Cessna 208B Grand Caravan EX. Airplane Flight Manual (2012)

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN ENTRY DOORS (Continued)

PASSENGER ENTRY DOOR (Passenger Version Only)

The entry door for passengers consists of an upper and lower section.

When opened, the upper section swings upward and the lower section

drops down providing integral steps to aid in boarding or exiting the

airplane. The upper door section incorporates a conventional exterior

door handle with a separate key-operated lock, a pushbutton exterior

door release, and an interior door handle which snaps into a locking

receptacle. The lower door section features a flush handle which is

accessible from either inside or outside the airplane. 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. A cabin door

unlatched warning system is provided as a safety feature so that if the

upper door is not properly latched, an amber DOOR UNLATCHED

annunciator located on the PFD will be shown to alert the pilot that.

To enter the airplane through the passenger entry door, depress the

exterior pushbutton door release, rotate the exterior door handle on the

upper door section counterclockwise to the open position, and raise the

door section to the overcenter position. Following this action, the

automatic door lift with the telescoping gas spring raises the door to the

full up position. When the upper section is open, 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. The door steps deploy automatically from their

stowed positions.

WARNING

The outside proximity of the lower door section

must be clear before opening the door.

(Continued Next Page)

7-38

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CABIN ENTRY DOORS (Continued)

PASSENGER ENTRY DOOR (Passenger Version Only)

(Continued)

To close the passenger entry door from the inside of the airplane, grasp

the support cables of the lower door section and pull the door up until

the top edge is within reach, then grasp the center of the door and pull

inboard until the door is held snugly against the fuselage door frame.

Rotate the inside handle forward to the CLOSE position and latch the

lower door section.

Check that the lower front and rear latches are correctly engaged. After

the lower door section is secured, grasp the pull strap on the upper

door section and pull down and inboard. As the door nears the closed

position, pull inboard firmly to make sure the latching pawls engage

correctly. When the latching pawls are engaged, rotate the inside

handle counterclockwise to the horizontal (latched) position, but do not

use excessive force. If the handle will not rotate easily, the door is not

fully closed. Use a more firm closing motion to get the latching pawls to

engage and rotate the door handle again to the latched position. Then

snap the interior handle into its locking receptacle.

CAUTION

Refer to Section 3, Emergency Procedures, for proper

operational procedures to be followed if the passenger

entry door should inadvertently open in flight.

To exit the airplane through the passenger entry door, pull the upper

door section inside handle from its locked position receptacle, rotating

the handle clockwise to the open position as you push the door

outward. When the door is partially open, the automatic door lift will

raise the upper door section to the fully open position. Next, rotate the

door handle of the lower section up and aft to the open position and

push the door outward. The telescoping gas spring will lower the door

to its fully open position and the integral steps will deploy.

WARNING

The outside proximity of the lower door section

must be clear before opening the door.

(Continued Next Page)

208BPHCUS-00

U.S.

7-39

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN ENTRY DOORS (Continued)

PASSENGER ENTRY DOOR (Passenger Version Only)

(Continued)

To close the passenger entry door from outside the airplane, raise the

lower door section until the door is held firmly against the door frame in

the fuselage. Rotate the inside handle of the lower door section forward

and down to the CLOSE position. After the lower door section is

secured, grasp the pull strap on the upper door section and pull down.

As the door nears the closed position, grasp the edge of the door and

push inward firmly to make sure the latching pawls engage correctly.

When engaged, rotate the outside door handle clockwise to the

horizontal

(latched) position. After entering the airplane, snap the

interior handle of the upper door into its locking receptacle (unless

cargo obstructs access to the door). If desired when leaving the

airplane parked, use the key in the outside key lock to lock the handle

in the horizontal position.

WARNING

Do not use the outside key lock to lock the door

prior to flight. The door could not be opened from

the inside if it were needed as an emergency exit.

CAUTION

Failure to properly latch the upper passenger door

section will result in the amber DOOR WARNING

annunciator being shown on the PFD. Inattention to this

safety feature may allow the upper cargo door to open

in flight.

The exterior pushbutton-type lock release located on the upper door

section just forward of the exterior door handle operates in conjunction

with the interior door handle. It is used whenever it is desired to open

the door from outside the airplane while the interior door handle is in

the locked position. Depress the pushbutton to release the lock of the

interior door handle and to allow the exterior door handle to function

normally to open the door.

7-40

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CARGO DOORS

A two-piece cargo door is installed on the left side of the airplane just

aft of the wing trailing edge. The cargo door is divided into an upper

and a lower section. When opened, the upper section swings upward

and the lower section swings forward to create a large opening in the

side of he fuselage which facilitates the loading of bulky cargo into the

cabin. The upper section of the cargo door incorporates a conventional

exterior door handle with a separate key-operated lock, and, on the

Passenger Version only, a pushbutton exterior emergency door release

and an interior door handle which snaps into a locking receptacle. The

upper door also incorporates two telescoping door lifts which raise the

door to the fully open position, when opened. A cargo door open

warning system is provided as a safety feature so that if the upper door

is not properly latched an amber annunciator, labeled DOOR

UNLATCHED, located on the PFD, illuminates to alert the pilot. The

lower door section features a flush handle which is accessible from

either inside or outside the airplane. The handle is designed so that

when the upper door is closed, the handle cannot be rotated to the

open position.

WARNING

In an emergency, do not attempt to exit the cargo

version through the cargo doors. Because the

inside of the upper door has no handle, exit from

the airplane through these doors without outside

assistance is not possible.

CAUTION

Failure to properly latch the upper cargo door section

will result in illumination of the amber DOOR WARNING

annunciator. Inattention to this safety feature may allow

the upper cargo door to open in flight.

(Continued Next Page)

208BPHCUS-00

U.S.

7-41

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CARGO DOORS (Continued)

To open the cargo door from outside the airplane, depress the upper

door section exterior pushbutton door release (Passenger Version only)

and rotate the exterior door handle clockwise to the open position.

Following this action, the telescoping door lifts will automatically raise

the door to the full up position. When the upper section is open, release

the lower section by pulling up on the inside door handle and rotating

the handle to the OPEN position. Open the door forward until it swings

around next to the fuselage where it can be secured to the fuselage by

a holding strap or chain.

To close the cargo door from outside the airplane, disconnect the

holding strap or chain from the fuselage, swing the door aft to the

closed position, and hold the door firmly against the fuselage door

frame to assure engagement of the latching pawls. Rotate the inside

handle forward and down to the CLOSE position to latch the lower door

section. After the lower door section is secured, grasp the pull strap on

the upper door section and pull down. As the door nears the closed

position, grasp the edge of the door and push inward firmly to assure

engagement of the latching pawls. When engaged, the exterior door

handle can be rotated counterclockwise to the horizontal (latched)

position. On the Passenger Version only, after entering the airplane,

snap the upper door interior handle into its locking receptacle (unless

cargo obstructs access to the door). If desired when leaving the

airplane parked, use the key in the outside key lock to lock the handle

in the horizontal position.

(Continued Next Page)

7-42

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CARGO DOORS (Continued)

To open the cargo door from inside the airplane (Passenger Version

only), pull the inside door handle of the upper door section from its

locked position receptacle. Rotate the handle counterclockwise to the

vertical position, and push the door outward. When the door is partially

open, the automatic door lifts will raise the upper door section to the

fully open position. Next, rotate the door handle of the lower section

door up and aft to the open position and push the aft end of the door

outward. The door may be completely opened and secured to the

fuselage with the holding strap or chain from outside.

WARNING

Do not attempt to exit the cargo version through the

cargo doors. Because the inside of the upper door

has no handle, exit from the airplane through these

doors is not possible without outside assistance.

To close the cargo door from inside the airplane (Passenger Version

only), disconnect the holding strap or chain from the fuselage and

secure it to the door. Pull the door aft to the closed position and hold the

aft edge of the door firmly against the fuselage door frame to assure

engagement of the latching pawls. Rotate the inside handle forward

and down to the CLOSE position to latch the lower door section (refer

to Section 2, Placards). After the lower door section is secured, grasp

the pull strap on the upper door section and pull down. As the door

nears the closed position, grasp the edge of the door and pull inward

firmly to assure engagement of the latching pawls. When engaged, the

interior door handle can be rotated clockwise to the horizontal position.

Snap the handle into its locking receptacle.

208BPHCUS-00

U.S.

7-43

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN WINDOWS

The airplane is equipped with a two-piece windshield reinforced with a

metal center strip. The passenger version has sixteen cabin side

windows of the fixed type including one each in the two crew entry

doors, two windows in the cargo door upper section, and one window in

the upper section of the passenger entry door. The pilot’s side window

incorporates a small triangular foul weather window. The foul weather

window may be opened for ground ventilation and additional viewing by

twisting the latch. The cargo version has only two cabin side windows,

one in each crew entry door.

CONTROL LOCKS

A control lock is provided to lock the aileron and elevator control

surfaces to prevent damage to these systems by wind buffeting while

the airplane is parked. The lock consists of a shaped steel rod and flag.

The flag identifies it as a control lock and cautions about its removal

before starting the engine. To install the control lock, align the hole in

the right side of the pilot’s control wheel shaft with the hole in the right

side of the shaft collar on the instrument panel and insert the rod into

the aligned holes. Installation of the lock will secure the ailerons in a

neutral position and the elevators in a slightly trailing edge down

position. Proper installation of the lock will place the flag over the left

sidewall switch panel.

The Rudder Gust Lock is a positive locking device consisting of a

bracket assembly and a bolt action lock attached to the rear bulkhead

inside the tailcone stinger below the rudder. When engaged, the rudder

is locked in the neutral position. A placard located below the lock

handle shaft on the left side of the tailcone explains the operation of the

rudder gust lock. The rudder gust lock is manually engaged and

disengaged on the ground by turning the airfoil-shaped handle

mounted on the shaft projecting from the left side of the tailcone. The

lock is engaged by turning the handle downward so that its trailing edge

points nearly due aft.

(Continued Next Page)

7-44

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CONTROL LOCKS (Continued)

The Rudder Gust Lock has a fail-safe connection to the elevator control

system to ensure that it will always be disengaged before the airplane

becomes airborne. This fail-safe connection automatically disengages

the lock when the elevator is deflected upward about one-fourth of its

travel from neutral. The pilot is responsible for disengaging the Rudder

Gust Lock during the preflight inspection and operating the fail-safe

disengagement mechanism by momentarily deflecting the elevator to

the full up position after the control lock is removed and before starting

the engine. If these procedures are not followed the rudder and rudder

pedals will be locked in the neutral position making ground steering

impossible. In the event that the engagement of the Rudder Gust Lock

goes completely unnoticed and the pilot commences a takeoff run with

the rudder system locked, the upward elevator deflection during

rotation will disengage the Rudder Gust Lock.

Because of the fail-safe system, the elevator lock should always be

engaged prior to engaging the Rudder Gust Lock when securing the

airplane after shutdown.

NOTE

The control lock and any other type of locking device should

be removed or unlocked prior to starting the engine.

208BPHCUS-00

U.S.

7-45

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE

The Pratt & Whitney Canada Inc. PT6A-140 powerplant is a free

turbine engine. It utilizes two independent turbines; one driving a

compressor in the gas generator section, and the second driving a

reduction gearing for the propeller.

Inlet air enters the engine 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, assembled as an integral unit.

A row of stator vanes located between each stage of compressor rotor

blades diffuses the air, raises its static pressure and directs it to the

next stage of compressor rotor blades. The compressed air passes

through diffuser ducts which turn it 90° in direction. It is then routed

through straightening vanes into the combustion chamber.

The combustion chamber liner located in the gas generator case

consists of an annular reverse-flow weldment provided with varying

sized perforations which allow entry of compressed air. The flow of air

changes direction to enter the combustion chamber liner where it

reverses direction and mixes with fuel. The location of the combustion

chamber liner eliminates the need for a long shaft between the

compressor and the compressor turbine, thus reducing the overall

length and weight of the engine.

Fuel is injected into the combustion chamber liner by

14 simplex

nozzles supplied by a dual manifold. the mixture is initially ignited by

two spark igniters which protrude into the combustion chamber liner.

The resultant gases expand from the combustion chamber liner,

reverse direction and pass through the compressor turbine guide vane

to the compressor turbine. The turbine guide vanes ensure that the

expanding gases impinge on the turbine blades at the proper angle,

with a minimum loss of energy. The still expanding gases pass forward

through a second set of stationary guide vanes to drive the power

turbine.

(Continued Next Page)

7-46

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE (Continued)

The compressor and power turbines are located in the approximate

center of the engine with their shafts extending in opposite directions.

The exhaust gas from the power turbine is directed through an exhaust

plenum to the atmosphere via a single exhaust port on the right side of

the engine.

The engine is flat rated at 867 shaft horsepower (2397 foot-pounds

torque at 1900 RPM. The speed of the gas generator (compressor)

turbine (N_g) is 37,468 RPM at 100% N_g. Maximum permissible speed

of the gas generator is 38,900 RPM which equals 103.7% N_g. The

power turbine speed is 33,000 RPM at a propeller shaft speed of 1900

RPM.

All engine-driven accessories, with the exception of the propeller

tachometer-generator and the propeller governors, are mounted on the

accessory gearbox located at the rear of the engine. These are driven

by the compressor turbine with a coupling shaft which extends the drive

through a conical tube in the oil tank center section.

The engine oil supply is contained in an integral tank which forms part

of the compressor inlet case. The tank has a drain and fill capacity of

9.5 U.S. quarts and is provided with a dipstick and drain plug.

The power turbine drives the propeller through a two-stage planetary

reduction gearbox located on the front of the engine. The gearbox

embodies an integral torquemeter device which is instrumented to

proved an accurate indication of the engine power output.

208BPHCUS-00

U.S.

7-47

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE COMPONENTS

1. Propeller Shaft

13. Accessory Gearbox Drive Shaft

2. Propeller Governor Drive Pad

14. Accessory Gearbox Cover

3. Second Stage Planetary Gear

15. Starter-Generator Drive Shaft

4. First Stage Planetary Gear

16. Oil Scavenge Pump

5. Power Turbine Shaft

17. Number 1 Bearing

6. Fuel Nozzle

18. Compressor Bleed Valve

7. Power Turbine

19. Number 2 Bearing

8. Combustion Chamber

20. Number 3 Bearing

9. Compressor Turbine

21. Number 4 Bearing

10. Centrifugal Compressor Impeller

22. Exhaust Outlet

11. Axial-Flow Compressor Impellers

23. Chip Detector

(3)

24. Roller Bearing

12. Compressor Air Inlet

25. Thrust Bearing

Figure 7-8

7-48

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE CONTROLS

The engine is operated by four separate controls consisting of a

POWER lever, EMERGENCY POWER lever, PROP RPM lever and a

FUEL CONDITION lever. The POWER and FUEL CONDITION levers

are engine controls while the PROP RPM lever controls propeller

speed and feathering.

POWER LEVER

The POWER lever is connected through linkage to a cam assembly

mounted in front of the fuel control unit at the rear of the engine. The

POWER lever controls engine power through the full range from

maximum takeoff power back through idle to full reverse. The lever also

selects propeller pitch when in the BETA range. The POWER lever has

MAX, IDLE, and BETA and REVERSE range positions. The range from

MAX position through IDLE enables the pilot to select the desired

power output from the engine. The BETA range enables the pilot to

control propeller blade pitch from idle thrust back through a zero or no-

thrust condition to maximum reverse thrust.

