is located on the power lever quadrant on the control pedestal. It is mechanically
interconnected through a cam box to the FCU, reverse lever, beta valve and follow-up mechanism,
and the propeller governor. The power lever quadrant permits movement of the power lever in the
forward thrust (alpha) range from idle to maximum thrust, and in the beta/ reverse range from idle to
maximum reverse. Mechanical stops in the power lever quadrant at the IDLE positions prevent
inadvertent movement of the lever into the beta/ reverse range. To select beta or reverse, the pilot
must lift up catches on the power lever to allow the power lever to be brought back past the stops.
Figure 6-8 Control Pedestal
In the forward thrust (alpha) range the power lever establishes NG by selecting a gas generator
governor speed which results in a fuel flow that will maintain the selected NG RPM.
In the beta range, the power lever controls the beta valve to reduce propeller blade angle, thus
reducing residual propeller thrust. NG RPM is not affected in the beta range.
In the reverse range, the power lever:
• selects a blade angle proportionate to the aft travel of the lever
• selects a fuel flow that will sustain the selected reverse power
• resets the power turbine (NF) overspeed governor from its normal 106% to 95% of the propeller
governor setting
Therefore, RPM in reverse is a function of the primary propeller governor acting through the FCU to
limit fuel flow and control propeller RPM in relation to power lever position.
Propeller Control Lever
Propeller RPM, within the primary governor range of 1600 to 1900 RPM, is set by the position of the
propeller control lever. The full forward position sets the primary governor at 1900 RPM. In the full aft
position, forward of the feathering detent, the primary governor is set at 1600 RPM. A detent at the
low RPM position prevents inadvertent movement of the propeller lever into feather.
Propeller Feathering
The propeller can be manually feathered by moving the propeller lever full aft, past the detent, into
feather. This action locks the governor’s pilot valve in the full up position, opens the feather valve, and
all oil quickly drains from the propeller pitch mechanism. As oil is dumped from propeller servo
chambers, the counterweights and springs drive the propeller blades to the feather position. Since the
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
propeller shaft and the NG shaft are not connected, the propeller can be feathered with the engine
running; however to avoid excessive torque loads on the propeller gearbox, the engine should be at
idle power when propellers are manually feathered.
© Solenta Aviation (Pty) Ltd
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Dec-2005
FUEL SYSTEM
Table of Contents
Fuel System Operation ............................... 7-1
Firewall Shutoff Valve ................................. 7-3
Fuel Tank Selectors .................................... 7-3
Auxiliary Boost Pump Switch ...................... 7-3
Fuel Flow Indicator...................................... 7-4
Fuel Quantity Indicators...............................7-4
Annunciators ................................................7-4
Fuel Drain Valves.........................................7-5
Fuel Drain (EPA) Can ..................................7-5
Fuel Pump Drain Reservoir .........................7-6
Fuel System Operation
Two vented, integral fuel tanks with shutoff valves, a fuel selectors off warning system, a fuel reservoir,
an ejector fuel pump, an electric auxiliary boost pump, a reservoir manifold assembly, a firewall shutoff
valve, a fuel filter, an oil-to-fuel heater, an engine-driven fuel pump, a fuel control unit, a flow divider,
dual manifolds, 14 fuel nozzle assemblies and a fuel drain can and drain.
Fuel flows from the tanks through two shutoff valves at each tank. Fuel flows by gravity from the shutoff
valves in each tank to the fuel reservoir. The reservoir is located at the low point in the fuel system.
A head of fuel is maintained around the ejector boost pump and auxiliary boost pumps which are
contained within the reservoir. Fuel in the reservoir is pumped by the ejector boost pump or by the
electric auxiliary boost pump to the reservoir manifold assembly. The ejector boost pump is driven by
motive fuel from the FCU. It normally provides fuel flow when the engine is operating. In the event of
failure of the ejector boost pump, the electric auxiliary boost pump will automatically turn on, on
condition that the boost pump switch is in the NORM position. The auxiliary boost pump is also used to
supply fuel during starting.
Fuel in the reservoir manifold assembly flows through the fuel shutoff valves located on the aft side of
the firewall. Fuel is then routed through the fuel filter located on the front side of the firewall. The filter
incorporates a bypass feature.
Fuel is then routed through the oil-to-fuel heater to the engine driven pump through a 74 micron inlet
screen. The inlet screen is spring-loaded and should it become blocked, the increase in differential
pressure will overcome the spring and allow unfiltered fuel to flow into the pump chamber. The pump
increases the pressure and delivers it to the FCU via a 10-micron filter. A bypass valve and cored
passages in the pump body enables unfiltered high pressure fuel to flow to the FCU in the event the
outlet filter becomes blocked.
The FCU meters the fuel and directs it to the flow divider. The flow divider schedules the metered fuel
between the 10 primary and 4 secondary fuel manifolds. The result is an even, efficient spray pattern
through all operational speeds. In the primary spin chamber a change in direction puts a spinning motion
on the fuel and establishes the proper spray angle and helps with atomisation. At that point the fuel is
discharged from the primary nozzle. As the engine accelerates from start-up, the fuel pressure begins to
increase. At approximately 90 psi, it causes the flow divider to open and part of the incoming fuel flow is
channelled to the secondary spin chamber. From the nozzles the fuel is distributed to the combustion
chamber.
When the fuel cut-off valve in the FCU closes during engine shutdown, both primary and secondary fuel
nozzles are connected to a dump valve port. Residual fuel in the manifold drains into a fireproof can on
the front left side of the firewall. This can should be drained daily.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Figure 7-1 C208B Fuel System Schematic
Fuel venting is essential to system operation. A complete blockage of the vent system will result in
decreased fuel flow and eventual engine stoppage. Venting is accomplished by check valve equipped
vent lines from each tank. One vent from each tank protrudes from the trailing edge of the wing at the
tips. The fuel reservoir is also vented to both wing tanks.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Firewall Shutoff Valve
A manual shutoff valve is located on the aft side of the firewall. It enables the
pilot to shut off all fuel from the reservoir to the engine. The shutoff valve is
controlled by a red push-pull knob labelled FUEL SHUTOFF-PULL OFF is
located on the pedestal. The push-pull knob has a press-to-release button in
the centre which locks the knob in position when the button is released.
Figure 7-2 C208B Firewall Shutoff Valve
Fuel Tank Selectors
Two fuel tank selectors mechanically control the position of the fuel tank shutoff valves at each wing
tank. When in the ON position, both shutoff valves in the tank are open, allowing fuel from that tank to
flow to the reservoir. When the aircraft is parked on a slope, both fuel tank selectors should be in the
OFF position. By leaving the fuel selector of the elevated wing in the OFF position only will not
completely eliminate this cross flow. Over a prolonged period there is enough seepage through the
reservoir fuel vent to create a wing-heavy condition for the next flight.
Figure 7-3 C208B Fuel Tank Selectors
Auxiliary Boost Pump Switch
The auxiliary boost pump switch is located on the pilot’s side panel,
and is labelled FUEL BOOST and has OFF, NORM and ON positions.
OFF Position
The auxiliary boost pump is inoperative.
NORM Position
The auxiliary boost pump is armed and will operate if the fuel pressure
in the fuel manifold assembly drops below 4.75 psi. This switch
position is used for all normal engine operations where main fuel flow
is provided by the ejector boost pump and the auxiliary pump is used
as a standby.
Figure 7-4 C208B Auxiliary Boost Pump Switch
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ON Position
When the auxiliary boost pump switch is placed in the ON position, the pump operates continuously.
This position is used for the following:
• Engine starting
• Any time the auxiliary boost pump cycles on and off with the switch in the NORM position
• Conditions of near fuel exhaustion
• For all operations using Avgas
Fuel Flow Indicator
The fuel flow indicator indicates fuel consumption in pounds per hour based on Jet A fuel. Flow of fuel is
measured downstream of the FCU just before being routed to the fuel divider. When battery power is
removed, the indicator needle will stow below zero in the OFF position.
Figure 7-5 C208B Fuel System Gauges
Fuel Quantity Indicators
Fuel quantity is measured by four transmitters in each tank and indicated by two electrically operated
fuel quantity indicators. The indicators measure volume and are calibrated in pounds (based on the
weight of Jet A fuel on a standard day 6.7 lbs/gal at 15°C) and gallons. An empty tank is indicated by a
red line and the letter E. When showing empty, approximately 2½ gallons remain in the tank as
unusable fuel.