CAUTION

The propeller reversing linkage can be damaged if the

power lever is moved aft of the idle position when the

propeller is feathered.

208BPHCUS-00

U.S.

7-49

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE CONTROLS (Continued)

EMERGENCY POWER LEVER

The EMERGENCY POWER lever is connected through linkage to the

manual override lever on the fuel control unit and governs fuel supply to

the engine should a pneumatic malfunction occur in the fuel control

unit. When the engine is operating, a failure of any pneumatic signal

input to the fuel control unit will result in the fuel flow decreasing to

minimum idle (about 48% N_g at sea level and increasing with altitude).

The EMERGENCY POWER lever allows the pilot to restore power in

the event of such a failure. The EMERGENCY POWER lever has

NORMAL, IDLE, and MAX positions. The NORMAL position is used for

all normal engine operation when the fuel control unit is operating

normally and engine power is selected by the POWER lever. The range

from IDLE position to MAX governs engine power and is used when a

pneumatic malfunction has occurred in the fuel control unit 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 (full aft) position to the

IDLE position.

NOTE

• The knob on the EMERGENCY POWER lever has

crosshatching. The crosshatching is visible when the

lever is in MAX position.

• The EMERGENCY POWER lever is annunciated by a

red (OFF and STRT modes) or amber (RUN mode)

EMERG PWR LVR on the PFD.

• The red annunciation will illuminate whenever the

EMERGENCY POWER lever is unstowed from the

NORMAL position with the ITT indications in either of the

OFF or STRT modes. This precaution is intended to

preclude starting of the engine with the EMERGENCY

POWER lever inadvertently placed in any position other

than NORMAL.

• The amber annunciation will illuminate whenever the

EMERGENCY POWER lever is unstowed during normal

operations.

(Continued Next Page)

7-50

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE CONTROLS (Continued)

EMERGENCY POWER LEVER (Continued)

CAUTION

• The EMERGENCY POWER lever and its associated

manual override system are considered to be an

emergency system and should be used only in the

event of a fuel control unit malfunction. When

attempting a normal start, the pilot must make sure

that the EMERGENCY POWER lever is in the

NORMAL (full aft) position; otherwise, an over-

temperature condition may result.

• When using the fuel control manual override system,

engine response may be more rapid than when

using the POWER lever. Additional care is required

during engine acceleration to avoid exceeding

engine limitations.

Operation of the EMERGENCY POWER lever is prohibited with the

primary POWER lever out of the IDLE position. The EMERGENCY

POWER lever overrides normal fuel control functions and results in the

direct operation of the fuel metering valve. The EMERGENCY POWER

lever will override the automatic fuel governing and engine acceleration

scheduling controlled during normal operation by the primary POWER

lever.

CAUTION

Inappropriate use of the EMERGENCY POWER lever

may adversely affect engine operation and durability.

Use of the EMERGENCY POWER lever during normal

operation of the POWER lever may result in engine

surges, or exceeding the ITT, N_G, and torque limits.

208BPHCUS-00

U.S.

7-51

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE CONTROLS (Continued)

PROPELLER CONTROL LEVER

The PROP RPM lever is connected through linkage to the propeller

governor mounted on top of the front section of the engine, and

controls propeller governor settings from the maximum RPM position to

full feather. The PROP RPM 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. PROP RPM lever settings

from the MAX position to MIN permit the pilot to select the desired

engine RPM for cruise. The FEATHER position is used during normal

engine shutdown to stop rotation of the power turbine and front section

of the engine. Since lubrication is not available after the gas generator

section of the engine has shut down, rotation of the forward section of

the engine is not desirable. Also, feathering the propeller when the

engine is shut down minimizes propeller windmilling during windy

conditions. A mechanical stop in the lever slot requires that the PROP

RPM lever be moved to the left to clear the stop before it can be moved

into or out of the FEATHER position.

FUEL CONDITION LEVER

The FUEL CONDITION lever is connected through linkage to a

combined lever and stop mechanism on the fuel control unit. The lever

and stop also function as an idle stop for the fuel control unit rod. The

FUEL CONDITION lever controls the minimum RPM of the gas

generator turbine (N_g) when the POWER lever is in the IDLE position.

The FUEL CONDITION lever has CUTOFF, LOW IDLE, and HIGH

IDLE positions. The CUTOFF position shuts off all fuel to the engine

fuel nozzles. LOW IDLE positions the control rod stop to provide an

RPM of 55% N_g. HIGH IDLE positions the control rod stop to provide an

RPM of 65% N_g.

QUADRANT FRICTION LOCK

A quadrant friction lock, located on the right side of the pedestal, is

provided to minimize creeping of the engine controls once they have

been set. The lock is a knurled knob which increases friction on the

engine controls when rotated clockwise.

7-52

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE INDICATING SYSTEM (EIS)

The G1000 Engine Indicating System

(EIS) provides graphical

indicators and numeric values for engine, fuel, and electrical system

parameters to the pilot. The EIS is shown in a vertical strip on the left

side of the PFD during engine starts and on the MFD during normal

operation. If either the MFD or PFD fails during flight, the EIS is shown

on the remaining display.

The EIS consists of two pages that are selected using the ENGINE

softkey. The ENGINE page provides indicators for Engine Torque,

Engine ITT, Gas Generator RPM%, Propeller RPM, Oil Pressure, Oil

Temperature, Fuel Quantity, Fuel Flow, Battery Amps and Bus Voltage.

When the ENGINE softkey is pressed, the SYSTEM softkey will appear

adjacent to the ENGINE softkey. The SYSTEM page provides

numerical values for parameters on the ENGINE page that are shown

as indicators only. Torque, ITT, N_g% and N_p RPM are displayed

identically on the SYSTEM page. The SYSTEM page also provides

numerical indication for fuel quantity, fuel totalizer (pounds remaining

and pounds used), generator amps, standby alternator amps, battery

amps and bus voltage.

The engine and airframe unit provides data to the EIS, which displays

the data for the ENGINE page described below. Engine operation is

monitored by: torque, ITT, N_g%, propeller RPM, oil pressure, oil

temperature, and fuel flow.

208BPHCUS-00

U.S.

7-53

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE INDICATING SYSTEM (EIS) (Continued)

TORQUE INDICATIONS

Torque (TRQ) indication is displayed at the top of both the ENGINE and

SYSTEM pages. The indicator is a round gage with a white pointer. The

transmitter senses the difference between the engine torque pressure

and the pressure in the engine case and transmits this data to the

G1000. Normal operating range is indicated by a green arc that

extends from 0 to redline. The Torque (TRQ) indicator incorporates a

dynamic redline varies with OAT and altitude.

DYNAMIC REDLINE

The dynamic redline is a graphical representation of takeoff power

below 16,000 feet MSL and Maximum Rated Power above 16,000 feet

MSL as depicted in Section 5, Performance, Figure 5-8, Maximum

Engine Torque For Takeoff chart and Figure 5-9 Maximum Engine

Torque For Climb. The dynamic redline automatically compensates for

altitude and temperature changes and adjusts displayed takeoff torque

for inertial separator deployment and bleed air heat switch position.

Failure to comply with the dynamic redline indication can result in

accelerated engine wear, unscheduled engine maintenance and

increased operating costs even though no other published engine

limitation has been exceeded.

STATIC REDLINE

The dynamic redline reverts to a static redline whenever the white

TORQUE GAGE annunciation is shown on the PFD. The white

TORQUE GAGE annunciation indicates when there is an ALT

MISCOMP message, red “X” through the OAT, Altitude, Airspeed and

Vertical Speed Indicators on PFD1 or a 5°C temperature difference

between OAT’s on PFD1 and PFD2. With a white TORQUE GAGE

annunciation, the dynamic redline will become fixed at 2397 FT-LB

regardless of temperature or altitude changes. The cruise torque bug

will become inoperative and will not be shown on the EIS torque

indicator. Refer to Section 5, Performance, for the appropriate power

settings.

7-54

U.S.

208BPHCUS-01

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE INDICATING SYSTEM (EIS) (Continued)

TORQUE INDICATIONS (Continued)

MAXIMUM CRUISE TORQUE BUG

For normal cruise flight when prop RPM is between 1600-1900, a

maximum cruise torque “bug” is included on the arc. This indicates

maximum allowed cruise torque per the cruise performance and

maximum torque charts in Section 5, Performance. The blue maximum

cruise torque bug located on the EIS torque indication is not to be used

as the primary means of setting cruise torque. Always refer to the

appropriate performance chart in Section 5 of the POH/AFM.

PROPELLER RPM INDICATIONS

The PROP RPM is indicated numerically below Gas Generator Ng %

RPM indicator. The digits are white with RPM between 0-1599 RPM,

green between

1600-1900 RPM, and white numerals on a red

background when RPM is greater than or equal to 1910 after a 20

second delay. The instrument is electrically operated from the propeller

tachometer-generator which is mounted on the right side of the front

case.

ITT INDICATION

Interstage Turbine Temperature (ITT) is indicated below torque gage

and is round dial gage with a white pointer. This instrument displays

gas temperature between the compressor and power turbines. With the

engine off, or during start, ITT indicator displays a green band from

100°C to 870°C and a redline at 871°C. The gage is graduated at 100

degree intervals from 600°C to 1100°C.

With the engine running, a green arc indicates normal operating range

from 100°C to 825°C, an amber caution arc from 826°C-849°C and a

redline at 850°C. The gage is graduated at 50 degree intervals from

600°C-950°C.

During any temperature exceedance, the digital readout will reverse to

white digits on a red background and the moving pointer will turn red.

208BPHCUS-00

U.S.

7-55

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE INDICATING SYSTEM (EIS) (Continued)

GAS GENERATOR RPM INDICATIONS

Gas generator RPM (Ng) is displayed below the ITT indicator and uses

a round dial style gage with a white pointer. RPM is displayed as a

percentage of maximum gas generator RPM. The Ng indicator displays

a green band from 55% to 103.6% and a redline at 103.7%. The Ng %

RPM labels are displayed in white with green digital values between

0%-103.6% RPM. Once Ng % RPM is greater than or equal to 103.7 %

for more than a 20 seconds, the pointer switches to red and numerals

switch to white on a red background. If Ng % RPM ever reaches

105.4% the pointer will immediately turn red and digital values revert to

white on a red background.

The Ng % RPM indicator is electrically operated from the gas generator

tachometer-generator mounted on the lower right portion of the

accessory case. The gage has major tick marks at 0, 12, 50, and 110%

and minor tick marks at 10% intervals from 50 to 100%, with a redline

at 103.7%.

FUEL FLOW INDICATIONS

Details of the fuel flow indicator are included under Fuel System in a

later paragraph in this section.

7-56

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE INDICATING SYSTEM (EIS) (Continued)

OIL PRESSURE INDICATION

Engine oil pressure is shown by the OIL PSI horizontal indicator on the

ENGINE page. The indicator range is 0 to 130 PSI with a minimum

redline at 39 PSI, a amber band from 40 to 84 PSI (caution range), a

green band from 85 to 105 PSI (normal operating range) and a

maximum redline at 105 PSI. A white pointer indicates actual oil

pressure. Oil pressure is shown numerically above the horizontal

indicator.

When oil pressure is the normal operating range, indications 85 to 105

PSI, the OIL PSI label and pointer will remain white and digital value

will be green.

When oil pressure is the caution range, indications 40 to 84 PSI, the

OIL PSI label and pointer will turn amber and the digital value will be

amber background with black text.

When oil pressure is the warning range, indications 0 to 39 PSI or 106

to 130 PSI, the OIL PSI label and pointer will turn red and the digital

value will change to red background with white text to show that oil

pressure is outside normal limits.

The oil pressure transducer, connected to the accessory case oil

pressure port, provides a signal to the engine display that is processed

and shown as oil pressure. A separate low oil pressure switch causes

an OIL PRESS LOW annunciation on the PFD when oil pressure is 0 to

39 PSI. A red X through the oil pressure indicator means that the

indicating system is inoperative.

OIL TEMPERATURE INDICATION

Oil temperature (OIL °C) is displayed using a varied color tape and

digital display; the display can be 3 digits on the ENGINE page. The

instrument is operated by an electrical-resistance type temperature

sensor. Normal operation is indicated between 32 and 99°C; Amber

caution regions are indicated from -40 to 31°C and from 100 to 104°C.

Red lines are included at -41 and 105°C. Digits vary in color between

green, amber or red in correlation with the pointer and tape.

208BPHCUS-00

U.S.

7-57

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE LUBRICATION SYSTEM

The lubrication system consists of a pressure system, a scavenge

system and a breather system. The main components of the lubrication

system include an integral oil tank at the back of the engine, an oil

pressure pump at the bottom of the oil tank, an external double-element

scavenge pump located on the back of the accessory case, an internal

double-element scavenge pump located inside the accessory gearbox,

an oil-to-fuel heater located on the bottom of the accessory case, an oil

filter located internally on the right side of the oil tank, and an oil cooler

located on the right side of the nose cowl.

Oil is drawn from the bottom of the oil tank through a filter screen where

it passes through a pressure relief valve for regulation of oil pressure.

The pressure oil is then delivered from the main oil pump to the oil filter

where extraneous matter is removed from the oil and precluded from

further circulation. Pressure oil is then routed through passageways to

the engine bearings, reduction gears, accessory drives, torquemeter,

and propeller governor. Also, pressure oil is routed to the oil-to-fuel

heater where it then returns to the oil tank

After cooling and lubricating the engine moving parts, oil is scavenged

as follows:

1. Oil from the number 1 bearing compartment is returned by gravity

into the accessory gearbox.

2. Oil from the number 2 bearing is scavenged by the front element

of the internal scavenge pump back into the accessory gearbox.

3. Oil from the number 3 and number 4 bearings is scavenged by

the front element of the external scavenge pump into the

accessory gearbox.

4. Oil from the propeller governor, front thrust bearing, reduction

gear accessory drives, and torquemeter is scavenged by the rear

element of the external scavenge pump where it is routed through

a thermostatically-controlled oil cooler and then returned to the oil

tank.

5. The rear element of the internal scavenge pump scavenges oil

from the accessory case and routes it through the oil cooler

where it then returns to the oil tank.

(Continued Next Page)

208BPHCUS-01

U.S.

7-59

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE LUBRICATION SYSTEM (Continued)

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 oil return lines. A bypass valve,

immediately upstream of the front element of the internal scavenge

pump, vents the accessory gearbox when the engine is operating at

high power.

An oil dipstick/filler cap is located at the rear of the engine on the left

side and is accessible when the left side of the upper cowling is raised.

Markings which indicate U.S. quarts low if the oil is hot are provided on

the dipstick to facilitate oil servicing. The oil tank capacity is 9.5 U.S.

quarts and total system capacity is 14 U.S. quarts. For engine oil type

and brand, refer to Section 8.

FIREWALL OIL SHUTOFF VALVE

A firewall oil shutoff valve, located on the forward side of the firewall,

enables the pilot to shut off all oil flow from the engine to the oil cooler

in the event of an engine fire. With the oil cooler shutoff valve closed, oil

bypasses the oil cooler through an external oil line and relief valve and

dumps back into the engine. An additional check valve is placed in the

oil cooler return line to prevent bypassed oil from flowing back to the oil

cooler. The shutoff valve is controlled by a red push-pull knob labeled

FUEL/OIL SHUTOFF located on the right side of the control pedestal.

The push-pull knob has a press-to-release button in the center which

locks the knob in position when the button is released.

7-60

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

IGNITION SYSTEM

The ignition system consists of two igniters, an ignition exciter, two

high-tension leads, an ignition monitor light, an IGNITION switch, and a

STARTER switch. Engine ignition is provided by two igniters in the

engine combustion chamber. The igniters are energized by the ignition

exciter mounted on the engine mount on the right side of the engine

compartment. Electrical energy from the ignition exciter is transmitted

through two high-tension leads to the igniters in the engine. The ignition

system is normally energized only during engine start.

Ignition is controlled by an IGNITION switch and a STARTER switch

located on the left sidewall switch and circuit breaker panel. The

IGNITION switch has two positions, ON and NORM. The NORM

position of the switch arms the ignition system so that ignition will be

obtained when the STARTER switch is placed in the START position.

The NORM position is used during all ground starts and during air

starts with starter assist. The ON position of the switch provides

continuous ignition regardless of the position of the STARTER switch.

This position is used for operation on water-covered runways, during

flight in heavy precipitation, during inadvertent icing encounters until

the inertial separator has been in bypass for 5 minutes, and when near

fuel exhaustion as indicated by illumination of the red RSVR FUEL

LOW annunciator.

(Continued Next Page)

208BPHCUS-01

U.S.

7-61

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

IGNITION SYSTEM (Continued)

The main function of the starter switch is control of the starter for

rotating the gas generator portion of the engine during starting.

However, it also provides ignition during starting. For purposes of this

discussion, only the ignition functions of the switch are described. For

other functions of the starter switch, refer to paragraph titled Starting

System, in this section. The starter switch has three positions, OFF,

START, and MOTOR. The OFF position shuts off the ignition system

and is the normal position at all times except during engine start or

engine clearing. The START position energizes the engine ignition

system provided the ignition switch is in the NORMAL position. After

the engine has started during a ground or air start, the starter switch

must be manually positioned to OFF for generator operation.

White IGNITION ON annunciator will illuminate when electrical power is

being applied to the igniters. The ignition system is protected by a pull-

off type circuit breaker, labeled IGN, on the left sidewall switch and

circuit breaker panel.

AIR INDUCTION SYSTEM

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

ducts and an inertial separator system and then enters the engine

through a circular plenum chamber where it is directed to the

compressor by guide vanes. The compressor air inlet incorporates a

screen which will prevent entry of large articles, but does not filter the

inlet air.

7-62

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

INERTIAL SEPARATOR SYSTEM

An inertial separator system in the engine air inlet duct prevents

moisture particles from entering the compressor air inlet plenum when

in bypass mode. The inertial separator consists of two movable vanes

and a fixed airfoil which, during normal operation, route the inlet air

through a gentle turn into the compressor air inlet plenum. When

separation of moisture particles is desired, the vanes are positioned so

that the inlet air is forced to execute a sharp turn in order to enter the

inlet plenum. This sharp turn causes any moisture particles to separate

from the inlet air and discharge overboard through the inertial separator

outlet in the left side of the cowling.

Inertial separator operation is controlled by a T-handle located on the

lower instrument panel. The T-handle is labeled BYPASS-PULL,

NORMAL-PUSH. The inertial separator control should be moved to the

BYPASS position prior to running the engine during ground or flight

operation in visible moisture (clouds, rain, snow, ice crystals) with an

OAT of 5°C (41°F) or less. It may also be used for ground operations or

takeoffs from dusty, sandy field conditions to minimize ingestion of

foreign particles into the compressor. The NORMAL position is used for

all other operations.

The T-handle locks in the NORMAL position by rotating the handle

clockwise 1/4 turn to its vertical position. To unlock, push forward

slightly and rotate the handle 90° counterclockwise. The handle can

then be pulled into the BYPASS position. Once moved to the BYPASS

position, air loads on the movable vanes hold them in this position.

CAUTION

Do not return the INERTIAL SEPARATOR to NORMAL

until after engine shutdown and inspection if icing

conditions are encountered.

NOTE

When moving the inertial separator control from BYPASS to

NORMAL position during flight, reduction of engine power

will reduce the control forces.

(Continued Next Page)

208BPHCUS-00

U.S.

7-63

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

EXHAUST SYSTEM

The exhaust system consists of a primary exhaust pipe attached to the

right side of the engine just aft of the propeller reduction gearbox. A

secondary exhaust duct, fitted over the end of the primary exhaust pipe

carries the exhaust gases away from the cowling and into the

slipstream. The juncture of the primary exhaust pipe and secondary

exhaust duct is located directly behind the oil cooler. Since the

secondary exhaust duct is of larger diameter than the primary exhaust

pipe, a venturi effect is produced by the flow of exhaust. This venturi

effect creates a suction behind the oil cooler which augments the flow

of cooling air through the cooler. This additional airflow improves oil

cooling during ground operation of the engine.

ENGINE FUEL SYSTEM

The engine fuel system consists of an oil-to-fuel heater, an engine-

driven fuel pump, a fuel control unit, a flow divider and dump valve, a

dual fuel manifold with 14 simplex nozzles, and two fuel drain lines. The

system provides fuel flow to satisfy the speed and power demands of

the engine.

Fuel from the airplane reservoir is supplied to the oil-to-fuel heater

which utilizes heat from the engine lubricating oil system to preheat the

fuel in the fuel system. 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 engine oil tank.

Fuel from the oil-to-fuel heater then enters the engine-driven fuel pump

chamber 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 fuel pressure and delivers

it to the fuel control unit via a 10-micron filter in the pump outlet. A

bypass valve and cored passages in the pump body enables unfiltered

high pressure fuel to flow to the fuel control unit in the event the outlet

filter becomes blocked.

(Continued Next Page)

208BPHCUS-00

U.S.

7-65

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ENGINE FUEL SYSTEM (Continued)

The fuel control unit consists of a fuel metering section, a temperature

compensating section, and a gas generator (N_g) pneumatic governor.

The fuel control unit determines the proper fuel schedule to provide the

power required as established by the power lever input. This is

accomplished by controlling the speed of the compressor turbine. The

temperature compensating section alters the acceleration fuel schedule

to compensate for fuel density differences at different fuel

temperatures, especially during engine start. The power turbine

governor, located in the propeller governor housing, provides power

turbine overspeed protection in the event of propeller governor failure.

This is accomplished by limiting fuel to the gas generator. During

reverse thrust operation, maximum power turbine speed is controlled

by the power turbine governor. The temperature compensator alters the

acceleration fuel schedule of the fuel control unit to compensate for

variations in compressor inlet air temperature. Engine characteristics

vary with changes in inlet air temperature, and the acceleration fuel

schedule must, in turn, be altered to prevent compressor stall and/or

excessive turbine temperatures.

The flow divider schedules the metered fuel, from the fuel control unit,

between the primary and secondary fuel manifolds. The fuel manifold

and nozzle assemblies supply fuel to the combustion chamber through

10 primary and 4 secondary fuel nozzles, with the secondary nozzles

cutting in above a preset value. All nozzles are operative at idle and

above.

When the fuel cutoff valve in the fuel control unit closes during engine

shutdown, both primary and secondary manifolds are connected to a

dump valve port and residual fuel in the manifolds is allowed to drain

into the fuel can attached to the firewall where it can be drained daily.

COOLING SYSTEM

No external cooling provisions are provided for the PT6A-140 engine in

this installation. However, the engine incorporates an extensive internal

air system which provides for bearing compartment sealing and for

compressor and power turbine disk cooling. For additional information

on internal engine air systems, refer to the engine maintenance manual

for the airplane.

7-66

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

STARTING SYSTEM

The starting system consists of a starter-generator, a starter switch,

and an amber STARTER ON annunciator. The starter-generator

functions as a motor for engine starting and will motor the gas

generator section until a speed of 46% N_g is reached, at which time, the

start cycle will automatically be terminated by a speed sensing switch

located in the starter-generator. The starter-generator is controlled by a

three-positioned starter switch located on the left sidewall switch and

circuit breaker panel. The switch has OFF, START, and MOTOR

positions. The OFF position deenergizes the ignition and starter circuits

and is the normal position at all times except during engine start. The

START position of the switch energizes the starter-generator which

rotates the gas generator portion of the engine for starting. Also, the

START position energizes the ignition system, provided the ignition

switch is in the NORMAL position. When the engine has started, the

starter switch must be manually placed in the OFF position to de-

energize the ignition system and activate the generator system. The

MOTOR position of the switch motors the engine without having the

ignition circuit energized and is used for motoring the engine when an

engine start is not desired. This can be used for clearing fuel from the

engine, washing the engine compressor, etc. The MOTOR position is

spring-loaded to the OFF position. Also, 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. Starter contactor operation is indicated by an amber

STARTER ON annunciator.

ENGINE ACCESSORIES

All engine-driven accessories, with the exception of the propeller

tachometer-generator and the propeller governors, are mounted on the

accessory gearbox located at the rear of the engine. These

accessories are driven from the compressor turbine by a coupling shaft

which extends the drive through a conical tube in the oil tank center

section.

208BPHCUS-00

U.S.

7-67

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

OIL PUMP

Pressure oil is circulated from the integral oil tank through the engine

lubrication system by a self-contained, gear-type pressure pump

located in the lowest part of the oil tank. The oil pump is contained in a

cast housing which is bolted to the front face of the accessory

diaphram, and is driven by the accessory gear shaft. The oil pump body

incorporates a circular mounting boss to accommodate a check valve,

located in the end of the filter housing. A second mounting boss on the

pump accommodates a pressure relief valve.

FUEL PUMP

The engine-driven pump is mounted on the accessory gearbox at the 2

o’clock position. The 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. Another splined

coupling shaft extends the drive to the fuel control unit which is bolted

to the rear face of the pump. Fuel from the oil-to-fuel heater enters the

fuel pump through a 74-micron inlet screen. Then, fuel enters the pump

gear chamber, is boosted to high pressure, and delivered to the fuel

control unit through a 10-micron pump outlet filter. A bypass valve and

cored passages in the pump casing enable unfiltered high pressure fuel

to flow from the pump gears to the fuel control unit should the outlet

filter become blocked. An internal passage originating at the mating

face with the fuel control unit returns bypass fuel from the fuel control

unit to the pump inlet downstream of the inlet screen. A pressure

regulating valve in this line serves to pressurize the pump gear

bushings.

N_g TACHOMETER-GENERATOR

The N_g tachometer-generator produces an electric current which is

used in conjunction with the gas generator% RPM indicator to indicate

gas generator RPM. The N_g tachometer-generator drive and mount

pad is located at the 5 o’clock position on the accessory gearbox and is

driven from the internal scavenger pump.

7-68

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

PROPELLER TACHOMETER-GENERATOR

The propeller tachometer-generator produces an electric signal which

is used in conjunction with the propeller RPM indicator. The propeller

tachometer- generator drive and mount pad is located on the right side

of the reduction gearbox case and rotates clockwise with a drive ratio of

0.1273:1.

TORQUEMETER

The torquemeter is a hydro-mechanical torque measuring device

located inside the first stage reduction gear housing to provide an

accurate indication of engine power output. The difference between the

torquemeter pressure and the reduction gearbox internal pressure

accurately indicates the torque being produced. The two pressures are

internally routed to bosses located on the top of the reduction gearbox

front case and to a pressure transducer which is electrically connected

to the G1000 which indicates the correct torque. For standby indication,

the pressures are routed to bosses on the top of the reduction

gearcase front case and plumbed to the standby torque indicator.

STARTER-GENERATOR

The starter-generator is mounted on the top of the accessory case at

the rear of the engine. The starter-generator is a 28-volt, 200-amp

engine-driven unit that functions as a motor for engine starting and,

after engine start, as a generator for the airplane electrical system.

When operating as a starter, a speed sensing switch in the

starter-generator will automatically shut down the starter, thereby

providing overspeed protection and automatic shutoff. The

starter-generator is air cooled by an integral fan, ram air ducted from

the front of the engine cowling and on airplanes equipped with the 300

amp starter generator, ram air is also supplied from a NACA scoop

located on the nosewheel fairing.

208BPHCUS-00

U.S.

7-69

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

INTERSTAGE TURBINE TEMPERATURE SENSING

SYSTEM

The interturbine temperature sensing system is designed to provide the

operator with an accurate indication of engine operating temperatures

taken between the compressor and power turbines. The system

consists of twin leads, two bus bars, and eight individual

chromel-alumel 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 power turbine

vanes. The probe is secured to the boss by means of a floating,

threaded fitting which is part of the thermocouple probe assembly.

Shielded leads connect each bus bar assembly to a terminal block

which provides a connecting point for external leads to the ITT indicator

in the airplane cabin.

PROPELLER GOVERNOR

The propeller governor is located in the 12 o’clock position on the front

case of the reduction gearbox. Under normal conditions, the governor

acts as a constant speed unit, maintaining the propeller speed selected

by the pilot by varying the propeller blade pitch to match the load to the

engine torque. The propeller governor also has a power turbine

governor section built into the unit. Its function is to protect the engine

against a possible power turbine overspeed in the event of a propeller

governor failure. If such an overspeed should occur, a governing orifice

in the propeller governor is opened by flyweight action to bleed off

compressor discharge pressure through the governor and computing

section of the fuel control unit. When this occurs, compressor discharge

pressure, acting on the fuel control unit governor bellows, decreases

and moves the metering valve in a closing direction, thus reducing fuel

flow to the flow divider.

7-70

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

TORQUE LIMITER

The torque limiter installed on the engine is a mechanical back-up unit

which prevents unintentional engine overtorques. On the PT6A-140

engine it limits the maximum torque value to 2500 FT-LB regardless of

propeller RPM.

The backup unit incorporates an oil bellows which senses torquemeter

oil pressure and is linked to a Py bleed orifice. Oil from the torquemeter

chamber passes through a restrictor before entering the bellows. The

restrictor dampens torque pressure fluctuation and prevents damage to

the bellows assembly. When torque pressure reaches 2500 FT-LB, the

bellows expands and compresses the spring.

Bimetallic disks are mounted on the spring to compensate for variation

of spring tension caused by change in ambient temperature.

The flapper valve then moves to allow Py air pressure from the fuel

control unit to vent to the atmosphere and therefore limit the fuel supply

to the engine reducing the engine speed and subsequently reducing

torque.

CAUTION

The pilot is always responsible for operating the engine

torque within limits and not depend on the mechanical

torque limiter.

208BPHCUS-00

U.S.

7-71

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

PROPELLER OVERSPEED GOVERNOR

This propeller overspeed governor is located at the 10 o’clock position

on the front case of the reduction gearbox. The governor acts as a

safeguard against propeller overspeed should the primary propeller

governor fail. The propeller overspeed governor regulates the flow of oil

to the propeller pitch-change mechanism by means of a flyweight and

speeder spring arrangement similar to the primary propeller governor.