Because of the long fuel tanks, fuel quantity indicator accuracy is affected by uncoordinated flight or a
sloping ramp if reading the indicators while on the ground. To obtain an accurate reading, ensure that
the aircraft is parked laterally level, or if in flight, in a coordinated and stabilised condition.
Annunciators
Figure 7-6 C208B Fuel System Annunciators
AUX FUEL PUMP ON
An amber auxiliary fuel pump on annunciator labelled AUX FUEL PUMP ON will illuminate when the
auxiliary boost pump switch is in the ON position or when the boost pump switch is in the NORM
position and the fuel pressure in the fuel manifold assembly drops below 4.75 psi.
FUEL PRESS LOW
An amber fuel pressure low warning annunciator labelled FUEL PRESS LOW will illuminate when the
fuel pressure in the reservoir fuel manifold assembly is below 4.75 psi.
RESERVOIR FUEL LOW
A red reservoir fuel low annunciator labelled RESERVOIR FUEL LOW will illuminate when the level of
fuel in the reservoir drops to approximately one-half full. There is only enough fuel in the reservoir for
1½ minutes of engine operation maximum continuous power after illumination.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
LEFT And RIGHT FUEL LOW
Two amber fuel low warning annunciators are labelled LEFT FUEL LOW and RIGHT FUEL LOW.
Each annunciator will illuminate when the fuel in the respective tank is less than 25 gallons. (95 ltr’s/
approximately 146 lb’s)
FUEL SELECT OFF
A red fuel selectors off annunciator labelled FUEL SELECT OFF will illuminate when one or both fuel
selectors are OFF, the fuel selector warning circuit breaker is not set or the start control circuit
breaker is not set.
Fuel Drain Valves
Fuel drain valves are located at the lower surfaces of each wing at the inboard end of the tank.
Outboard fuel tank drain valves may be installed and their use is recommended if the aircraft is parked
with one wing low on a sloping ramp. The drain valves are constructed so that a Phillips screwdriver can
be used to depress the valve and then twist to lock it in the open position.
Figure 7-7 C208B Wing Fuel Tank, And Reservoir Tank Drains
The drain valve for the reservoir consists of a recessed T-handle. When pulled out, fuel from the
reservoir drains out the rear fuel drain pipe located adjacent to the drain valve. The drain valve for the
fuel filter consists of a drain pipe which can be depressed upwards to drain fuel from the filter. If
contamination is detected, drain all fuel drain points again. Take repeated samples of all fuel drain
points until all contamination has been removed.
Fuel Drain (EPA) Can
When the engine is shut down, residual fuel in the engine drains into a fuel drain can mounted on the
front left side of the firewall. The can should be drained once a day or at intervals not exceeding 6
shutdowns. A drain valve on the bottom side of the cowling is provided to drain the fuel into a suitable
container.
Figure 7-8 C208B Fuel Filter, EPA Can And Fuel Pump Drain
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Fuel Pump Drain Reservoir
To control expended lubrication oil from the engine fuel pump drive coupling area and provide a way
to determine if fuel is leaking past the fuel pump seal a drainable reservoir collects allowable
discharge of oil and any fuel seepage. The reservoir is mounted on the front left side of the firewall. It
should be drained once a day or at intervals not exceeding 6 shutdowns. A quantity of up to 3 cc of oil
and 20 cc of fuel discharge per hour of engine operation is allowable. In addition, a FCU bearing
failure will be indicated by a blue dye in the expended oil. This requires immediate attention and
under no circumstances should the aircraft be flown before this situation is corrected.
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Dec-2005
ELECTRICAL SYSTEM
Table of Contents
Introduction ................................................. 8-1
Electrical Buses........................................... 8-2
Starter-generator......................................... 8-2
Generator Control Unit ................................ 8-2
Ground Power Monitor ................................ 8-2
Ground service Plug Receptacle ................ 8-3
External Power Switch ................................ 8-3
Battery ......................................................... 8-3
Battery Switch ..............................................8-4
Generator Switch .........................................8-4
Avionics Power Switches.............................8-4
Avionics Bus Tie Switch...............................8-4
Circuit Breakers ...........................................8-5
Volt/Ammeter ...............................................8-5
Standby Electrical System ...........................8-5
Introduction
Figure 8-1 C208B Electrical Schematic
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
The aircraft is equipped with a 28 volt dc electrical system. The system uses a 24 volt battery as a
source of electrical energy and a dual purpose starter-generator. An optional standby alternator is
available for the use as a standby source in the event of the main generator system malfunctioning.
Figure 8-2 C208B Electrical System Annunciators
Electrical Buses
Power is supplied to the electrical and avionics circuits through 2 general, 2 avionics and a battery bus.
The battery bus is energised continuously for the memory keep alive, clock and cabin/courtesy lights.
The two general buses are on any time the battery switch is turned on. All DC buses are on any time the
battery switch and two avionics switches are turned on.
Starter-generator
The starter-generator is mounted on the engine accessory gearbox and is driven by the engine
through a splined shaft. It is cooled by an integral fan and a blast tube located above the right forward
cowling. It functions as a motor during engine start and as a generator after starting.
In the starter mode it is powered by either the battery or an external power source. After engine start, the
generator output is 28 volt and 200 amp. When operating as a generator, it supplies power to operate
the electrical systems and maintains the battery's state of charge. The unit is controlled by a generator
control unit.
Generator Control Unit (GCU)
The GCU is mounted inside the cabin on the left forward fuselage sidewall. It provides the electrical
control functions necessary for the operation of the starter-generator. It provides automatic starter cutout at 46% NG. Below 46% the starter-generator functions as a starter. Above 46% it functions as a
generator when the starter switch is OFF.
The GCU provides voltage regulation plus high voltage protection reverse current protection. A rheostat
increases the resistance in the field circuit and less current flows through the field winding resulting in a
decrease in the strength of the magnetic field. The rotating armature now turns in a weaker field with the
result of a lower generator output voltage.
A differential relay switch protects the generator from reverse current from battery voltage. It is
essentially an on/off switch that is controlled by the difference in voltage between the battery bus and
the generator output. The differential relay switch connects the generator to the electrical system’s main
bus whenever generator voltage exceeds bus voltage by at least 0.5 volt. If on the other hand, bus
voltage exceeds generator output voltage, the reverse current relay opens and takes the generator offline. In the event of the above, the generator is automatically disconnected from the buses and the red
GENERATOR OFF annunciator illuminates.
Ground Power Monitor
The ground power monitor is located inside the electrical power assembly mounted on the left side of
the firewall. This unit senses voltage applied to the external power receptacle and closes the external
power contactor when the applied voltage is within limits (24 - 28 volts and 800 - 1700 amp).
It also senses airplane bus voltage and will illuminate the VOLTAGE LOW annunciator when the bus
voltage drops to 24.5 volts.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Ground Service Plug Receptacle
The external power receptacle is installed on
the left side of the engine compartment near
the firewall. It permits the use of an external
power source for starting or maintenance.
Figure 8-3 C208B Ground Service Plug
Receptacle
External power control circuitry is provided to
prevent external power and the battery from
being connected together during starting. The
ground service circuit incorporates polarity
reversal and over-voltage protection.
External Power Switch
The external power is a guarded three-position switch.
OFF Position
Battery power is provided to the main bus and to the starter
circuit. External power cannot be applied to the main bus. With
the generator switch in the ON position, power is applied to the
generator control circuit.
STARTER Position
External power is applied to the starter circuit only. Battery
power is applied to the main bus. No generator power is
available in this position. If external power drops off during the
start cycle, the external power switch must be placed in the
OFF position to reconnect the battery to the starter if motoring is
required.
Figure 8-4 C208B Left Sidewall Panel
BUS Position
External power is applied to the main bus. No power is available to the starter. The battery, if desired,
can be connected to the main bus and the external power by the battery switch. Battery charge should
however be monitored to avoid overcharge.
Battery
The system uses a 24 volt lead-acid battery with a capacity of 45 amp-hr or a 24 volt Ni-cad battery with
a 40 amp-hr capacity. The Ni-cad battery has a longer service life, high discharge and short recharge
capability.
• Thermal Runaway
If for any reasons a cell’s temperature increases, its voltage and internal resistance decreases. Three
factors lead to increasing a cell’s temperature:
• Excessively high discharge rate.
• Excessively high ambient conditions, particularly with a poorly or improperly ventilated battery
compartment.