Because it has no mechanical controls, the overspeed governor is

equipped with a test solenoid that resets the governor below its normal

overspeed setting for ground test. The OVERSPEED GOVERNOR

PUSH TO TEST Switch is located on the left side of the instrument

panel. For a discussion of this switch, refer to the paragraph titled

Propellers in this section.

ENGINE FIRE DETECTION SYSTEM

The engine fire detection system consists of a heat sensor in the

engine compartment, a red ENGINE FIRE annunciator located on the

PFD, and a warning horn above the pilot. The heat sensor consists of

three flexible closed loops. When high engine compartment

temperatures are experienced, the heat causes a change in resistance

in the closed loops. This change in resistance is sensed by a control

box, located on the aft side of the firewall, which will illuminate the red

ENGINE FIRE annunciator and trigger the audible warning horn. Fire

warning is initiated when temperatures in the engine compartment

exceed 425ºF (218ºC) on the first section (firewall), 625ºF (329ºC) on

the second section (around the exhaust), or 450ºF (232ºC) on the third

section (rear engine compartment).

A test switch, labeled TEST SWITCH, FIRE DETECT - UP, is located

on the lower left corner of the instrument panel. When this switch is

placed in the UP position, the red ENGINE FIRE annunciator will

illuminate on the CAS system and the warning horn will sound

indicating that the fire warning circuitry is operational. The system is

protected by a pull-off type circuit breaker, labeled FIRE DET, on the left

sidewall switch and circuit breaker panel.

7-72

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ENGINE GEAR REDUCTION SYSTEM

The reduction gear and propeller shaft, located in the front of the

engine, are housed in two magnesium alloy castings which are bolted

together at the exhaust outlet. The gearbox contains a two-stage

planetary gear train, 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 second

stage reduction gear, and then to the propeller shaft. The reduction

ratio is from a maximum power turbine speed of 33,000 RPM down to a

propeller speed of 1900 RPM.

The accessories, located on the front case of the reduction gearbox,

are driven by a bevel gear mounted at the rear of the propeller shaft

thrust bearing assembly. Drive shafts from the bevel drive gear transmit

rotational power to the three pads which are located at the 12, 3 and 9

o'clock positions. Propeller thrust loads are absorbed by a flanged ball

bearing assembly located on the front face of the reduction gearbox

center bore. The bevel drive gear adjusting spacer, thrust bearing, and

seal runner are stacked and secured to the propeller shaft by a key

washer and spanner nut. A thrust bearing cover assembly is secured

by bolts at the front flange of the reduction gearbox front case.

208BPHCUS-00

U.S.

7-73

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CHIP DETECTORS

Two chip detectors are installed on the engine, one on the underside of

the reduction gearbox case and one on the underside of the accessory

gearbox case. The chip detectors are installed to trigger an amber

CHIP DETECT annunciations anytime metal chips are present in one

or both of the chip detectors. Illumination of the amber CHIP DETECT

annunciator indicates the need for engine inspection for abnormal

wear. The amber CHIP DETECT annunciation will also be shown if

either chip detectors electrical connector has come loose. The engine

oil must be drained prior to removing either of the chip detector

sensors. Refer to the 208 Maintenance Manual, for more information

on inspection and removal of the engine chip detectors.

OIL BREATHER DRAIN CAN

Model 208 airplanes have an oil breather drain 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 (if installed), and

the propeller shaft seal. This can should be drained after every flight. A

drain valve on the bottom right side of the engine cowling enables the

pilot to drain the contents of the oil breather drain can into a suitable

container. The allowable quantity of oil discharge per hour of engine

operation is 14 cc for airplanes with air conditioning and 11 cc for

airplanes without air conditioning. If the quantity of oil drained from the

can is greater than specified, the source of the leakage should be

identified and corrected prior to further flight.

7-74

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

PROPELLER

The airplane is equipped standard with a Hartzell aluminum material,

three-bladed propeller. The propeller is constant-speed, full-feathering,

reversible, single-acting, governor-regulated propeller. A setting

introduced into the governor with the PROP RPM lever establishes the

propeller speed. The propeller utilizes oil pressure which opposes the

force of springs and counter-weights to obtain correct pitch for the

engine load. Oil pressure from the propeller governor drives the blades

toward low pitch (increases RPM) while the springs and counterweights

drive the blades toward high pitch (decreasing RPM). The source of oil

pressure for propeller operation is furnished by the engine oil system,

boosted in pressure by the governor gear pump, and supplied to the

propeller hub through the propeller flange.

To feather the propeller blades, the PROP RPM lever on the control

pedestal is placed in the FEATHER position; counterweights and spring

tension will continue to twist the propeller blades through high pitch and

into the streamlined or feathered position. Unfeathering the propeller is

accomplished by positioning the PROP RPM lever forward of the

feather gate. The unfeathering system uses engine oil pressure to force

the propeller out of feather.

Reversed propeller pitch is available for decreasing ground roll during

landing. To accomplish reverse pitch, the power lever is retarded

beyond IDLE and well into the BETA range. Maximum reverse power is

accomplished by retarding the power lever to the MAX REVERSE

position which increases power output from the gas generator and

positions the propeller blades at full reverse pitch. An externally

grooved feedback ring is provided with the propeller.

Motion of the feedback ring is proportional to propeller blade angle, and

is picked up by a carbon block running in the feedback ring. The

relationship between the axial position of the feedback ring and the

propeller blade angle is used to maintain control of blade angle from

idle to full reverse.

CAUTION

The propeller reversing linkage can be damaged if the

power lever is moved aft of the idle position when the

propeller is feathered.

208BPHCUS-00

U.S.

7-75

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

PROPELLER (Continued)

OVERSPEED GOVERNOR TEST SWITCH

An overspeed governor test switch is located on the left side of the

instrument panel. The switch is the push-to-test type and is used to test

the propeller overspeed governor during engine run-up. The switch,

when depressed, actuates a solenoid on the propeller overspeed

governor which restricts propeller RPM when the power lever is

advanced. To check for proper operation of the overspeed governor,

during engine run-up, depress the press-to-test switch and advance the

power lever until propeller RPM stabilizes; propeller RPM should not

exceed 1750 +/- 60 RPM.

FUEL SYSTEM

The airplane fuel system (see Fuel System figure) consists of 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, and 14 fuel nozzle

assemblies. An ecology tank is also provided to collect unused fuel

from the fuel manifold during engine shutdown. This unused fuel is

returned to the engine through an ejector pump during the next engine

start. Refer to the Fuel Quantity Data Chart for information pertaining to

this system.

WARNING

Unusable fuel levels for this airplane were

determined in accordance with Federal Aviation

Regulations. Failure to operate the airplane in

compliance with the fuel limitations specified in

Section 2 may further reduce the amount of fuel

available in flight.

Fuel flows from the tanks through the two fuel tank shutoff valves at

each tanks. The fuel tank shutoff valves are mechanically controlled by

two fuel selectors, labeled LEFT, ON and OFF, located on the overhead

panel. By manipulating the fuel selectors, the pilot can select either left

or right fuel tanks or both at the same time. Normal operation is with

both tanks on. Fuel flows by gravity from the shutoff valves in each tank

to the fuel reservoir.

(Continued Next Page)

7-76

U.S.

208BPHCUS-01

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

FUEL SYSTEM (Continued)

The reservoir is located at the low point in the fuel system which

maintains a head of fuel around the ejector boost pump and auxiliary

boost pump which are contained within the reservoir. This head of fuel

prevents pump cavitation in low-fuel quantity situations, especially

during in-flight maneuvering. 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, which is driven

by motive fuel flow from the fuel control unit, normally provides fuel flow

when the engine is operating. In the event of failure of the ejector boost

pump, the electric boost pump will automatically turn on, thereby

supplying fuel flow to the engine. The auxiliary boost pump is also used

to supply fuel flow during starting. Fuel in the reservoir manifold then

flows through a fuel/oil shutoff valve located on the aft side of the

firewall. This shutoff valve enables the pilot to cut off all fuel to the

engine.

After passing through the shutoff valve, fuel is routed through a fuel

filter located on the right side of the engine. The fuel filter incorporates

a bypass feature which allows fuel to bypass the filter in the event the

filter becomes blocked with foreign material. Fuel from the filter is then

routed through the oil-to-fuel heater to the engine-driven fuel pump

where fuel is delivered under pressure to the fuel control unit. The fuel

control unit meters the fuel and directs it to the flow divider which

distributes the fuel to dual manifolds and 14 fuel nozzles located in the

combustion chamber. For additional details concerning the flow of fuel

at the engine, refer to the Engine Fuel System paragraph in this

section.

Fuel rejected by the engine during shutdown drains into the ecology

tank located on the front left side of the firewall. This unused fuel is then

drawn back into the fuel system through an ejector pump that uses flow

from the motive flow line and when the motive flow shutoff valve is

reenergized.

Fuel system venting is essential to system operation. 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, one from each fuel tank, which protrude from the

trailing edge of the wing at the wing tips. also the fuel reservoir is

vented to both wing tanks.

208BPHCUS-01

U.S.

7-77

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

FUEL SYSTEM (Continued)

FUEL QUANTITY DATA

FUEL LEVEL

TOTAL

UNITS OF

(QUANTITY EACH TOTAL

TOTAL

USABLE ALL

MEASURE

TANK)

FUEL UNUSABLE

FLT

CONDITIONS

FULL

(OUTBOARD

FILLERS)

POUNDS

1124.25

2272

24.1

2246.5

GALLONS

(U.S.)

167.8

339.1

3.6

335.3

NOTE

Pounds are based on a fuel specific weight of 6.7 pounds

per U.S. gallon.

WARNING

To achieve full capacity, fill fuel tank to the top of

the fuel filler neck. Filling fuel tanks to the bottom

of the fuel filler collar (level with the flapper valve)

allows space for thermal expansion and results in a

decrease in fuel capacity of four gallons per side

(eight gallons total).

Figure 7-11

208BPHCUS-00

U.S.

7-79

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

FUEL SYSTEM (Continued)

FIREWALL FUEL SHUTOFF VALVE

A firewall fuel shutoff valve, located on the aft side of the firewall,

enables the pilot to shut off all fuel flow from the fuel reservoir to the

engine. The shutoff valve is controlled by a red push-pull knob labeled

FUEL/OIL SHUTOFF located on the right side of the control pedestal.

The push-pull knob has a press-to-release button in the center which

locks the knob in position when the button is released.

FUEL TANK SELECTORS

Two FUEL SELECTORS, one for each tank, are located on the

overhead console. The selectors, labeled LEFT, ON and OFF and

RIGHT, ON and OFF, mechanically control the position of the two fuel

tank shutoff valves at each wing tank. When a FUEL TANK SELECTOR

is in the OFF position, the shutoff valves in the tank are closed. When in

the ON position, both shutoff valves in the tank are open, allowing fuel

from that tank to flow to the reservoir. Normal fuel management is with

both FUEL TANK SELECTORS are in the ON position.

Before refueling, or when the airplane is parked on a slope, turn off one

of the FUEL TANK SELECTORS (if parked on a slope, turn high wing

tank off). This action prevents crossfeeding from the fuller or higher

tank and reduces any fuel seepage tendency from the wing tank vents.

7-80

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

FUEL SYSTEM (Continued)

FUEL SELECTORS OFF WARNING SYSTEM

A fuel selectors off warning system is incorporated to alert the pilot if

one or both of the FUEL TANK SELECTORS are in the OFF position

inadvertently. The system included redundant warning horns, a red

FUEL SELECT OFF annunciation of the selected tank, actuation

switches, and miscellaneous electrical hardware. The dual aural

warning system is powered through the START CONT circuit breaker

with a non-pullable FUEL SEL WARN circuit breaker installed in series

to protect the integrity of the start system.

The warning system functions as follows:

1. If both the LEFT and RIGHT FUEL TANK SELECTORS are in

the OFF position (fuel tank shutoff valves are closed), the red

FUEL SELECT OFF annunciator illuminates and one of the fuel

selector off warning horns is activated;

2. During an engine start operation (STARTER switch in START or

MOTOR position) with either the left or right fuel tank selectors

in the OFF position, the red FUEL SELECT OFF annunciator

illuminates and both of the fuel select off warning horns are

activated;

3. With one fuel tank selector in the OFF position and fuel

remaining in the tank being used is less than approximately 25

gallons, the red FUEL SELECT OFF annunciator illuminates

and one of the fuel selector off warning horns is activated.

The warning system has the ability to annunciate which fuel selector is

selected off by displaying OFF next to the respective fuel quantity

indicator on the EIS. There is no annunciation when the fuel selector is

turned ON.

If the FUEL SEL WARN circuit breaker has popped or the START

CONT circuit breaker has been pulled

(possible for ground

maintenance), the red FUEL SELECT OFF annunciator will be

illuminated even with both fuel tank selectors ON. This is a warning to

the pilot that the fuel selector warning system has been deactivated.

A test switch, labeled TEST SWITCH, FUEL SELECT OFF - DN, is

located on the lower left corner of the instrument panel. When this

switch is placed in the DOWN position, the two warning horns will

sound simultaneously indicating that the fuel selector warning horns

are operational.

208BPHCUS-00

U.S.

7-81

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

FUEL SYSTEM (Continued)

FUEL BOOST PUMP SWITCH

An auxiliary boost pump switch, located on the left sidewall switch and

circuit breaker panel, is labeled FUEL BOOST and has OFF, NORM,

and ON positions. When the FUEL BOOST switch is in the OFF

position, the auxiliary boost pump is inoperative. When the FUEL

BOOST switch is in the NORM position, the auxiliary boost pump is

armed and will operate when fuel pressure in the fuel manifold

assembly drops below 2.5 psi. The NORM position is used for all

normal engine operation where main fuel flow is provided by the ejector

boost pump and the auxiliary boost pump is used as a standby. When

the FUEL BOOST switch is placed in the ON position, the auxiliary

boost pump will operate continuously and the motive flow pump will be

shut off. The ON position is used for engine start and any other time

that the auxiliary boost pump cycles on and off with the switch in the

NORM position due to low fuel pressure.

The high pressure motive flow shutoff valve needs to be closed (FUEL

BOOST switch ON) during engine starts so there is sufficient fuel

pressure delivered to the Fuel Control Unit to open the minimum

pressurizing valve to allow fuel flow to the fuel nozzles. The fuel boost

pump incorporates a timer so when the fuel boost pump switch is

moved from ON to NORM or OFF the pump will continue to run for a

short period of time. This is to avoid the motive flow shutoff valve from

opening and the pump switching off at the same time. This prevents

potential low fuel pressure to the engine causing the boost pump to

cycle back on.

NOTE

If the FUEL BOOST pump switch is not turned ON prior to

engine start, (motive flow shutoff valve not energized) there

will be insufficient fuel pressure to open the minimum

pressurizing valve and the engine will not start.

7-82

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

FUEL SYSTEM (Continued)

FUEL FLOW INDICATION

A fuel flow indicator, located beneath the quantity indicators on both the

ENGINE and SYSTEM pages, indicates the fuel consumption of the

engine in pounds per hour based on Jet A fuel. The indicator measures

the flow of fuel downstream of the fuel control unit just before being

routed into the flow divider. The fuel flow indicator receives power from

a pull-off type circuit breaker labeled F FLOW NP & NG, on the left

sidewall switch and circuit breaker panel.