• Deterioration of the plates’ separator material, which allows oxygen from the positive plate to
interact with a negative plate where it will chemically interact with the cadmium and generate heat.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
The beginning of the problem is an excessive discharge rate. On short legs in IMC at night, the battery is
constantly being drained and recharged, which generates excessive battery temperature. Due to the
way individual cells are installed in the battery case, the outer cells dissipate heat through the sides of
the case and as a result run slightly cooler than the middle cells. As the battery temperature rises, it is
the inner cells that get the hottest.
The generator attempts to recharge the battery by supplying sufficient current to the cells. As the cell
temperature increases, its resistance decreases simply helping the generator. The excessive current
causes the cells to overheat even more, further reducing their resistance. The porous plastic strip
between the individual plates breaks down, further complicating the problem. At this point, shutting off
the generator will stop the problem.
Meanwhile, the inner cell’s internal resistance and voltage becomes so low that the voltage of the good
cells surrounding it will be higher by comparison and they will begin to feed the bad cells. When any cell
begins to receive voltage from a surrounding cell, isolating the battery from the generator will no longer
have any effect; the battery is self-destructing. The only remaining solution is to land as soon as
possible.
• Battery Hot Annunciator
This amber annunciator illuminates when the NiCad battery temperature is 140 - 160°C.
• Battery Overheat Annunciator
This red annunciator illuminates when the NiCad battery temperature exceeds 160°C. Immediately turn
the battery switch OFF.
Battery Switch
In the ON position battery power is supplied to the two general buses. The OFF position cuts power to
all buses except the battery bus.
Generator Switch
The generator has ON, RESET and TRIP positions.
RESET and TRIP
These positions are momentary and are spring loaded back to ON position. If the GENERATOR OFF or
VOLTAGE LOW lights illuminate, place the switch momentarily in the RESET position to restore
generator power. If erratic operation is observed, the system can be shut off by placing the switch
momentarily in the TRIP position. After a suitable waiting period, generator operation may be recycled
by placing the generator switch momentarily to RESET.
ON Position
The GCU will automatically control the generator line contactor for normal generator operation.
Avionics Power Switches
The avionics power switches control power to number 1 or number 2 avionics buses. They should be in
the OFF position prior to turning the battery switch ON or OFF, engine starting or applying external
power. All avionics may be turned on or off by operating the AVIONICS power switches rather than by
operating all of the individual avionics equipment switches.
Avionics Bus Tie Switch
The avionics bus tie switch is guarded in the OFF position. It connects the number 1 and number 2
avionics buses together in the event of failure of either bus feeder circuit. Since each avionics bus is
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
supplied power from a separate current limiter on the power distribution bus, failure of the current limiter
can cause failure of the affected bus. Placing the bus tie switch to the ON position will restore power to
the failed bus. Operation without both bus feeder circuits may require an avionics load reduction.
Circuit Breakers
Most of the electrical circuits are protected by “pull-off” type circuit beakers. The circuit breaker panel is
mounted on the left sidewall. Once a circuit breaker has popped, allow it to cool for approximately 3
minutes before resetting. A circuit breaker should be reset once only and at the discretion of the pilot
consistent with crew safety.
Ensure that all circuit breakers are in before all flights. Never operate with disengaged circuit breakers
without a thorough knowledge of the consequences.
Volt/Ammeter
A volt/ammeter and a four position rotary-type selector switch are mounted on the left side of the
instrument panel so that the electrical system can be monitored. The selector switch has GEN, ALT,
BATT and VOLT positions.
Figure 8-5 C208B Volt/Ammeter
GEN Position
Measures generator current. The meter is connected to the generator shunt and will display current in
amps flowing from the starter/generator to the distribution bus.
ALT Position
Current from the standby alternator to the number 1 and 2 bus power circuit breakers is displayed.
BATT Position
Battery charge rate or discharge is measured because the meter displays both positive and negative
currents.
VOLT Position
System voltage is measured from either the battery or the generator.
Standby Electrical System
A standby electrical system is installed in the event the main generator system malfunctions in flight.
This system includes an alternator at a 75 amp capacity rating. The alternator is belt driven from the
accessory pad on the rear of the engine.
Initially battery power is required to excite the magnetic field. Field excitation to the alternator control is
supplied through diode logic from either a circuit breaker in the standby alternator assembly or KEEP
ALIVE 2 circuit breaker in the main power relay box. A diode allows electron flow in one direction only.
After alternator operation is initiated, the alternator is self-excited.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
System monitoring is provided by two amber STBY ELEC PWR ON and STBY ELEC PWR INOP
annunciator lights. Supplied amperage can be monitored on the voltammeter by placing the switch to
ALT position.
The STBY PWR switch is turned on after start. The standby power system is now armed and will
automatically supply power to the main bus only if the system voltage drops below 27.5 volts.
In order to utilise the 75-ampere capacity from the standby alternator, the AVIONICS STBY PWR and
AVIONICS BUS TIE switch/breaker must be turned ON. AVIONICS 1 and 2 switches should be turned
OFF to avoid connecting the standby power system to a possible fault in the primary power when
operating on standby power. Power to the two main buses is limited to 30 amps per bus. The load may
have to be reduced to prevent battery discharge. In the event of a fault in the primary relay box, the
primary power supply system can be isolated by pulling the six 30 amp bus feeder circuit breakers.
Immediately after start, the STBY ELECT PWR ON annunciator may illuminate, indicating the standby
electrical system is operating to help replenish the electrical power used for starting. The STBY ELECT
PWR INOP annunciator will not illuminate except in the event of a broken alternator drive belt or an
electrical malfunction in the standby electrical system.
The standby alternator receives field current from the KEEP ALIVE 2 circuit. In an emergency the
standby alternator can be brought on line without turning the battery switch on. Normal engine shutdown
procedures call for the standby power to be switched off prior to shutting the engine down and turning
the battery switch off. If the standby power switch is inadvertently left on, several red lights in the
annunciator panel will remain illuminated after the battery is turned off.
© Solenta Aviation (Pty) Ltd
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Dec-2005
AIR-CONDITIONING & VENTILATION SYSTEMS
Table of Contents
Ventilating Air Inlets .................................... 9-1
Ventilation Fans .......................................... 9-1
Instrument Panel Vent Knobs ..................... 9-2
Cabin Heating System ................................ 9-2
Heating Controls ..........................................9-3
Air-conditioning System ...............................9-5
Air-conditioning Controls..............................9-6
Air-conditioner Operation.............................9-6
Ventilating Air Inlets
Ventilating air is obtained from an inlet on each side at the forward fuselage and through two ram air
inlets, one on each wing at the upper end of the struts. The wing inlet ventilating air is routed through the
wing into the plenum chamber located in the centre of the cabin top. The plenum chamber distributes
the air to the individual overhead outlets at each seat position.
Ventilation Fans (Non-Air-conditioner Equipped Aircraft)
An optional ventilation fan system may be installed to provide supplemental cabin ventilation. The
system controls are located on the overhead console and consists of two rotary-type vent air controls,
labelled VENT AIR.
Figure 9-1 C208B Ventilation Fan Controls
The vent air controls operate shutoff valves in the left and right wing to control the flow of ram ventilating
air which enters the inlet in each upper wing strut fairing. The vent air controls also operate switches in
two ventilation fan circuits to control fan operation. When the vent air controls are rotated to the CLOSE
position, the shutoff valves are closed. Rotating the knobs beyond the OPEN to FAN position, a
mechanism on the control actuates a switch to operate the duct-mounted fan just inboard of the shutoff
valve in each wing. The switch will not operate the fan until the shutoff valve is open, thus assuring a
supply of cool air to the fan motor. Whenever the vent air controls are in the OPEN position, ram airflow
is ducted to the overhead ventilating outlets. This airflow can be augmented by rotating the controls to
the FAN position.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
System electrical protection is provided by two “pull-off” type circuit breakers labelled LEFT VENT
BLWR and RIGHT VENT BLWR.
Instrument Panel Vent Knobs
Two vent knobs, labelled VENT, PULL ON, are located on each side of the instrument panel. Each knob
controls the flow of ventilating air from an outlet located adjacent to each knob. Pulling each knob opens
a small air door on the fuselage exterior.
Cabin Heating System
Hot compressor outlet air is routed from the engine through a flow control valve. This hot air is then
mixed with cabin return air or warm compressor bleed air in the mixer/muffler. Air is then routed to the
cabin distribution system. Controls direct the heated air to forward and/or aft portions of the cabin and to
the windshield for defrosting.
Figure 9-2 C208B Cabin Air System, Including Ventilation Fans
(For ease of display the Cabin Heater Outlets, and Cabin Overhead Ventilating Outlets have been removed
from the diagram)
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Cabin Heating Controls
The cabin heating controls are located at the bottom of the instrument panel, above the pedestal.
Figure 9-3 C208B Cabin Heating Controls
• Temperature Selector Knob
A rotary temperature selector knob, labelled TEMP, is located on the cabin heat switch and control
panel. This selector modulates the opening and closing action of the flow control valve to control the
amount and temperature of air flowing into the cabin. Clockwise rotation of the knob increases the mass
flow and temperature of the air.
If more cabin heat in needed on the ground, move the fuel condition lever to HIGH IDLE and/or select
the GRD position of the mixing air control. For best results turn the selector to the full clockwise position
and then slowly counter-clockwise to decrease the bleed airflow to the desired amount.
A temperature sensor, located in the outlet duct from the mixer/muffler operates in conjunction with the
temperature selector knob. In the event of an overheat in the outlet duct, a temperature sensor will be
energised, closing the flow control valve and shutting off the source of hot bleed air from the engine.
• Bleed Air Heat Switch
The BLEED AIR HEAT switch is controls the operation of the bleed air flow control valve. The ON
position of the switch opens the flow control valve, allowing hot bleed air to flow to the cabin heating
system. The OFF position closes the valve, shutting off flow of air to the bleed air heating system.
• Mixing Air Push-Pull Control
A push-pull control, labelled MIXING AIR, GRD-PULL, FLT-PUSH, is located on the cabin heat switch
and control panel.
GRD Position
With the push-pull control in the GRD position (pulled out), warm compressor bleed valve air (P2.5) is
mixed with hot compressor outlet air (P3) in the mixer/muffler. This mode is used for ground operation
when warm compressor bleed air (P2.5) is available. P2.5 is only available below 92% NG. It can be
used to augment hot compressor outlet bleed air (P3) during periods of cold ambient temperature.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
FLT Position
With the push-pull control in the FLT position (pushed in) cabin return air is mixed with hot compressor
outlet air (P3) in the mixer/muffler. This recirculation of cabin return air enables the heating system to
maintain the desired temperature.
If desired, the FLT position can be used on the ground when ambient temperatures are mild and
maximum heating is not required. In this mode, excess warm compressor bleed valve air (P2.5)
available at power settings below 92% NG, is exhausted overboard from the mixing air valve.
The mixing air push-pull control valve should always be in the FLT position in flight. Cabin return air
must be allowed to flow through the mixing valve and blend with hot compressor outlet air during high
engine power operation otherwise the system may overheat and cause automatic shutdown.
• Aft/Forward Cabin Push-Pull Control
A push-pull control, labelled AFT CABIN-PULL, FWD CABIN-PUSH, is located on the cabin heat switch
and control panel.
With the control in the AFT (pulled out) position, heated air is directed to aft cabin heater outlets located
in the floor directly behind the pilots in the Super Cargomaster version, and on the cabin sidewalls at
floor level for the Grand Caravan version.
With the control in the FWD CABIN position (pushed in), heated air is directed to the forward cabin
through four heater outlets located behind the instrument panel and/or two windshield defroster outlets.
The controls can be positioned at any intermediate setting desired for proper distribution of heated air to
the forward and aft cabin areas.
• Defrost/Forward Cabin Push-Pull Control
A push-pull control, labelled DEFROST-PULL, FWD CABIN-PUSH, is located on the cabin heat switch
and control panel.
With the control in the DEFROST position (pulled out), forward cabin air is directed to two defroster
outlets located at the base of the windshield. The aft/fwd push-pull control must also be pushed in for
availability of forward cabin air for defrosting.
With the control in the FWD CABIN position (pushed in), heated air will be directed to the four heater
outlets behind the instrument panel.
• Cabin Heat Firewall Shutoff Knob
A push-pull shutoff knob, labelled CABIN HEAT FIREWALL
SHUTOFF, PULL OFF, is located on the lower right side of the
pedestal. When pulled out, the knob actuates two firewall shutoff
knobs to the off position. One in the bleed-air supply line to the
cabin heating system and the other in the cabin return air line.
This knob should normally be pushed in unless a fire is
suspected in the engine compartment.
Figure 9-4 Cabin Heat Firewall Shutoff Knob
Do not place this knob in the OFF position when the mixing air
control is in the GRD position. This will result in a compressor
stall at low power settings when the compressor bleed valve is
open. The engine must be shut down to relieve back pressure
on the valves prior to opening the valves.
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Air-conditioning System
An optional air-conditioning system may be installed into Grand Caravan version aircraft. The airconditioning system is able to provide cooling air both on the ground and in flight.
A belt-driven compressor is located on the engine accessory section. There are three evaporator units
with integral blowers, one in each wing root area, and one in the tail cone behind the aft bulkhead.
The evaporator units direct cooled air to a series of overhead outlets in the cabin headliner. The
system condenser is mounted beneath the engine, and is provided with a louvered air inlet and outlet
on the lower left engine cowling. Refrigerant lines under the floorboards and in the fuselage sides
connect the compressor, evaporators and the condenser.
Figure 9-5 C208B Air-conditioning System
When the air-conditioning system is operating, cooled air is supplied to the cabin through 16 overhead
adjustable outlets, one above each pilot, 11 above the passengers, and 3 directing air forward from
the rear cabin bulkhead.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Air-conditioning Controls
The cabin heating controls are located at the bottom of the instrument panel, above the pedestal, next
to the cabin heat controls.
Figure 9-6 C208B Air-conditioning Controls
System electrical protection is provided by for 15 Amp circuit breakers, labelled AIR COND CONT,
LEFT VENT BLWR, RIGHT VENT BLWR, and AFT VENT BLWR.
• Air-conditioning Switch
A three-position master switch controls the air-conditioning system. Placing the switch in the COOL
position starts the system compressor and evaporator fans. With the switch in the VENTILATE position,
power is only provided to the system evaporator fans, providing uncooled ventilating air to the cabin.
• Air-conditioning Fan Switches
The 3 two-position AC Fan switches provide separate HIGH or LOW speed control of each evaporator
fan.
Air-conditioner Operation
As the air-conditioning compressor is belt-driven through the engine accessory section, significant
engine power loss may be evident with the engine at low power on the ground, and at take-off power.
Under very hot OAT, or high ground elevation conditions, the idle ITT may exceed the maximum idle
limitation of 685ºC. In this case, advance the fuel condition lever towards HIGH IDLE to increase the
NG as required to maintain the ITT within limits.
For increased cooling while on the ground, select the fan switches to HIGH, and increase NG to 6065% for a higher air-conditioning compressor RPM. Once the cabin has reached a comfortable
temperature the fan switches can be repositioned to LOW.
Ground operation of the air-conditioner with the propeller in beta range for prolonged periods will
cause the air-conditioning compressor pressure safety switch to disengage the compressor clutch,
and therefore should be avoided.
If the temperature of the air coming from the outlets does not start to cool within a minute or two, the
system may be malfunctioning, and should be turned off.
Due to the loss of engine power at take-off, it is recommended to place the air-conditioning master
switch to VENTILATE immediately before take-off, and to return it to COOL only once safely airborne
if air-conditioning is still required in flight. Likewise, before landing the air-conditioning master switch
should be placed to VENTILATE to ensure maximum power is available in the event of a go-around.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
During extended flight in high temperatures and humidity, the evaporator coils may frost over.
Normally the compressor cycles off when temperatures in the evaporators near freezing. If frost does
form, as evidenced by reduced cooling airflow, turn the air-conditioning master switch to VENTILATE
and select the fan switches to HIGH. This should increase evaporator discharge temperature
sufficiently to clear the frost.
There is a 10–25 fpm reduction in aircraft climb performance, 1-2 KTAS decrease in cruise
performance, and approximately 1% increase in fuel required for a given flight with the air-conditioner
operational.
© Solenta Aviation (Pty) Ltd
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Dec-2005
OXYGEN SYSTEM
Table of Contents
Oxygen System......................................... 10-1
Oxygen System
Oxygen is supplied by a 116.95 cubic foot capacity oxygen cylinder located in the fuselage tail cone.
The oxygen is under a pressure of 1850 psi. Below 200 psi flow rates are not predictable.