208BPHCUS-00

U.S.

7-83

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

FUEL QUANTITY INDICATIONS

Fuel quantity is measured by four fuel quantity transmitters (two in each

tank) and indicated in the EIS section on the MFD below the oil

temperature indicators. The FUEL QTY indicators utilize twin vertical

scales that show fuel quantity for both the L and R tanks in LBS from 0

to 1000 in 200 pound increments. The fuel quantity indicators, which

measure volume, are calibrated in pounds (based on the weight of Jet

A fuel on a standard day). An empty tank is indicated by a red line.

When an indicator shows an empty tank, approximately 2.8 gallons

remain in the tank as unusable fuel. The left and right fuel level senders

each receive power from a pull-off type circuit breaker. The breakers

are labeled LEFT FUEL QTY and RIGHT FUEL QTY, respectively, and

are located on the left sidewall switch and circuit breaker panel.

WARNING

Because of the relatively long fuel tanks, fuel

quantity indicator accuracy is affected by

uncoordinated flight or a sloping ramp if reading

the indicators while on the ground. Therefore, to

obtain accurate fuel quantity readings, verify that

the airplane is parked in a laterally level condition,

or if in flight, make sure the airplane is in a

coordinated and stabilized condition for at least 1

minute.

WING TANK FUEL LOW CAUTION ANNUNCIATORS

Two float sensors one for each wing tank will trigger the appropriate

amber L FUEL LOW or R FUEL LOW annunciation when the fuel in the

respective tank is 25 gallons (170 lbs) or less. When the fuel quantity in

each tank is less than 25 gallons (170 lbs), amber L-R FUEL LOW

annunciator will replace the previously displayed L or R FUEL LOW

annunciator.

7-84

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

RESERVOIR FUEL LOW WARNING ANNUNCIATOR

A red RSVR FUEL LOW annunciator is located on the PFD, and will

come on when the level of fuel in the reservoir drops to approximately

one-half full. With the fuel reservoir full, there is adequate fuel for

approximately

3 minutes of maximum continuous power or

approximately 9 minutes at idle power.

FUEL PRESSURE LOW WARNING ANNUNCIATOR

An amber FUEL PRESS LOW annunciator is located on the PFD, and

will illuminate when fuel pressure drops below 2.5 psi.

FUEL BOOST PUMP ON ANNUNCIATOR

An amber FUEL BOOST ON annunciator is located on the PFD and will

come on when the electric boost pump is operating, such as when the

FUEL BOOST pump switch is placed in the ON position or when the

FUEL BOOST pump switch is in the NORM position and fuel pressure

drops below 2.5 psi.

208BPHCUS-00

U.S.

7-85

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

DRAIN VALVES

The fuel system is equipped with drain valves to provide a means for

the examination of fuel in the system for contamination and grade.

Drain valves are located on the lower surface of each wing at the

inboard end of the fuel tank, in fuel tank external sumps, on the left side

of the cargo pod for the reservoir tank, and on the underside of the fuel

filter. Outboard fuel tank drain valves and their use is recommended if

the airplane is parked with one wing low on a sloping ramp. The drain

valves for the wing tanks and their external sumps are tool-operated

poppet type and are flush-external mounted. The wing tank and

external sump drain valves are constructed so that the phillips

screwdriver on the fuel sampler which is provided can be utilized to

depress the valve and then twist to lock the drain valve in the open

position. The drain valve for the reservoir is controlled by a double-

button push-pull drain control knob. 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 upward to drain fuel from the filter. The fuel sampler

can be used in conjunction with these drain valves for fuel sampling

and purging of the fuel system. The fuel tanks should be filled after

each flight when practical to minimize condensation.

Before each flight of the day and after each refueling, use a clear

sampler and drain fuel from the inboard fuel tank sump, external sump

quick-drain valves, fuel reservoir quick-drain valve, and fuel filter quick-

drain valve to determine if contaminants are present, and that the

airplane has been fueled with the proper fuel. If the airplane is parked

with one wing low on a sloping ramp, draining of the outboard fuel tank

sump quick-drain valves is also recommended. If contamination is

detected, drain all fuel drain points again. Take repeated samples from

all fuel drain points until all contamination has been removed. If after

repeated sampling, evidence of contamination still exists, the fuel tanks

should be completely drained and the fuel system cleaned. Do not fly

the airplane with contaminated or unapproved fuel.

WARNING

JP-4 and other NAPHTHA based fuels can cause

severe skin and eye irritation.

7-86

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

FUEL ECOLOGY TANK

A fuel ecology tank is supplied with the engine and is mounted on the

engine side of the firewall, capturing fuel drained from the fuel nozzle

manifolds at engine shutdown. When the engine is restarted, motive

flow fuel feeds an ejector pump in the ecology tank, which picks up the

discharged fuel and recirculates it back into the fuel system.

FUEL PUMP DRAIN RESERVOIR

To control expended lubricating oil from the engine fuel pump drive

coupling area and provide a way to determine if fuel is leaking past the

fuel pump seal, this airplanes is equipped with a drainable reservoir to

collect this 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 an interval not to exceed six engine

shutdowns. A drain valve on the bottom side of the cowling enables the

pilot to drain the contents of the reservoir into a suitable container. A

quantity of up to 3 cc of oil and 20 cc of fuel discharge per hour of

engine operation is allowable. If the quantity of oil or fuel drained from

the reservoir is greater than specified, the source of leakage should be

identified and corrected prior to further flight.

208BPHCUS-00

U.S.

7-87

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

BRAKE SYSTEM

The airplane has a single-disc, hydraulically-actuated brake on each

main landing gear wheel. Each brake is connected, by a hydraulic line,

to a master cylinder attached to each of the pilot's rudder pedals. The

brakes are operated by applying pressure to the top of either the left

(pilot's) or right

(copilot's) set of rudder pedals, which are

interconnected. When the airplane is parked, both main wheel brakes

may be set by utilizing the parking brake which is operated by a handle

below to the right of the pilot’s control wheel. To apply the parking

brake, set the brakes with the rudder pedals and pull the handle aft. To

release the parking brake, push the handle fully in.

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 checked for

proper level prior to each flight.

For maximum brake life, keep the brake system properly maintained.

Airplanes are equipped with metallic type brakes, and require a special

brake burn-in before delivery

(or after brake replacement). When

conditions permit, hard brake application is beneficial in that the

resulting higher brake temperatures tend to maintain proper brake

glazing and will prolong the expected brake life. Conversely, the

habitual use of light and conservative brake application is detrimental to

metallic brakes.

Some of the symptoms of impending brake failure are: gradual

decrease in braking action after brake application, noisy or dragging

brakes, soft or spongy pedals, and excessive travel and weak braking

action. If any of these symptoms appear, the brake system is in need of

immediate attention. If, during taxi or landing roll, braking action

decreases, let up on the pedals and then re-apply the brakes with

heavy pressure. If the brakes become spongy or pedal travel

increases, pumping the pedals should build braking pressure. If one

brake becomes weak or fails, use the other brake sparingly while using

opposite rudder, as required, to offset the good brake.

7-88

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ELECTRICAL SYSTEM

The airplane is equipped with a 28-volt, direct-current electrical system,

refer to Figure 7-12, Electrical System. The system uses a 24-volt

sealed lead acid battery; located on the front right side of the firewall,

as a source of electrical energy. A 200-amp engine-driven starter-

generator is used to maintain the battery's state of charge. Power is

supplied to most general electrical and all avionics circuits through two

general buses, two avionics buses, and a hot battery bus. The battery

bus is energized continuously for cabin/courtesy lights and functions

requiring power when the two general buses are off. The two general

buses are on anytime the BATTERY switch is turned ON. All DC buses

are on anytime the BATTERY switch and the two AVIONICS switches

are turned ON.

STANDBY ELECTRICAL SYSTEM

The standby electrical system serves as a power source in the event

the main generator system malfunctions in flight. The system includes

an alternator operated at a 75-amp capacity rating. The alternator is

belt-driven from an accessory pad on the rear of the engine. The

system also includes an alternator control unit located forward of the

circuit breaker panel, a standby alternator contactor assembly on the

left front side of the firewall and two switches on the left sidewall switch

panel, labeled STBY ALT PWR and AVIONICS STBY PWR.

Circuit protection and isolation is provided by two circuit breakers,

labeled STBY PWR, on the left sidewall circuit breaker panel. Field

excitation to the alternator control unit is supplied through diode logic

from a circuit breaker in the standby alternator relay assembly or from

the HOURMETER/ACU circuit breaker in the main power relay box.

Standby electrical system monitoring is provided by annunciators; white

STBY PWR ON annunciator and amber STBY PWR INOP annunciator.

Total amperage supplied from the standby electrical system can be

monitored on the EIS SYSTEMS DISPLAY. Additionally, an amber

ALTNR AMPS annunciator is provided if the standby alternator

amperage draw exceeds normal operating ranges.

208BPHCUS-00

U.S.

7-89

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

GENERATOR CONTROL UNIT

The generator control unit (GCU) is mounted inside the cabin on the left

forward fuselage sidewall. The unit provides the electrical control

functions necessary for the operation of the starter-generator. The GCU

provides for automatic starter cutoff when engine RPM is above 46%.

Below 46%, the starter-generator functions as a starter, and above

46%, the starter-generator functions as a generator when the

STARTER switch is OFF. The GCU provides voltage regulation plus

high voltage protection and reverse current protection. In the event of a

high-voltage or reverse current condition, the generator is automatically

disconnected from the buses. The generator contactor (controlled by

the GCU) connects the generator output to the airplane bus. If any

GCU function causes the generator contactor to de-energize, the

amber GENERATOR OFF annunciator will illuminate.

GROUND POWER MONITOR

The ground power monitor is located inside the electrical power control

assembly mounted on the left hand side of the firewall in the engine

compartment. This unit senses the voltage level applied to the external

power receptacle and will close the external power contactor when the

applied voltage is within the proper limits.

BATTERY SWITCH

The BATTERY switch is a two-position toggle-type switch, labeled

BATTERY, and is located on the left sidewall switch and circuit breaker

panel. The BATTERY switch is ON in the forward position and OFF in

the aft position. When the BATTERY switch is in the ON position,

battery power is supplied to the two general buses. The OFF position

removes power to all buses except the battery bus.

7-90

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

STARTER SWITCH

The STARTER switch is a three-position toggle-type switch, labeled

STARTER, on the left sidewall switch and circuit breaker panel. The

switch has OFF, START, and MOTOR positions. For additional details

of the STARTER switch, refer to the starting system paragraph in this

section.

IGNITION SWITCH

The IGNITION switch is a two-position toggle-type switch, labeled

IGNITION, on the left sidewall switch and circuit breaker panel. The

switch has ON and NORMAL positions. For additional details of the

IGNITION switch, refer to the ignition system paragraph in this section.

GENERATOR SWITCH

The GENERATOR switch is a three-position toggle-type switch, labeled

GENERATOR, on the left sidewall switch and circuit breaker panel. The

switch has ON, RESET, and TRIP positions. With the switch in the ON

position, the GCU will automatically control the generator line contactor

for normal generator operation. The RESET and TRIP positions are

momentary positions and are spring-loaded to the ON position. If a

momentary fault should occur in the generating system (as evidenced

by the amber GENERATOR OFF annunciator, red VOLTAGE LOW

annunciator and/or red VOLTAGE HIGH annunciator), the

GENERATOR switch can be momentarily placed in the RESET position

to restore generator power. If erratic operation of the generating system

is observed, the system can be shutoff by momentarily placing the

GENERATOR switch to the TRIP position. After a suitable waiting

period, generator operation may be recycled by placing the

GENERATOR switch momentarily to RESET.

208BPHCUS-00

U.S.

7-91

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

STANDBY ALTERNATOR POWER SWITCH

The STBY ALT PWR switch is a two-position toggle-type switch,

labeled STBY ALT PWR. There is also an amber LED light located

above the switch that comes on when the BATTERY switch is in the

OFF position with STBY ALT PWR switch in the ON position. This is an

alert to the operator to help prevent accidental discharging of the

battery that can occur if the STBY ALT PWR switch is left ON after

shutdown.

AVIONICS POWER SWITCHES

Electrical power from the airplane power distribution bus to the avionics

buses, refer to Figure 7-12, Electrical System, is controlled by two

toggle-type switch/breakers located on the left sidewall switch and

circuit breaker panel. One switch controls power to the No. 1 avionics

bus while the other switch controls power to the No. 2 avionics bus. The

switches are labeled AVIONICS 1 and 2 and are ON in the forward

position and OFF in the aft position. The AVIONICS power switches

should be placed in the OFF position prior to turning the BATTERY

switch ON or OFF, or applying an external power source. The

AVIONICS No. 1 Switch must be ON prior to engine start to display EIS

information.

AVIONICS STANDBY POWER SWITCH

The AVIONICS STBY PWR switch is a guarded two-position switch/

breaker, labeled AVIONICS STBY PWR. The guard covering this

switch must be lifted in order to select the ON position. When switched

ON, the standby electrical system directly provides power to the AVN

BUS 1. When switched OFF, the standby electrical system may provide

extra power to the avionics buses via the main power distribution bus,

provided the STBY PWR circuit breakers on the electrical buses are not

pulled. When operating solely on standby power, both AVIONICS No. 1

and No. 2 power switches should be OFF to avoid feeding a possible

fault in the primary power system.

7-92

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

AVIONICS BUS TIE SWITCH

The AVIONICS BUS TIE switch is a two-position guarded toggle-type

switch located on the left sidewall switch and circuit breaker panel. The

switch connects the AVN BUS 1 and AVN BUS 2 together in the event

of failure of either bus feeder circuit. Because power for each avionics

bus is supplied from a separate current limiter on the power distribution

bus, failure of a current limiter can cause failure of the affected bus.

Placing the AVIONICS 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, depending on equipment installed.

208BPHCUS-00

U.S.

7-93

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

EXTERNAL POWER SWITCH

The EXTERNAL POWER switch is a three-position guarded toggle-

type switch located on the left sidewall switch and circuit breaker panel.

The switch has OFF, STARTER, and BUS positions and is guarded in

the OFF position. When the switch is in the OFF position, battery power

is supplied to the main bus and to the starter-generator circuit, external

power cannot be applied to the main bus, and, with the generator

switch in the ON position, power is applied to the generator control

circuit. When the EXTERNAL POWER switch is in the STARTER

position, external power is applied to the starter circuit only and battery

power is supplied to the main bus. No generator power is available in

this position. When the EXTERNAL POWER switch is in the BUS

position, external power is applied to the main bus and no power is

available to the starter. The battery can be connected to the main bus

with external power connected to the airplane by placing the BATTERY

switch to the ON position. The battery charge must be monitored to

prevent a overcharge condition.

CIRCUIT BREAKERS

Most of the electrical circuits in the airplane are protected by pull-off

type circuit breakers mounted on the left sidewall switch and circuit

breaker panel. Should an overload occur in any circuit, the controlling

circuit breaker will trip, opening the circuit. After allowing the circuit

breaker to cool for approximately three minutes, it may be reset

(pushed in). If the breaker trips again, it should not be reset until

corrective action is taken.

WARNING

Make sure all circuit breakers are in before all

flights. Never operate with tripped circuit breakers

without a thorough knowledge of the

consequences.