Cylinder pressure is reduced to 70 psi by pressure regulator attached to the cylinder. A shutoff valve
forms part of the regulator assembly. An altitude compensating regulator, located between the pressure
regulator and the oxygen supply lines, varies the flow of oxygen.
Partial re-breathing type oxygen masks are equipped with vinyl plastic hoses and flow indicators. Hoses
are high-flow type and colour coded with a blue band adjacent to the plug-in fitting.
Figure 10-1 C208B Oxygen Controls
Shutoff valves are located in the overhead consul. The valve shuts off the supply when supply not in
use. It is mechanically connected to the shutoff valve on the cylinder.
© Solenta Aviation (Pty) Ltd
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Dec-2005
ICE & RAIN PROTECTION SYSTEMS
Table of Contents
Introduction ............................................... 11-1
Pneumatic De-ice Boots............................ 11-1
Propeller Anti-ice Boots............................. 11-2
Windshield Anti-ice Panel ......................... 11-2
Pitot And Stall Warning Heat .....................11-3
Ice Detector Light.......................................11-3
Engine Inertial Separator ...........................11-3
Introduction
The flight into known icing equipment includes the following:
• Pneumatic de-icing boots on the wings, wing struts, horizontal and vertical stabiliser leading edges
• Electrically heated propeller blade anti-ice boots
• A detachable electric windshield anti-ice panel
• A pitot-static heat system and heated stall warning system
• An ice detector light
• Engine inertial separator
Figure 11-1 C208B Ice Protection Controls
All ice protection, with the exception of the manually activated engine inertial separator, are
electrically controlled.
Pneumatic De-ice Boots
The system components include a pressure line which leads from the engine bleed-air system pressure
regulator to the vacuum ejector, three flow control valves and pressure switches, a timer, a system
switch and circuit breaker, an annunciator and supply lines and pneumatically operated surface de-ice
boots.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Boots expand and contract using pressure from the engine bleed-air system to the flow control valves.
Normally the flow control valves are open to create a vacuum to hold the boots against the leading edge
surfaces. When a de-icing cycle is initiated, the control valves are closed in order to remove the vacuum
and allow the pressure to inflate the boots.
Figure 11-2 C208B Pneumatic De-ice Boots And Propeller De-ice Boots
When the switch is held in the AUTO (upper) position and released one de-ice cycle will be activated.
The switch is off when placed in the middle position. Boots inflate according to the following sequence:
• Horizontal and vertical stabilizer boots
• Inboard wing
• Outboard wing and wing strut boots
The total time required is 18 seconds and each sequence lasts 6 seconds. In the event of a malfunction
of the timer, causing erratic operation of a sequence of a cycle, the switch can be held momentary in
MANUAL (lower) position. This results in simultaneous inflation of all the boots. The system can be
stopped (deflating the boots) by pulling the circuit breaker labelled DE-ICE BOOT.
The pressure indicator annunciator illuminates within 3 seconds after initiating a cycle and remains on 3
additional seconds to the end of the first sequence. Through each of the remaining two sequences of
the cycle, the annunciator will remain off during pressure build-up for about 3 seconds and then
illuminate for about 3 seconds. When illuminated the pressure in the boot is about 15 psi. The absence
of illumination during any one of the three sequences of a cycle indicates insufficient pressure for proper
boot inflation and effective de-icing ability.
In rime ice conditions wait until approximately 1/2 to 3/4 inch of ice has accumulated. In clear ice
conditions wait until approximately 1/4 to 3/8 inch of ice has accumulated. This procedure is
recommended due to the high drag penalties associated with clear ice shapes. Do not cycle the boots
during the approach and landing since inflation may increase the stall speeds by much as 10 kts. Do not
use more than 20° flaps with heavy ice accumulations on the horizontal stabilizer leading edge.
Propeller Anti-ice Boots
The system is operated by a three-position toggle switch labelled PROP. In the AUTO position electric
current flows to an anti-ice timer which cycles the current simultaneously to the heating elements in the
anti-ice boots on the three propeller blades in intervals of 90 seconds on and 90 seconds off.
The MANUAL position can be used in the event of failure of the anti-ice timer. The switch can be held in
the MANUAL position to achieve emergency propeller anti-icing. Hold the switch in the lower position for
90 seconds.
The operation can be monitored on the ammeter labelled PROP ANTI-ICE AMPS. The ammeter should
indicate 20 to 24 amps. An oil-operated pressure switch installed in the electrical circuit prevents the
anti-ice system from being turned on without the engine running. A failure of this switch will be
undetected unless the ammeter is monitored continuously. The PROP ANTI-ICE CONT circuit breaker
protects the control circuit. The PROP ANTI-ICE circuit breaker protects the heater circuit.
Windshield Anti-ice Panel
A detachable, electrically heated windshield anti-ice panel can be mounted to the base of the pilot's
windshield utilising a spring-loaded quick-release pin. Windshield anti-icing is controlled by a threeposition toggle switch labelled W/S. In the AUTO position electric current regulated by a controller flows
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
to the anti-ice panel. In the event of malfunction in the system controller circuitry, the switch can be held
in the MANUAL position to achieve windshield anti-icing.
Figure 11-3 C208B Windshield Anti-ice Panel
Pitot And Stall Warning Heat
The pitot-static heat system consists of a heating element in the pitot-static tube, a two position toggle
switch, labelled PITOT/STATIC HEAT and a “pull-off” type circuit breaker. When the switch is turned on,
the element in the pitot-static tube is heated electrically to maintain proper operation in possible icing
conditions.
The stall warning vane and sensor unit is equipped with a heating element operated by the STALL
HEAT switch and is protected by the STALL WRN circuit breaker.
Ice Detector Light
An ice detector light is flush mounted near the upper left corner of the windshield, to illuminate the pilot’s
wing. The switch is spring-loaded to the off position and must be held in the ON position to illuminate the
light.
Engine Inertial Separator
The inertial separator is controlled by a T-handle located on the lower instrument panel. The handle is
labelled BYPASS-PULL, NORMAL-PUSH. The inertial separator should be in the BYPASS position
under the following conditions:
• running the engine on the ground or in flight in visible moisture with an OAT of 4°C or less
• ground operation on dusty, sandy fields
With the inertial separator in the BYPASS position and power set below the maximum torque limit,
decrease the maximum cruise torque by 100 foot-pounds. Do not exceed 740°C ITT. Fuel flow for a
given torque setting will be 15 pph higher.
To unlock, push slightly forward and rotate the handle 90° counter-clockwise. The handle can then be
pulled into the BYPASS position. Once moved into the BYPASS position, air loads on the moveable
vanes hold them in position. The handle is very stiff to move back into the NORMAL position due to
the airloads on the moveable vanes.
When using the inertial separator in icing conditions, ensure that the handle is only returned to the
NORMAL position after engine shutdown, due to the fact that ice build-up can still be present inside
the inertial separator even though external signs of icing have dissipated.
© Solenta Aviation (Pty) Ltd
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Dec-2005
INSTRUMENTS
Table of Contents
Pitot-Static System.................................... 12-1
Vacuum System ........................................ 12-2
Suction Gauge .......................................... 12-2
Vacuum Low Warning Annunciator ...........12-2
Electrically Operated Gyros .......................12-3
Pitot-Static System
The pitot-static system supplies ram air pressure to the airspeed indicator and static pressure to
airspeed indicator, vertical speed indicator and altimeter. The system is composed of a heated pitotstatic tube mounted on the leading edge of the left wing, a static pressure alternate source valve located
below the de-ice/anti-ice switch panel, a drain valve located on the left sidewall beneath the instrument
panel, an airspeed pressure switch located behind the instrument panel and associated plumbing.
The pitot-static heat system consists of a heating element in the pitot-static tube, a two position toggle
switch, labelled PITOT/STATIC HEAT and a “pull-off” type circuit breaker. When the switch is turned on,
the element in the pitot-static tube is heated electrically to maintain proper operation in possible icing
conditions.
Figure 12-1 C208B Static Controls
A static alternate source is installed below the de-ice/anti-ice switch panel and can be used if the static
source is malfunctioning. This valve supplies static pressure from inside the cabin instead of from the
pitot-static tube. If erroneous instrument readings are suspected due to water or ice in the pressure line
to the static pressure source, the alternate source valve should be pulled on. The pressure inside the
cabin will vary with vents open or closed. Cabin static is usually lower; causing instruments to overread.
A drain valve is located on the cabin sidewall below the instrument panel. The valve is used to drain
suspected moisture in the system by lifting the drain valve lever to the OPEN position. The valve must
be returned to the CLOSED position prior to flight.
An airspeed pressure switch in the pitot- static system is used to actuate an airspeed warning horn. The
horn is located behind the headliner in the area above the pilot. It sounds when the airspeed exceeds
175 KIAS. A warning will also be heard in the pilot’s headset.
A second, independent pitot-static system is included for the right flight instrument panel.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Vacuum System
A vacuum system provides the suction necessary to operate the left hand attitude indicator and the right
hand directional indicator. Vacuum is obtained by passing regulated compressor outlet bleed air through
a vacuum ejector. Bleed air flowing through an orifice in the ejector creates the suction to operate the
instruments. The vacuum system consists of the bleed air pressure regulator, a vacuum ejector on the
forward left side of the firewall, a vacuum relief valve and vacuum system air filter, vacuum operated
instruments, a suction gage and a vacuum-low warning annunciator.
Figure 12-2 C208B Typical Vacuum System
• Suction Gauge
The suction gage is calibrated in inches of mercury. The desired suction ranges are 4.5 to 5.5 up to 15
000 feet, 4.0 to 5.5 from 15 000 to 20 000 feet, 3.5 to 5.5 from 20 000 to 25 000 feet. The 15K, 20K, 25K
and 30K markings at the appropriate step locations indicate the altitude at which the lower limit of that
arc segment is acceptable.
A suction reading out of these ranges may indicate a system malfunction or improper adjustment. In this
case, the attitude and directional indicators should not be considered reliable.
• Vacuum Low Warning Annunciator
A red VACUUM LOW warning annunciator will illuminate when the suction is below 3.0 in. Hg.
© Solenta Aviation (Pty) Ltd
12-2
Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Electrically Operated Gyros
The right hand horizontal situation indicator and the left hand directional indicator have electrically
operated gyros. These instruments are energised anytime the battery switch is ON and the circuit
breakers are in. If takeoff is soon after engine start, cage the gyro immediately after start by pulling the
knob for approximately 5 seconds before releasing it. If time between start and takeoff is 10 minutes or
more, allow the gyro to run as caging is not necessary.
© Solenta Aviation (Pty) Ltd
12-3
Dec-2005
AVIONICS
Table of Contents
King KFC-150 Autopilot System ............... 13-1
Definitions And Terms............................... 13-2
KC-192 Autopilot And FD Computer......... 13-2
KA-185 Remote Mode Annunciator .......... 13-4
Autopilot Control Yoke Switches............... 13-5
Before Take-off Reliability Tests ............... 13-5
Autopilot Limitations .................................. 13-6
Flight Command Indicator (FCI) ................13-6
Horizontal Situation Indicator (HSI) ...........13-7
Slaving Accessory And Compensator .......13-8
Weather Radar...........................................13-9
Radar Safety Precautions........................13-10
Operational Procedures...........................13-10
King KFC-150 Autopilot System
The 150 AFCS is certified as a 2 axis autopilot control system. A third axis autopilot control for yaw is
available as an option. The system incorporates an electric pitch trim system which provides autotrim
during autopilot operation and manual electric trim for the pilot. Trim faults are visually and aurally
annunciated. A lockout device prevents autopilot engagement until the system has been successfully
preflight tested.
The following conditions will cause the autopilot to automatically disengage:
• Electrical power failure.
• Internal flight control system failure.
• A loss of a valid compass signal (HDG flag) when a mode using heading information is engaged.
With the HDG flag present the autopilot may be reengaged in the basic wings-level mode with any
vertical mode.
• Roll rates in excess of 14° per second except when the CWS switch is held depressed.
• Pitch rates in excess of 6° per second except when the CWS button is held depressed.
• Actuating the manual electric trim.
Figure 13-1 C208B KFC-150 Autopilot System Components
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Definitions And Terms
Autopilot
This autopilot system incorporates electrically-driven actuators that fly the aircraft by moving the control
surfaces which are the ailerons, rudder (when optional yaw damper is installed), elevators and the
elevator trim.
Flight Director (FD)
The flight director system incorporates a panel-mounted computer that calculates intercept angles and
displays them to the pilot as pitch and steering recommendations on the Flight Command Indicator
(FCI). The recommended pitch and steering display serves as a reminder to the pilot as to which way he
should fly the aircraft to get the desired results of his mode selector input.
The flight director may be used with the autopilot engaged or disengaged. In the latter case, the pilot
manually flies the aircraft to satisfy the command bar on the FCI which is positioned by the computer
rather than allowing the autopilot to satisfy the computed commands.
•
•
•
The autopilot can only be coupled to the NAV 1 receiver.
The flight director (FD) mode must be selected before the autopilot engage mode (AP ENG) can be
selected.
The AVIONICS POWER #1 switch supplies power to the autopilot and ELEV TRIM circuit breakers.
KC-192 Autopilot And Flight Director Computer
The autopilot system is controlled by the Autopilot and FD computer, located at the lower right side of
the avionics stack.
Figure 13-2 KC-192 Autopilot And Flight Director Computer
• Controls And Indicators
YD – Yaw Damper Annunciator
Illuminates when the yaw damper (optional) is engaged.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Mode Annunciators
FD – Flight Director Mode
ALT – Altitude Hold Mode
HDG – Heading Mode
NAV – Navigation Mode
APR – Approach Mode
BC – Back Course Mode
Illuminate when a mode is selected by the corresponding mode selector button.
GS – Glide Slope Annunciator
Illuminates continuously when the autopilot is coupled to the glide slope signal. The GC annunciator will
flash if the glide slope signal is lost (GS flag in CDI). The autopilot reverts to pitch hold operation. If a
valid glide slope signal returns within 6 seconds, the autopilot will automatically recouple in the GS
mode. If the valid signal does not return within 6 seconds, the autopilot will remain in pitch attitude hold
until such time that a valid glide slope returns and the aircraft passes through the glide slope. At that
point, GS couple will re-occur.
TRIM – Trim Warning Light
Illuminates continuously whenever trim power is not on or the system has not been preflight tested.
The TRIM warning light illuminates and an audible warning will sound whenever a trim fault is
detected. The autotrim system is monitored for the following failures:
• Trim servo running without a command.
• Trim servo not running when commanded to.
• Trim servo running in the wrong direction.
AP – Autopilot Annunciator
Illuminates continuously when the autopilot is engaged. Flashes 12 times whenever the autopilot is
disengaged (an aural alert will sound for 2 seconds).
AP ENG – Autopilot Engage Button
When pushed, engages the autopilot when all logic conditions are met.
TEST – Preflight Test Button
When momentarily pushed, initiates preflight test sequence which:
• Automatically turns on all lights.
• Tests the roll and pitch rate monitors.
• Tests the autotrim fault monitor.
• Checks the manual trim drive voltage.
• Tests all autopilot valid and dump logic.
• If the preflight is successfully passed, the AP annunciator light will flash for 6 seconds (an aural
tone will also sound simultaneously with the annunciator flashes).
• The autopilot cannot be engaged until the autopilot preflight tests are successfully passed.
BC – Back Course Approach Mode Selector Button
When pushed, selects Back Course Approach mode. This mode functions identically to the approach
mode except that response to the LOC signals is reversed. Glide slope coupling is inhibited.
APR –Approach Mode Selector Button
When pushed, selects Approach mode. This mode provides all angle intercept, automatic beam
capture and tracking of VOR, RNAV or LOC signals plus glide slope coupling in case of an ILS. The
tracking gain of the APR is greater than the gain in NAV mode. The APR annunciator will flash until
the automatic capture sequence is initiated. On the Remote Mode Annunciator, APR ARM will
annunciate until the automatic capture sequence is initiated. At beam capture, APR CPLD will
annunciate.
NAV – Navigation Mode Selector Button
When pushed, selects NAV mode. This mode provides all angle intercept (with HSI), automatic beam
capture and tracking of VOR, RNAV or LOC. The NAV annunciator will flash until the automatic
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
capture sequence is initiated. On the Remote Mode Annunciator, NAV ARM will annunciate until the
automatic capture sequence is initiated. At beam capture, NAV CPLD will annunciate.
HDG – Heading Mode Selector Button
When pushed, selects the heading mode which commands the aircraft to turn to and maintain the
heading selected by the heading bug on the HSI. A new heading may be selected at any time and will
result in the aircraft turning to the new heading with a maximum bank angle of about 20°. Selecting
HDG mode will cancel NAV, APR or BC track modes.