208BPHCUS-00

U.S.

7-97

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

ELECTRICAL SYSTEM (Continued)

VOLTAGE AND AMPERAGE DISPLAY

The status of the electrical system can be monitored on the MFD (non-

reversionary mode). Battery current (BAT AMPS) and bus voltage

(BUS VOLTS) are displayed on the default EIS-ENGINE display page.

By pressing the ENGINE softkey and the SYSTEM softkey, the EIS

pages changes to the EIS-Systems display where generator current

(GEN AMPS), and bus voltage (BUS VOLTS) can be monitored

simultaneously. A negative display on BAT AMPS indicates battery

discharge, while a positive display indicates battery charging. A

negative display on BAT AMPS will be shown in amber to remind the

pilot to reduce electrical load, or increase Ng, to maintain battery

charge.

GROUND SERVICE PLUG RECEPTACLE

A ground service plug receptacle permits the use of an external power

source for cold weather starting and during lengthy maintenance work

on the electrical and avionics equipment. External power control

circuitry is provided to prevent the external power and the battery from

being connected together during starting. The external power

receptacle is installed on the left side of the engine compartment near

the firewall.

The ground service circuit incorporates polarity reversal and

overvoltage protection. Power from the external power source will flow

only if the ground service plug is correctly connected to the airplane. If

the plug is accidentally connected backwards or the ground service

voltage is too high, no power will flow to the electrical system, thereby

preventing any damage to electrical equipment.

7-98

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

LIGHTING SYSTEMS

EXTERIOR LIGHTING

Exterior lighting consists of three navigation lights, two landing lights,

two taxi/recognition lights, two strobe lights, a flashing beacon, two

underwing courtesy lights and one wing inspection light. All exterior

lights are controlled by toggle switches located on the lighting control

panel on the left side of the instrument panel. The toggle switches are

ON in the up position and OFF in the down position. All exterior lights

are LED.

NAVIGATION LIGHTS

LED navigation lights are installed on the wing tips along with rear-

facing wing LED navigation lights that replace the traditional tailcone

white navigation light. The lights are protected by a pull-off type circuit

breaker, labeled NAV LIGHT, on the left sidewall switch and circuit

breaker panel.

LANDING LIGHTS

Two LED landing lights are installed on the airplane, one in each wing

leading edge mounted outboard of each taxi light. The lights provide

illumination forward and downward during takeoff and landing. The

lights are protected by two pull-off type circuit breakers, labeled LEFT

LDG LIGHT and RIGHT LDG LIGHT, on the left sidewall switch and

circuit breaker panel.

TAXI/RECOGNITION LIGHTS

Two LED taxi/recognition lights are mounted inboard of each landing

light in each wing leading edge. The lights are focused to provide

illumination of the area forward of the airplane during ground operation

and taxiing. The lights are also used to enhance visibility of the airplane

in the traffic pattern or enroute. The taxi/recognition lights are protected

by a pull-off type circuit breaker, labeled TAXI LIGHT, on the left

sidewall switch and circuit breaker panel.

208BPHCUS-00

U.S.

7-99

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

LIGHTING SYSTEMS (Continued)

EXTERIOR LIGHTING (Continued)

STROBE LIGHTS

A high intensity LED strobe light system is installed on the airplane. The

system includes two strobe lights located one on each wing tip. The

lights are used to enhance anti-collision protection for the airplane and

are required anti-collision lights for night operations. The strobe lights

are protected by a pull-off type circuit breaker, labeled STROBE LIGHT,

on the left sidewall switch and circuit breaker panel.

WARNING

Strobe lights should be turned off when taxiing.

Ground operation of the high intensity anti-

collision lights can be considerable annoyance to

ground personnel and other pilots. Do not operate

the anti-collision lights in conditions of fog, clouds,

or haze as the reflection of the light beam can

cause disorientation or vertigo.

FLASHING BEACON LIGHT

A red flashing LED beacon light is installed on the top of the vertical fin

as additional anti-collision protection in flight and for recognition during

ground operation. The light is visible through most angles. The flashing

beacon light is protected by a pull-off type circuit breaker, labeled

BEACON LIGHT, on the left sidewall switch and circuit breaker panel.

WARNING

The flashing beacon should not be used when

flying through clouds or overcast; the flashing light

reflected from water droplets or particles in the

atmosphere, particularly at night, can cause

disorientation or vertigo.

7-100

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

LIGHTING SYSTEMS (Continued)

EXTERIOR LIGHTING (Continued)

WING INSPECTION LIGHT

One LED wing inspection light is installed in the left hand wing root

fairing. The light illuminates the left hand wing and left hand wing strut

for ice detection. The WING LIGHT switch is on the left hand switch

panel and is protected by a pull-type circuit breaker labeled WING ICE

DET LIGHT.

COURTESY LIGHTS

Two LED courtesy lights are installed, one under each wing. The lights

illuminate the area outside of the airplane adjacent to the crew entry

doors. The lights operate in conjunction with the cabin lights and are

controlled by the CABIN light switch as described in the cabin lights

paragraph in this section.

208BPHCUS-00

U.S.

7-101

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

LIGHTING SYSTEMS (Continued)

INTERIOR LIGHTING

Instrument and control panel lighting is provided by integral and flood

lights. Six lighting control knobs are grouped together on the lower part

of the instrument panel to the left of the control pedestal. These

controls vary the intensity of the lighting for the instrument panels,

pedestal, overhead panel, left sidewall panel, LED panels, Garmin

displays, and internally lit standby instruments. The following

paragraphs describe the function of these controls. The circuits for

these lights are protected by two pull-off type circuit breakers, labeled

AVN/LED/STBY LIGHTS and COCKPIT FLOOD LIGHTS, on the left

sidewall switch and circuit breaker panel. Other miscellaneous lighting

provided or available includes control wheel map LED lights, cabin

lights, LED passenger reading lights, and a no smoking/seat belt sign.

Discussion of these lights and their controls is also included in the

following paragraphs. Most interior lighting is LED.

GARMIN DISPLAYS, OPTIONAL ADF, AND HF DISPLAYS (if

installed)

The AVIONICS knob varies the intensity of the Garmin and optional

displays (if installed). Clockwise rotation of the knob increases display

brightness and counterclockwise rotation decreases brightness. The

displays cannot be dimmed to full dark. Rotating this knob

counterclockwise past the dimmest setting will place the displays in

photosensitive mode.

STANDBY INDICATOR CONTROL KNOB

The STANDBY IND knob varies the intensity of the Non-LED integral

lighting of the standby airspeed indicator, attitude indicator, altimeter,

torque indicator and magnetic compass. Clockwise rotation of the knob

increases light brightness and counterclockwise rotation decreases

brightness.

7-102

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

LIGHTING SYSTEMS (Continued)

INTERIOR LIGHTING (Continued)

LED PANELS/ANNUN CONTROL KNOB

The LED PANELS/ANNUN knob varies the intensity of the backlit LED

panels. These panels are inscribed with labels for the switches,

controls, and circuit breakers mounted on the instrument panel.

Clockwise rotation of the knob increases panel brightness and

counterclockwise rotation decreases brightness.

CENTER FLOOD/MAP PANEL KNOB

The CENTER FLOOD knob varies the intensity of the light that

illuminates either pilot’s seat. Clockwise rotation of the knob increases

panel brightness and counterclockwise rotation decreases brightness.

LEFT FLOOD/MAP LIGHTING CONTROL KNOB

The LEFT FLOOD knob varies the brightness of the floodlight located

on the right aft side of the overhead panel. This floodlight may be used

to illuminate the pilot’s map or chart. Clockwise rotation of this control

knob increases lamp brightness while counterclockwise rotation

decreases brightness.

RIGHT FLOOD/MAP LIGHTING CONTROL KNOB

This RIGHT FLOOD knob varies the brightness of the floodlight located

on the left aft side of the overhead panel. This floodlight may be used to

illuminate the co-pilot’s map or chart. Clockwise rotation of this control

knob increases lamp brightness while counterclockwise rotation

decreases brightness.

CONTROL WHEEL MAPLIGHTS

A control wheel maplight is mounted on the bottom of each control

wheel. These lights illuminate the lower portion of the cabin in front of

the pilot and copilot, and are used for checking maps and other flight

data during night operations. Brightness of these lights is adjusted with

a rheostat control knob on the bottom of the control wheel. Rotating the

near side of the knob to the right increases light brightness and to the

left decreases brightness.

208BPHCUS-00

U.S.

7-103

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

LIGHTING SYSTEMS (Continued)

INTERIOR LIGHTING (Continued)

CABIN LIGHTS WITHOUT TIMER (208B Passenger)

The 208B passenger cabin light system without timer consists of four

LED cabin lights installed on the interior of the airplane and courtesy

lights under each wing to facilitate boarding or loading cargo during

night operations. Two lights are located above the center cabin area,

one above the aft cargo door, and one above the aft passenger door.

Controls for the lighting system consists of one 2-way toggle switch

labeled CABIN on the lighting control panel as well as a rocker switch

just forward of both the aft passenger and cargo doors. All three of

these switches will toggle all cabin lights on or off at any time

regardless of the other switch positions.

The circuit for the cabin lights is protected by a pull-off type circuit

breaker, labeled CABIN LTS, on the J-Box panel in the engine bay.

CABIN LIGHTS WITH TIMER (if installed)

The 208B cabin light system with timer consists of four LED cabin lights

installed on the interior of the airplane and courtesy lights under each

wing to facilitate boarding or loading cargo during night operations. Two

lights are located above the center cabin area, one above the aft cargo

door, and one above the aft passenger door.

Controls for the lighting system consists of one 3-way momentary

switch labeled CABIN on the lighting control panel as well as a rocker

switch just forward of both the aft passenger and cargo doors. The

passenger door toggle switch will control all lights except the cargo

door light, while the cargo door toggle switch will control only the cargo

door light. The 3-way momentary switch labeled CABIN on the lighting

control panel will control all lights regardless of the other switch

positions.

The timer circuitry includes a solid state timer that will turn off all lights

after 30 minutes automatically unless they are switched off manually.

The circuit for the cabin lights is protected by a “pull-off” type circuit

breaker, labeled CABIN LTS, on the J-Box panel in the engine bay.

7-104

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

LIGHTING SYSTEMS (Continued)

INTERIOR LIGHTING (Continued)

CABIN LIGHTS WITH TIMER (Super CargoMaster)

The Super CargoMaster cabin light system consists of four LED cabin

lights installed on the interior of the airplane and courtesy lights under

each wing to facilitate boarding or loading cargo during night

operations. Two lights are located above the center cabin area, one

above the aft cargo door and one opposite the aft cargo door.

Controls for the lighting system consists of one 3-way momentary

switch labeled CABIN on the lighting control panel as well as a rocker

switch just forward of the cargo door.

PASSENGER READING LIGHTS (Passenger Version Only)

Passenger reading lights may be installed near each of the aft

passengers positions. The LED lights are located in

14 small

convenience panels above each seat. A pushbutton-type ON, OFF

switch, mounted in each panel, controls the lights. The lights can be

pivoted in their mounting sockets to provide the most comfortable angle

of illumination for the passenger.

NO SMOKE/SEAT BELT SIGN (Passenger Version Only)

A lighted warning sign may be installed in the airplane to facilitate

warning passengers of impending flight operations necessitating the

fastening of seat belts and/or the extinguishing of all smoking materials.

This installation consists of a small lighted panel mounted in the cabin

headliner immediately aft of the overhead console and two toggle-type

switches, labeled SEAT BELT and NO SMOKE, on the lighting control

panel. When these switches are placed in the ON position, the warning

signs illuminate, displaying the international graphic symbolism for

fasten seat belts and no smoking to the rear cabin passengers. The

circuit for the warning sign lights is protected by a pull-off type circuit

breaker, labeled SEAT BELT SIGN, on the left sidewall switch and

circuit breaker panel.

208BPHCUS-00

U.S.

7-105

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM

The temperature and volume of airflow to the cabin is regulated by the

cabin heating, ventilating and defrosting system, refer to Figure 7- 13,

Cabin Heating, Ventilating and Defrosting System. In the heating

system, hot compressor outlet air is routed from the engine through a

flow control valve, then through a mixer/muffler where it is mixed with

cabin return air or warm air from the compressor bleed valve

(depending on the setting of the mixing air valve) to obtain the correct

air temperature before the air is routed to the cabin air distribution

system.

Controls are provided to direct the heated air to the forward and/or aft

portions of the cabin for heating and to the windshield for defrosting.

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 wing struts. The wing inlet ventilating air is routed through

the wing into a plenum chamber located in the center of the cabin top.

The plenum distributes the ventilating air to individual overhead outlets

near each seat position. Two electric blowers are available for the

overhead ventilating system. Refer to Section 9, Supplement 7, for

additional information on ventilating and Air Conditioning systems.

BLEED AIR HEAT SWITCH

A two-position toggle switch, labeled BLEED AIR HEAT, is located on

the CABIN HEAT switch and control panel. The switch 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 (down) closes the valve,

shutting off flow of hot bleed air to the heating system.

7-106

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM (Continued)

TEMPERATURE SELECTOR KNOB

A rotary temperature selector knob, labeled TEMP Control Knob, is

located on the CABIN HEAT switch and control panel. The 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 TEMP Control Knob increases the mass flow

and temperature of the air.

NOTE

• If more cabin heat is needed while on the ground, move

the FUEL CONDITION lever to HIGH IDLE and/or select

the GRD position (pulled out) of the MIXING AIR control.

• Some hysteresis may be encountered when adjusting

bleed air temperature. The resulting amount and

temperature of bleed air may be different when

approaching a particular temperature selector knob

position from a clockwise versus a counterclockwise

direction. Best results can usually be obtained by turning

the temperature selector knob full clockwise and then

slowly turning it counterclockwise to decrease bleed

airflow to the desired amount.

A temperature sensor, located in the outlet duct from the mixer/muffler

operates in conjunction with the TEMP control knob. In the event of a

high temperature condition

(overheat) in the outlet duct, the

temperature sensor will be energized, closing the flow control valve and

thus shutting off the source of hot bleed air from the engine.

208BPHCUS-00

U.S.

7-107

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM (Continued)

MIXING AIR PUSH-PULL CONTROL

A push-pull control, labeled MIXING AIR, GRD-PULL, FLT-PUSH, is

located on the CABIN HEAT switch and control panel. With the push-

pull control in the GRD position (pulled out), warm compressor bleed

valve air is mixed with hot compressor outlet air in the mixer/muffler.

This mode is used during ground operation when warm compressor

bleed valve air is available (at power setting below 92% N_g) and can be

used as additional bleed air heat to augment the hot compressor outlet

bleed air supply during periods of cold ambient temperature. With the

push-pull control in the FLT position (pushed in), cabin return air is

mixed with the hot compressor outlet air in the mixer/muffler. This

recirculation of cabin return air enables the heating system to maintain

the desired temperature for proper cabin heating. If desired, the FLT

position of the push-pull control can be used on the ground when

ambient temperatures are mild and maximum heating is not required.

In this mode, the excess warm compressor bleed valve air available at

power settings below 92% N_g is exhausted overboard from the mixing

air valve.