ALT – Altitude Hold Mode Selector Button
When pushed, selects altitude hold mode which commands the aircraft to maintain the pressure
altitude existing at the moment of selection. Engagement may be accomplished in the climb, cruise or
descent. In the APR mode, altitude hold will automatically disengage when the glide slope is captured.
FD – Flight Director Mode Selector Button
When pushed, selects the flight director mode, bringing the command bar in view on the AI and will
command wings-level and pitch attitude hold. The FD must be selected prior to autopilot engagement.
DN/UP – Vertical Trim Control
A spring-loaded-to-centre rocker switch will provide up or down pitch command changes. While in
ALT, will adjust altitude at a rate of 500 fpm. When not in ALT, will adjust pitch attitude at a rate of 0.7
deg/sec. Will cancel GS couple. The aircraft must pass through the glide slope again to allow GS
recouple.
YAW DAMP – Yaw Damper Switch
May be used to engage the optional yaw damper independently of the autopilot.
KA-185 Remote Mode Annunciator
The remote mode annunciator provides annunciators within the pilot’s normal instrument scan area,
as well as three marker beacon lights.
Figure 13-3 KA-185 Remote Mode Annunciator
ARM – Armed Annunciator
Illuminates continuously along with NAV or APR when either mode button is depressed. The ARM
annunciator will remain illuminated until the automatic capture sequence is initiated, at which time
ARM will extinguish and CPLD will illuminate.
CPLD – Coupled Annunciator
Illuminates continuously along with NAV or APR when either mode button is depressed. The ARM
annunciator will remain illuminated until the automatic capture sequence is initiated, at which time
ARM will extinguish and CPLD will illuminate.
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Autopilot Control Yoke Switches
Autopilot and electric trim controls are provided only on the captain’s control yoke.
Figure 13-4 Autopilot Control Yoke Switches
A/P DISC/TRIM INTER – Autopilot Disconnect And Trim Interrupt Switch
When depressed will disengage the autopilot and yaw damper and cancel all operating flight director
modes. When depressed and held, will do all above and also interrupt all trim power (stops trim
motion), disengage the autopilot and yaw damper and cancel all operating flight director modes.
Manual Electric Trim Control Switches
The left half of the switch provides power to engage the trim servo clutch. The right half controls the
direction of motion of the trim servomotor. In order to operate the electric trim, both switches must be
selected simultaneously. When the autopilot is engaged, operation of the electric trim will
automatically disconnect the autopilot.
CWS – Control Wheel Steering Button
When depressed, allows the pilot to manually control the aircraft (disengages the pitch and roll
servos) without cancellation of any of the selected modes. Will engage the flight director if not
previously engaged. Automatically synchronises the flight director/ autopilot to the pitch attitude
present when the CWS switch is released or to the present pressure altitude when in ALT hold mode.
Will cancel GS couple. The aircraft must pass through the glide slope again to allow GS recouple.
Before Take-off Reliability Tests
Gyros....................................................Allow 3 – 4 minutes for gyros to come up to speed
Avionics Power #1 Switch ....................ON
Preflight Test Button.............................PRESS and OBSERVE
a.
All annunciator lights on.
b.
TRIM annunciator flashing.
c.
After approximately 5 seconds, all annunciator lights off except the AP light, which will
flash approximately 12 times and then remain off.
If the TRIM warning light stays on, the autotrim did not pass the preflight test. The autopilot
circuit breaker should be pulled. Autopilot and manual electric trim will be inoperative.
Manual Electric Trim ............................................... TEST as follows:
a.
Actuate the left side split switch unit to the fore and aft positions. The trim wheel
should not move. Rotate the trim wheel manually against the engaged clutch to check
for override capability.
b.
Actuate the right side split switch unit to the fore and aft positions. The trim wheel
should not move and normal trim wheel force is required to move it manually.
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AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
c.
Press the A/P DISC/TRIM INTER switch down and hold. The electric trim should not
operate.
Flight Director.......................................ENGAGE by pressing the FD or CWS button
Autopilot ...............................................ENGAGE by pressing the AP ENG button
Flight Controls ......................................MOVE fore, aft, left and right to verify that the autopilot can be
overpowered
A/P DISC/TRIM Button.........................PRESS. Verify that the autopilot disconnects and all flight
director modes are cancelled
TRIM.....................................................SET for takeoff
Autopilot Limitations
•
•
•
•
•
•
During autopilot operation, a pilot with seat belt fastened must be seated in the left seat.
The autopilot and yaw damper must be OFF during takeoff and landing.
The system is approved for Cat I ILS conditions only – Approach Mode Selected.
The autopilot must be disconnected at 200 feet AGL.
The autopilot must be OFF during use of standby flap system.
In accordance with FAA recommendation, use of basic pitch attitude hold mode is recommended
during operation in severe turbulence.
• Maximum Altitude Loss Due To Malfunction
Cruise, Climb and Descent ..................................... 500 ft
Manoeuvring ........................................................... 100 ft
Approach ................................................................. 100 ft
Flight Command Indicator (FCI)
Command Bar
Displays computed steering commands referenced to the
symbolic aircraft. The command bar is visible only when
the FD is selected. The command bar is visible only when
the FD mode is selected. The command bar will be biased
out of view whenever the system is invalid or a flight
director mode is not selected.
Figure 13-5 KI-256 Flight Command Indicator
FCI Symbolic Aircraft
The aircraft pitch and roll attitude is displayed by the
relationship between the fixed symbolic aircraft and the
moveable background. During flight director operation, the
symbolic aircraft is flown to align it with the command bar
to satisfy the flight director commands.
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Horizontal Situation Indicator (HSI)
Figure 13-6 KI-525A Horizontal Situation Indicator
The HSI provides a pictorial presentation of the aircraft deviation relative to VOR radials or localizer
beams. It also displays glide slope deviations and gives heading reference with respect to magnetic
north. The gyro is driven electrically.
Nav Flag
The flag is in view when the NAV receiver signal in inadequate. When a NAV flag is present in the
HSI, the autopilot operation is not affected. The pilot must monitor the navigation indicator for a NAV
flag to ensure that the autopilot and/or flight director are tracking valid navigation information.
Lubber Line
Indicates aircraft magnetic heading on the compass card.
Heading Warning Flap (HDG)
When the flag is in view, heading display is invalid. If a HDG flag appears and a lateral mode (HDG,
NAV or APP) is selected, the autopilot will be disengaged. The autopilot may be re-engaged in the
basic wings-level mode along with any vertical mode. The CWS switch would be used to manually
manoeuvre the aircraft laterally.
Course Bearing Pointer
Indicates the selected VOR course or localizer course on the compass card. The selected VOR radial
or localizer course remains set on the compass card when the compass card rotates.
TO/FROM Flag Indicator
Indicates direction of the VOR station relative to the selected course.
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Dual Scale Slope Pointers
Indicates the glide slope scale aircraft displacement from the glide slope beam centre. Glide slope
pointers in view indicate a usable glide slope signal is being received.
Glide Slope Scales
Indicates displacement from the glide slope beam centre. A glide slope deviation bar displacement of
2 dots represents a full scale deflection (0.7 degree) deviation above or below glide slope beam
centreline.
Course Deviation Bar
The centre position of the omni bearing pointer moves laterally to pictorially indicate the relationship of
the aircraft to the selected course. It indicates degrees of angular displacement from VOR radials or
displacement in nautical miles from RNAV courses.
Course Deviation Scale
A course deviation bar displacement of 5 dots represents full scale (VOR = 10 degrees, LOC = 2½
degrees, RNAV = 5 nm)
Heading Bug
Moved by the heading selector knob to select the desired heading.
Slaving Accessory And Compensator
This unit controls the compass system.
Figure 13-7 KA-51B Slaving Accessory And Compensator Unit
MAN/AUTO – Free/Slave Compass Slave Switch
Selects either the manual or automatic slaving mode for the compass system.
CW/CCW – Compass Manual/Slave Switch
With the switch in the FREE position, allows manual compass card slaving in clockwise or counterclockwise position. The switch is spring loaded to the centre position.
Slaving Meter
Indicates the difference between the displayed heading and magnetic heading. Up deflection indicates
a clockwise error of the compass card.
© Solenta Aviation (Pty) Ltd
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Weather Radar
The Weather Radar System consists of a receiver-transmitter and antenna mounted in a pod on the
right-hand wing, and a panel mounted radar indicator. There a several different radar installations for
the C208B. The most popular radar installation is the Narco (Bendix/King) KWX-56, however the
same basic details should also be applicable to other installations.
Figure 13-8 Narco KWX-56 Radar Indicator, And Wing Pod Installation
Function Selector Switch
The function selector switch controls application of power and selects mode of operating for
transmitting, testing, and warm-up. Switch positions are as follows:
OFF
Primary power is removed from the system.
SBY – Standby
After 30 seconds in this mode, places the system in operational ready status. Use during warm-up
and in-flight periods when the system is not in use. The word STBY is displayed in the lower left
corner.
TST – Test
Selects test function to determine the operational ready status. A test pattern is displayed. No
transmission is possible. The word TEST is displayed in the lower left corner.
ON
Selects the position for normal operation. Commences radar transmission. Defaults to WX mode
and 80 nm range. WX will be displayed in the lower left corner and 80 will be displayed just above
the right end of the top concentric range mark.
TILT - Antenna Tilt Control
The tilt control adjusts the antenna to move the radar beam up to +15 degrees above the horizontal or
to a maximum of -15 degrees below the horizontal position. The horizontal position is indicated as
zero degrees on the control. The tilt angle selected is displayed in the upper right corner of the
indicator.
STAB – Stabilisation Pushbutton
When pushed in, stabilisation is disabled. The words “STAB OFF” will flash on and off in the upper left
corner. When in out position, the antenna stabilisation is restored. This feature keeps the antenna
azimuth parallel to the ground, independent of aircraft attitude. The antenna compensates for up to ±
25° of aircraft pitch and roll.
GAIN – Gain Control
Permits adjusting the radar receiver gain in the terrain MAP mode only. In the test function as well as
in all weather modes the receiver gain is preset.
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
MAP – Ground Mapping Pushbutton
Selects ground mapping when pressed. When MAP mode is selected, the word MAP is displayed in
the lower left corner. GAIN control becomes an operator function. Manual GAIN control is important in
obtaining a definitive presentation during varying topographic conditions. Prominent terrain features
are presented in green, yellow and red. Magenta is not used.
WxA – Weather Alert Pushbutton
Selects weather alert mode when selected. WxA is displayed in the lower left corner of the screen.
Storm cells are displayed in up to four colours depending on level of intensity. Green, yellow, red and
magenta. Magenta area of storm flashes between magenta and black to indicate strong storm cell.
Wx – Weather Pushbutton
Selects weather mode when selected. Wx is displayed in the lower left corner of the screen.
Operation for the Wx mode is the same as the WxA except the areas of strong rainfall appear as a
steady magenta colour and will not flash between magenta and black as it does in WxA mode.
Radar Safety Precautions
•
•
Do not turn on or operate within 15 feet of ground personnel or metal objects.
Do not turn on or operate during refuelling operations.
Operational Procedures
•
•
•
•
The system is designed so that full operation is possible 30 seconds after turn on.
The pilot may choose to leave the function switch in OFF rather than SBY if no significant weather
is in the immediate area.
Always turn the indicator to SBY or OFF before turning the avionics master switch off.
The system will power-down in about 5 seconds after switched to the OFF position, to allow time
for the antenna to move to the down position.
• Margin Of Avoidance
•
•
•
•
•
Avoid any cell by 5 nm if the temperature is above freezing and by 10 nm if below freezing.
When penetrating an area of storms and passing between two cells, a minimum corridor of 10 to
20 nm is required.
If the radar return is very intense with hooks and scallops, increase this distance accordingly.
Avoid by 10 nm or any cell that is changing shape rapidly.
Never fly under an overhang from a mature cumulonimbus cloud.
© Solenta Aviation (Pty) Ltd
13-10
Dec-2005
LIMITATIONS
Table of Contents
Airspeeds .................................................. 14-1
Engine Start Cycles................................... 14-1
Power Plant............................................... 14-1
Operating Altitudes.................................... 14-2
Weights ..................................................... 14-2
Flight Load Factor Limits ...........................14-2
Maximum Full Rudder Sideslip Duration ...14-2
Fuel ............................................................14-3
Use Of Avgas.............................................14-3
Minimum Oil Quantity ................................14-3
Airspeeds
Maximum Operating Speed (VMO) ..................175 KIAS
Manoeuvring Speed (VA)
8,750 lbs. ....................................................................148 KIAS
7,500 lbs. ....................................................................137 KIAS
6,250 lbs. ....................................................................125 KIAS
5,000 lbs. ....................................................................112 KIAS
Maximum Flap Extended Speed (VFE)
0° To 10° Flaps...........................................................175 KIAS
10° To 20° Flaps.........................................................150 KIAS
20° To 30° Flaps.........................................................125 KIAS
Figure 14-1 C208B ASI
Engine Start Cycles
Aircraft Battery Start Cycle ................................ 30 seconds ON - 60 seconds OFF,
30 seconds ON - 60 seconds OFF,
30 seconds ON - 30 MINUTES OFF.
External Power Start Cycle ............................... 20 seconds ON - 120 seconds OFF,
20 seconds ON - 120 seconds OFF,
20 seconds ON - 60 MINUTES OFF.
Powerplant (PT6A-114A)
Minimum NG For Starting ................................... 12% (preferably 18%)
Emergency Airstart NG ....................................... Minimum 10%
Rise In NG And ITT During Start ..................... Maximum 10 seconds
Battery Start Minimum Voltage........................ 24 volts minimum
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
APU Start Minimum Voltage............................. 20 volts minimum
APU Capacity ........................................................ 800 to 1700 Amp
Figure 14-2 C208B Engine Instruments
MAXIMUM ITT
°C
GAS
GENERATOR
RPM % NG
PROPELLER
RPM
805
101.6
1900
765
101.6
1900
1865
740
101.6
1900
----
685
52 (Minimum)
----
1865
805
101.6
1825
2400
850
102.6
2090
----
1090
----
----
1865
805
101.6
1900
POWER
SETTING
TORQUE
FT-LBS
TAKEOFF
1865
MAXIMUM CLIMB
1865
MAXIMUM CRUISE
IDLE
MAXIMUM REVERSE
(1 Minute Maximum)
TRANSIENT
(2 Second Maximum)
STARTING
(2 Second Maximum)
MAXIMUM
CONTINUOUS
(20 sec’s max.)
(5 mins max.)
Operating Altitudes
Non-icing Conditions .....................................................25,000 ft
Icing Conditions.............................................................20,000 ft
Conditions With Any Ice On The Airframe ...................20,000 ft
Weights
Maximum Ramp Weight............................................... 8,785 lbs
Maximum Take-off Weight........................................... 8,750 lbs
Maximum Landing Weight ........................................... 8,500 lbs
Known Icing Operation................................................. 8,550 lbs
Forward Cargo Compartment......................................... 230 lbs
Centre Forward Cargo Compartment............................. 310 lbs
Centre Aft Cargo Compartment...................................... 270 lbs
Aft Cargo Compartment .................................................. 280 lbs
Flight Load Factor Limits
Flaps UP ................................................................ +3.8g, -1.52g
Flaps DOWN ......................................................................+2.4g
© Solenta Aviation (Pty) Ltd
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Dec-2005
AIRCRAFT TECHNICAL NOTES
Cessna 208B Grand Caravan
Maximum Full Rudder Sideslip Duration
3 minutes (due to possible fuel starvation).With low fuel reserves (FUEL LOW annunciator ON),
continuous uncoordinated flight with more than one quarter “ball” out of centre position is prohibited.
Unusable fuel quantity increases when more severe sideslip is maintained.
Fuel
Total Fuel...................................................................... 2,244 lbs
Total Unusable Fuel .......................................................... 20 lbs
Total Useable Fuel ....................................................... 2,224 lbs
Maximum Fuel Imbalance............................................... 200 lbs
Use Of Avgas
The use of avgas is restricted to emergency use, and shall not be used for more than 150 hours in
one overhaul period. A mixture of one part avgas and three parts of Jet A, Jet A-1, JP-1 or JP-5 may
be used for emergency purposes for a maximum of 450 hours per overhaul period. Avgas contains
tetraethyl lead (TEL) to increase its critical temperature and pressure. When mixed with jet fuel, the
result is that the lead ends up sticking to the turbine blades and vanes.
Minimum Oil Quantity
Fill to 1½ quarts of MAX HOT or MAX COLD position on dipstick.
© Solenta Aviation (Pty) Ltd
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Dec-2005