CAUTION

The MIXING AIR push-pull control should always be in

the FLT position (pushed in) when the airplane is 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 in order to

maintain proper temperature in the cabin heat

distribution system. If the FLT position is not used

during flight, the system may overheat and cause an

automatic shutdown.

7-110

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM (Continued)

AFT/FORWARD CABIN PUSH-PULL CONTROL

A push-pull control, labeled AFT CABIN-PULL, FWD CABIN-PUSH, is

located on the CABIN HEAT switch and control panel. With the control

in the AFT CABIN position (pulled out), heated air is directed to the aft

cabin heater outlets located on the cabin sidewalls at floor level on the

Passenger Version 208 and the outlets in the floor behind the pilot and

copilot on the Cargomaster.

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 the two windshield defroster outlets. The

push-pull control 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, labeled 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

CABIN/FWD CABIN push-pull control also must be pushed in for

availability of forward cabin air for defrosting). With the DEFROST/

FWD CABIN push-pull control in the FWD CABIN position (pushed in),

heated air will be directed to the four heater outlets behind the

instrument panel.

208BPHCUS-00

U.S.

7-111

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM (Continued)

CABIN HEAT FIREWALL SHUTOFF KNOB

A push-pull shutoff knob, labeled CABIN HEAT FIREWALL SHUTOFF,

PULL OFF, is located on the lower right side of the control pedestal.

When pulled out, the knob actuates two firewall shutoff valves, one in

the bleed air supply line to the cabin heating system and one in the

cabin return air line, to the off position. This knob should normally be

pushed in unless a fire is suspected in the engine compartment.

CAUTION

Do not place the CABIN HEAT FIREWALL SHUTOFF

knob in the OFF position when the MIXING AIR control

is in the GRD position because a compressor stall will

occur 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.

VENT AIR CONTROL KNOBS

Two vent air control knobs, labeled VENT AIR, are located on the

overhead console. The knobs control the operation of the shutoff valves

in each wing which control the flow of ventilating air to the cabin. The

knob on the right side of the console controls the right wing shutoff

valve and similarly, the knob on the left side controls the left wing

shutoff valve. When the VENT AIR control knobs are rotated to the

CLOSE position, the wing shutoff valves are closed; rotating the knobs

to the OPEN position progressively opens the wing shutoff valves.

When the optional cabin ventilation fans are installed, rotating the

knobs to the full OPEN position also turns on the ventilation fans.

7-112

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CABIN HEATING, VENTILATING AND DEFROSTING

SYSTEM (Continued)

INSTRUMENT PANEL VENT KNOBS

Two vent knobs, labeled VENT - PULL ON, are located one 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 which pulls in ram air for

distribution through the ventilating outlet.

VENTILATING OUTLETS

Adjustable ventilating outlets (one located above each seat position)

permits individual ventilation to the airplane occupants. The outlets are

the swivel type for optimum positioning, and airflow volume is

controlled by rotating the outlet nozzle controlling an internal valve. In

addition to the pilot and front passenger outlets, the Passenger Version

has 11 outlets in the rear cabin area for use by rear seat passengers.

OXYGEN SYSTEM

Some Cargo Versions are equipped with a two-port oxygen system

having quick-don type masks for the pilot and passenger; other Cargo

Versions can be equipped with a two-port oxygen system utilizing

conventional masks. The Passenger Version can be equipped with up

to 17-port oxygen system utilizing conventional masks. Refer to Section

9, Supplement 6, for complete details and operating instructions.

208BPHCUS-00

U.S.

7-113

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

PITOT-STATIC SYSTEM AND INSTRUMENTS

There are two independent pitot-static systems on the airplane. The left

pitot-static system supplies ram air pressure to Air Data Computer #1

and to the standby airspeed indicator, and supplies static pressure to

Air Data Computer 1 and to the standby airspeed indicator and standby

altimeter. The right pitot-static system provides ram air and static

pressure to Air Data Computer 2. Each system is composed of a

heated pitot-static tube mounted on the leading edge of the

corresponding wing, a drain valve located on the sidewall beneath the

instrument panel, and the associated plumbing necessary to connect

the instruments and sources. In addition, the left system includes a ALT

STATIC AIR source valve located on the lower left corner of the

instrument panel.

The static pressure alternate source valve in the left system 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 going to the static pressure source, the ALT STATIC AIR

source valve should be PULLED ON. Pressures within the cabin will

vary with vents open or closed. Refer to Section 5, Performance for the

effect of varying cabin pressures on airspeed and altimeter readings.

The drain valves incorporated in each system, located on the sidewall

beneath the instrument panel, are used to drain suspected moisture

accumulation by lifting the drain valve lever to the OPEN position as

indicated by the placard adjacent to the valve. The valve must be

returned to the CLOSED position prior to flight.

A left and right pitot-static heat system is installed to assure proper

airspeed indications in the event icing conditions are encountered. The

system is designed to prevent ice formation rather than remove it. The

pitot-static heat system consists of a heating element in each

pitot-static tube, a two-position toggle switch, labeled PITOT/STATIC

HEAT, on the ANTI-ICE switch panel, and two pull-off type circuit

breakers, labeled LEFT PITOT HEAT and RIGHT PITOT HEAT, on the

left sidewall switch and circuit breaker panel. When the PITOT-STATIC

HEAT switch is turned ON, elements in the pitot-static tubes are heated

electrically to maintain proper operation in possible icing conditions.

Both pitot and static systems are monitored by the G1000 system for

insufficient current and alerting is provided to the flight crew by a single

chime and an amber L P/S HEAT, R P/S HEAT, or L-R P/S HEAT

annunciator.

7-114

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

PITOT-STATIC SYSTEM AND INSTRUMENTS

AIRSPEED INDICATORS

The Garmin PFDs are the primary sources of airspeed information.

Standby airspeed information is depicted by a mechanical indicator

calibrated in knots, connected to the left pitot-static system. Limitation

and range markings (in KIAS) match the markings on the PFD as listed

in Section 2, Limitations.

The standby airspeed indicator is a true airspeed indicator and is

equipped with a knob which works in conjunction with the airspeed

indicator dial in a manner similar to the operation of a flight computer.

To operate the indicator, first rotate the knob until pressure altitude is

aligned with outside air temperature in degrees Centigrade. To obtain

pressure altitude, momentarily set the barometric scale on the standby

altimeter to 29.92 and read pressure altitude on the standby altimeter.

Be sure to return the standby altimeter barometric scale to the original

barometric setting after pressure altitude has been obtained. Having

set the knob to correct for altitude and temperature, read the true

airspeed shown in the window by the indicator pointer. For best

accuracy, the indicated airspeed should be corrected to calibrated

airspeed by referring to the Airspeed Calibration chart in Section 5,

Performance. Knowing the calibrated airspeed, read true airspeed in

the window opposite the calibrated airspeed.

VERTICAL SPEED INDICATION

The vertical speed indication on the PFDs depict airplane rate of climb

or descent in feet per minute. The pointers are actuated by atmospheric

pressure changes resulting from changes of altitude as supplied by the

static sources through the respective ADC.

ALTIMETER (STANDBY INSTRUMENT PANEL)

Airplane altitude is depicted by a barometric type altimeter. A knob near

the lower left portion of the indicator provides adjustment of the

instrument's barometric scale to the current altimeter setting.

208BPHCUS-00

U.S.

7-115

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

VACUUM SYSTEM AND INSTRUMENTS

A vacuum system, refer to Figure 7-14, Vacuum System, provides the

suction necessary to operate the standby attitude 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 necessary to operate the indicator. 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 on the aft side of the firewall, and the standby

attitude indicator.

ATTITUDE INDICATOR (Standby Instrument Panel)

Standby attitude information is depicted by a vacuum-driven attitude

indicator. Bank attitude is presented by a pointer at the top of the

indicator relative to the bank scale which has index marks at 10°, 20°,

30°, 60°, and 90° either side of the center mark. Pitch and roll attitudes

are presented by a miniature airplane superimposed over a symbolic

horizon area divided into two sections by a white horizon bar. The

upper blue sky and the lower ground area have arbitrary pitch

reference lines useful for pitch attitude control. A knob at the bottom of

the instrument is provided for inflight adjustment of the miniature

airplane to the horizon bar for a more accurate flight attitude indication.

LOW-VACUUM WARNING FLAG

The standby attitude indicator includes an orange low-vacuum warning

flag (GYRO) that comes into view when the vacuum is below the level

necessary for reliable gyroscope operation.

WARNING

The orange low-vacuum warning flag (gyro) is the

only indication of the loss of the vacuum system.

7-116

U.S.

208BPHCUS-00

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

STALL WARNING SYSTEM

The airplane is equipped with a vane-type stall warning unit, in the

leading edge of the left wing, which is electrically connected to a stall

warning horn located overhead of the pilot's position. The vane in the

wing senses the 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 stall warning system should be checked during the preflight

inspection by momentarily turning on the BATTERY switch and

actuating the vane in the wing. The system is operational if the warning

horn sounds as the vane is pushed upward. The elevator must be off

the forward stop before the stall warning horn is enabled due to the fact

that the airplane is equipped with a stall warning ground disconnect

switch.

A pull-off type circuit breaker, labeled STALL WARN, protects the stall

warning system. Also, it is provided to shut off the warning horn in the

event it should stick in the on position.

WARNING

This circuit breaker must be closed (pushed in) for

approach and landing.

The vane and sensor unit in the wing leading edge is equipped with a

heating element. The heated part of the system is operated by the

STALL HEAT switch on the ANTI-ICE switch panel, and is protected by

the STALL WARN circuit breaker on the left sidewall switch and circuit

breaker panel.

7-118

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

AVIONICS SUPPORT EQUIPMENT

Various avionics support equipment is installed in the airplane, and

includes a microphone/speaker, mic/phone jacks, avionics cooling fans,

12VDC power outlets, an auxiliary audio input jack, and control surface

static dischargers. Description and operation of radio equipment is

covered in the Garmin CRG or Section 9 of this POH/AFM.

AVIONICS COOLING FAN

Two DC electric deck skin fans mounted on the underside of the cowl

deck draw warm air from behind the instrument panel to maintain

proper operating temperatures. In addition, three DC electric fans blow

air directly onto the display heat sinks for prolonged equipment life. The

deck skin fans will operate when the BATTERY switch is ON and the

AVIONICS No. 1 power switch is ON.

MICROPHONE-HEADSET INSTALLATIONS

Radio communications are accomplished by the use of a hand-held

microphone and the airplane speaker, or by aviation-style headsets.

The hand-held microphone stows in a hanger on the front of the

pedestal and plugs into a mic jack located on the right side of the

pedestal. It includes an integral push-to-talk button. The airplane

speakers are located above the pilot's and copilot's positions in the

cabin headliner.

The headsets plug into microphone (MIC) and headset (PHONE) jacks

located on the left side of the instrument panel for the pilot and the right

side of the instrument panel for the copilot. Push-to-talk switches for the

headsets are mounted on the control wheels.

Audio is controlled by the individual audio selector switches and

adjusted for volume level by using the selected receiver volume

controls. The system is designed so that microphones are voice

activated, with transmission over the COM radios controlled by the

push-to-talk switches.

208BPHCUS-00

U.S.

7-119

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

AVIONICS SUPPORT EQUIPMENT (Continued)

STATIC DISCHARGERS

As an aid in IFR flights, wick-type static dischargers are installed to

improve radio communications during flight through dust or various

forms of precipitation (rain, snow or ice crystals). Under precipitation

static

(P-Static) conditions, the build-up and discharge of static

electricity from the trailing edges of the wings, rudder, elevator,

propeller tips, and radio antennas can result in loss of usable radio

signals on all communications and navigation radio equipment. Usually

the ADF is first to be affected and VHF communication equipment is the

last to be affected.

Installation of static dischargers reduces interference from P-Static, but

it is possible to encounter severe P-Static conditions which might cause

the loss of radio signals, even with static dischargers installed.

Whenever possible, avoid known severe precipitation areas to prevent

loss of dependable radio signals. If avoidance is impractical, minimize

airspeed and anticipate temporary loss of radio signals while in these

areas.

Static dischargers lose their effectiveness with age, and therefore,

should be checked periodically (at least at every annual inspection) by

qualified avionics technicians, etc. If testing equipment is not available,

it is recommended that the wicks be replaced every two years,

especially if the airplane is operated frequently in IFR conditions. The

discharger wicks are designed to unscrew from their mounting bases to

facilitate replacement.

12 VDC POWER OUTLET

A power converter, located below the copilot seat, reduces the

airplane's 28 VDC power to 12 VDC. This converter provides up to 10

amps of power to operate portable devices such as notebook computer

and audio players. The power output connector (POWER OUTLET

12V) is located on the center pedestal, refer to Figure 7-2, Instrument

Panel.

7-120

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

AVIONICS SUPPORT EQUIPMENT (Continued)

AUXILIARY AUDIO INPUT JACK

An auxiliary audio input jack (AUX AUDIO IN) is mounted on the lower

aft face of the pedestal, refer to

7-2, Instrument Panel. It allows

connection of entertainment audio devices such as cassette, compact

disc, and MP3 players to play music over the airplane's headsets.

The signal from AUX AUDIO IN is automatically muted during radio

communications or pilot selection of CREW ICS ISOLATION modes

located on the audio panel. The AUX key on the audio panel does not

control the AUX AUDIO IN signal. For a more complete description and

operating instructions of the audio panel, refer to the Garmin G1000

CRG.

NOTE

Since the entertainment audio input is not controlled by a

switch, there is no way to deselect the entertainment

source except to disconnect the source at the audio input

connector.

208BPHCUS-00

U.S.

7-121

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

CABIN FEATURES

CABIN FIRE EXTINGUISHER

A portable fire extinguisher is installed on the cargo barrier in some

Cargo Versions and on the inside of the pilot’s entry door in other Cargo

Versions and the Passenger Version. The extinguisher in both

airplanes is readily accessible in case of fire. The extinguisher should

be checked prior to each flight to ensure that its bottle pressure, as

indicated by the gage on the bottle, is within the green arc and the

operating lever lock pin is securely in place.

To operate the fire extinguisher:

Loosen retaining clamp and remove extinguisher from bracket.

Hold extinguisher upright, pull operating lever lock pin, and press

lever while directing the discharge at the base of the fire at the

near edge. Progress toward the back of the fire by moving the

nozzle rapidly with a side-to-side sweeping motion.

CAUTION

Care must be taken not to direct the initial discharge

directly at the burning surface at close range (less than

five feet) because the high velocity stream may cause

splashing and/or scattering of the burning material.

Anticipate approximately ten seconds of discharge duration.

WARNING

Ventilate the cabin promptly after successfully

extinguishing the fire to reduce the gases produced

by thermal decomposition. Occupants should use

oxygen masks until the smoke clears.

Fire extinguishers should be recharged by a qualified fire extinguisher

agency after each use. After recharging, secure the extinguisher to its

mounting bracket; do not allow it to lie loose on floor or seats.

7-122

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

CABIN FEATURES (Continued)

SUN VISORS

Two sun visors are mounted overhead of the pilot and copilot. The

visors are mounted on adjustable arms which enable them to be swung

and telescoped into the desired windshield area.

CHART AND STORAGE COMPARTMENTS

A map compartment is located in the lower right side of the instrument

panel. A hinged door covers the compartment and can be opened to

gain access into the compartment. Storage pockets are also installed

on the back of the pilot's and copilot's seats and along the bottom edge

of each crew entry door and can be used for stowage of maps and

other small objects.

MISCELLANEOUS EQUIPMENT

ENGINE INLET COVERS AND PROPELLER ANCHOR

Various covers and an anchor are available to close engine openings

and restrain the propeller during inclement weather conditions and

when the airplane is parked for extended periods of time, such as

overnight. The covers preclude the entrance of dust, moisture, bugs,

etc. into the engine and engine compartment.

Two covers are provided which plug into the two front inlets, thereby

closing off these openings. The engine inlet covers may be installed

after the engine has cooled down (ITT indicator showing off scale

temperature). To prevent the propeller from windmilling during windy

conditions, the propeller anchor can be installed over a blade of the

propeller and its anchor strap secured around the nose gear or to the

bracket located on the lower right-hand cowl.

208BPHCUS-00

U.S.

7-123

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

MISCELLANEOUS EQUIPMENT (Continued)

CREW ENTRY STEP ASSEMBLY

The airplane may be equipped with a crew entry step for each crew

entry door. The step assembly attaches to the floorboard just inside the

entry door and extends toward ground level, providing two steps for

entering or exiting the airplane. When not in use, the step assembly

folds and stows just inside the cabin, inboard of each entry door.

CARGO BARRIER AND NETS

A cargo barrier and three cargo barrier nets may be installed directly

behind the pilot's and copilot's seats. The barrier and nets preclude

loose cargo from moving forward into the pilot's and copilot's stations

during an abrupt deceleration. The barrier consists of a U-shaped

assembly of honeycomb composite construction. The assembly

attaches to the four seat rails at the bottom at station 153 and to

structure at the top at approximately station 166. The cargo barrier nets

consist of three nets: one for the left sidewall, one for the right sidewall,

and one for the center. The left and right nets fill in the space between

the barrier assembly and the airplane sidewalls.

The side nets are fastened to the airplane sidewalls and the edge of the

barrier with six anchor-type fasteners each, three on each side. The

center net fills in the opening in the top center of the barrier. The center

net is fastened with four anchor-type fasteners, two on each side.

7-124

U.S.

208BPHCUS-00

CESSNA

SECTION 7

MODEL 208B 867 SHP AIRPLANE AND SYSTEMS DESCRIPTION

GARMIN G1000

MISCELLANEOUS EQUIPMENT (Continued)

CARGO PARTITIONS

Cargo partitions are available and can be installed to divide the cargo

area into convenient compartments. Partitions may be installed in all of

the five locations at stations 188.7, 246.8, 282.0, 307.0, and 332.0. The

cargo partitions are constructed of canvas with nylon webbing

reinforcement straps crisscrossing the partition for added strength. The

ends of each strap have fittings which attach to the floor tracks and

anchor-type fasteners on the sides and top of the fuselage. Four straps

have adjustable buckles for tightening the straps during installation of

the partition.

CARGO DOOR RESTRAINING NET

A restraining net may be installed on the inside of the airplane over the

cargo door opening. The net precludes loose articles from falling out

the cargo door when the doors are opened. The restraining net consists

of two halves which part in the center of the door opening. The front

and rear halves slide fore and aft, respectively, on a rod to open the net.

The net is attached to the sidewall by screws and nutplates along the

front and rear edges of the net. When the net is closed, the two halves

are held together by snap-type fasteners.

CARGO/AIRPLANE TIE-DOWN EQUIPMENT

Various items of tie-down equipment are available for securing cargo

within the airplane and/or tying down the airplane. This equipment

consists of tie-down belt assemblies having various load ratings and

adjustment devices and two types of quick-release tie-down ring

anchors for securing the belts to the cabin seat tracks and anchor

plates. Refer to Section 6 for the recommended use and restrictions of

this equipment.

208BPHCUS-00

U.S.

7-125

SECTION 7

CESSNA

AIRPLANE AND SYSTEMS DESCRIPTION MODEL 208B 867 SHP

GARMIN G1000

MISCELLANEOUS EQUIPMENT (Continued)

HOISTING RINGS

Provisions are made for the installation of four hoisting rings which

attach to the left and right sides of both front and rear spar wing-to-

fuselage attach fittings. Each hoisting ring consists of a hinge which

replaces the washer on the attachment bolt of the fitting. The upper half

of the hinge contains a ring which is used for attaching the hoist when

the airplane is being hoisted. When not in use, the upper hinge half

folds down out of the way. To gain access to the hoisting rings, when

installed, it is necessary to remove the wing-to-fuselage fairing strips.

RELIEF TUBE

Provisions are made for the installation of a relief tube in the aft cabin

area on the Passenger Version. The relief tube is installed on the right

sidewall, just aft of the passenger entry door.

OIL QUICK-DRAIN VALVE

An oil quick-drain valve is available to replace the drain plug on the

bottom of the engine oil tank, and provides quicker, cleaner draining of

the engine oil. To drain the oil with this valve, slip a hose over the end of

the valve, cut the safety wire securing the valve on-off lever in the off

position, and rotate the lever to the on position. After draining, rotate

the valve on-off lever to the off position, remove the hose to check for

leakage, and resafety the on-off lever in the off position.

7-126

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

AIRPLANE HANDLING,

SERVICE AND MAINTENANCE

TABLE OF CONTENTS

Page

Introduction

8-3

Identification Plate

8-3

Cessna Owner Advisories

8-4

United States Airplane Owners

8-4

International Airplane Owners

8-4

Publications

8-5

Airplane File

8-6

Airplane Inspection Periods

8-7

FAA Required Inspections

8-7

Inspection Programs

8-7

Cessna Customer Care Program

8-8

CESCOM System

8-8

Pilot Conducted Preventive Maintenance

8-9

Alterations or Repairs

8-9

Ground Handling

8-10

Towing

8-10

Parking

8-11

Tiedown

8-12

Jacking

8-13

Leveling

8-15

Servicing

8-16

Oil

8-17

Oil Specification

8-17

Oil System Servicing

8-18

Oil Capacity

8-19

Oil Quantity Operating Range

8-19

(Continued Next Page)

208BPHCUS-00

U.S.

8-1

SECTION 8

CESSNA

AIRPLANE HANDLING, SERVICE

MODEL 208B 867 SHP

AND MAINTENANCE

GARMIN G1000

TABLE OF CONTENTS (Continued)

Page

Fuel

8-20

Fuel Grade (Specification) and Fuel Additives

8-20

Fuel Additives

8-21

Fuel Capacity

8-26

Fuel Contamination

8-27

Landing Gear

8-29

Oxygen

8-30

Ground Deice/Anti-Ice Operations

8-31

Holdover Timetable

(Type I, Type II, Type III, and Type IV Fluids)

8-35

Essential Areas to be Deiced

8-42

Essential Areas to Apply Anti-Ice Fluid

8-43

Deice and Anti-Ice Fluid

Direct Spray Avoidance Areas

8-44

Deicing and Anti-Icing Application

8-45

Cleaning and Care

8-46

Painted Surfaces

8-46

Windshield and Windows

8-47

Stabilizer Abrasion Boot Care

8-51

Propeller Care

8-51

Engine

8-52

Interior Care

8-55

Avionics Care

8-56

Prolonged Out of Service Care

8-57

8-2

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

INTRODUCTION

This section contains factory recommended procedures for proper

ground handling and routine care and servicing of your airplane. It also

identifies certain inspection and maintenance requirements which must

be followed if your airplane is to retain that new airplane performance

and dependability. It is important to follow a planned schedule of

lubrication and preventive maintenance based on climatic and flying

conditions encountered in your local area.

Keep in touch with your local Cessna Authorized Service Facility and

take advantage of their knowledge and experience. Your Cessna

Authorized Service Facility knows your airplane and how to maintain it,

and will remind you when lubrications and oil changes are necessary,

as well as other seasonal and periodic services.

The airplane should be regularly inspected and maintained in

accordance with information found in the airplane maintenance manual

and in any company issued service bulletins and service letters. All

service bulletins pertaining to the airplane by serial number should be

accomplished and the airplane should receive repetitive and required

inspections. Cessna does not condone modifications, whether by

Supplemental Type Certificate

(STC) or otherwise, unless these

certificates are held and/or approved by Cessna. Other modifications

may void warranties on the airplane since Cessna has no way of

knowing the full effect on the overall airplane. Operation of an airplane

that has been modified may be a risk to the occupants, and operating

procedures and performance data set forth in the POH may no longer

be considered accurate for the modified airplane.

IDENTIFICATION PLATE

All correspondence regarding your airplane should include the Serial

Number. The Serial Number, Model Number, Production Certificate

Number (PC) and Type Certificate Number (TC) can be found on the

Identification Plate, located on the aft left tailcone. The Finish and Trim

Plate, which is installed on the lower part of the left forward doorpost,

contains a code describing the exterior paint combination of the

airplane. The code may be used in conjunction with an applicable

Illustrated Parts Catalog if finish and trim information is needed.

208BPHCUS-00

U.S.

8-3

SECTION 8

CESSNA

AIRPLANE HANDLING, SERVICE

MODEL 208B 867 SHP

AND MAINTENANCE

GARMIN G1000

CESSNA OWNER ADVISORIES

Cessna Owner Advisories are sent to Cessna Aircraft FAA Registered

owners of record at no charge to inform them about mandatory and/or

beneficial airplane service requirements and product changes. Copies

of the actual bulletins are available from Cessna Authorized Service

Facilities and Cessna Customer Care.

As a convenience, service documents are now available online to all

our customers through a simple, free-of-charge registration process. If

you would like to sign up, please visit the “Customer Access” link at

www.cessnasupport.com to register.

UNITED STATES AIRPLANE OWNERS

If your airplane is registered in the U.S., appropriate Cessna Owner

Advisories will be mailed to you automatically according to the latest

airplane registration name and address which you have provided to the

FAA. Therefore, it is important that you provide correct and up to date

mailing information to the FAA.

If you require a duplicate Owner Advisory to be sent to an address

different from the FAA aircraft registration address, please complete

and return an Owner Advisory Application (otherwise no action is

required on your part).

INTERNATIONAL AIRPLANE OWNERS

To receive Cessna Owner Advisories, please complete and return an

Owner Advisory Application.

Receipt of a valid Owner Advisory Application will establish your

Cessna Owner Advisory service for one year, after which you will be

sent a renewal notice. It is important that you respond promptly to

update your address for this critical service.

8-4

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

PUBLICATIONS

Various publications and flight operation aids are furnished in the

airplane when delivered from the factory. These items are listed below.

• Customer Care Program Handbook

• Pilot’s Operating Handbook and FAA Approved Airplane

Flight Manual

• Pilot’s Checklist

• Passenger Briefing Card

• Cessna Authorized Service Facility Directory

To obtain additional publications or owner advisory information, you

may contact Cessna Customer Care at (316) 517-5800. Fax (316) 517-

7271 or write to Cessna Aircraft Company, P.O. Box 7706, Attn. Dept.

569, Wichita, KS 67277.

The following additional publications, plus many other supplies that are

applicable to your airplane, are available from a Cessna Authorized

Service Facility.

• Information Manual

(contains Pilot’s Operating

Handbook Information)

• Maintenance Manual, Wiring Diagram Manual and

Illustrated Parts Catalog

Cessna Authorized Service Facilities have access to a Customer Care

Supplies and Publications Catalog covering all available items, many of

which the Authorized Service Facility keeps on hand. The Authorized

Service Facility can place an order for any item which is not in stock.

NOTE

A Pilot's Operating Handbook and FAA Approved Airplane

Flight Manual which is lost or destroyed may be replaced by

contacting Cessna Customer Care.

208BPHCUS-00

U.S.

8-5

SECTION 8

CESSNA

AIRPLANE HANDLING, SERVICE

MODEL 208B 867 SHP

AND MAINTENANCE

GARMIN G1000

AIRPLANE FILE

There are miscellaneous data, information and licenses that are a part

of the airplane file. The following is a checklist for that file. In addition, a

periodic check should be made of the latest Federal Aviation

Regulations to ensure that all data requirements are met.

To be displayed in the airplane at all times:

1. Aircraft Airworthiness Certificate (FAA Form 8100-2).

2. Aircraft Registration Certificate (FAA Form 8050-3).

3. Aircraft Radio Station License, (if applicable).

To be carried in the airplane at all times:

1. Current Pilot's Operating Handbook and FAA Approved Airplane

Flight Manual.

2. Garmin G1000 Cockpit Reference Guide (190-00384-00 Rev. B

or subsequent).

3. Weight and Balance, and associated papers (latest copy of the

Repair and Alteration Form, FAA Form 337, if applicable).

4. Equipment List.

To be made available upon request:

1. Airplane Logbook.

2. Engine Logbook.

3. Propeller Logbook.

Most of the items listed are required by the United States Federal

Aviation Regulations. Since the Regulations of other nations may

require other documents and data, owners of airplanes not registered in

the United States should check with their own aviation officials to

determine their individual requirements.

Cessna recommends that these items, plus the Pilot's Checklists,

CESCOM/Customer Care Program Handbook and Customer Care

Card, be carried in the airplane at all times.

8-6

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

AIRPLANE INSPECTION PERIODS

FAA REQUIRED INSPECTIONS

As required by U.S. Federal Aviation Regulations, all civil aircraft of

U.S. registry must undergo a complete inspection (annual) each twelve

calendar months. In addition to the required annual inspection, aircraft

operated commercially (for hire) must have a complete inspection

every 100 hours of operation.

The FAA may require other inspections by the issuance of

Airworthiness Directives

(ADs) applicable to the airplane, engine,

propeller and components. It is the responsibility of the owner/operator

to ensure compliance with all applicable airworthiness directives, and

when the inspections are repetitive, to take appropriate steps to prevent

inadvertent noncompliance.

If an airplane is being operated under a CFR Part 135 Certificate, the

operator can choose to use an Approved Aircraft Inspection Program.

INSPECTION PROGRAMS

Refer to the 208 Maintenance Manual, Chapter 4-00-00, Airworthiness

Limitations, for FAA approved mandatory replacement times and

inspection intervals for components and structures that are life-limited.

The section also gives the scheduled inspection requirements for

structural and fatigue components that are considered a part of the

certification process. Refer to Chapter 5-00-00 for approved time limits

and maintenance checks for the Model 208B airplanes.

Regardless of the inspection method selected, the owner should keep

in mind that 14 CFR 43 and 14 CFR 91 establishes the requirement

that properly certified agencies or personnel accomplish all required

FAA inspections and most of the manufacturer recommended

inspections.

NOTE

Airplanes operating in other than U.S. registry should refer

to the regulations of the country of certification for

information on approved maintenance inspection programs.

208BPHCUS-00

U.S.

8-7

SECTION 8

CESSNA

AIRPLANE HANDLING, SERVICE

MODEL 208B 867 SHP

AND MAINTENANCE

GARMIN G1000

AIRPLANE INSPECTION PERIODS (Continued)

CESSNA CUSTOMER CARE PROGRAM

Specific benefits and provisions of the Cessna Warranty plus other

important benefits are contained in the CESCOM/Customer Care

Program Handbook supplied with the airplane. The CESCOM/

Customer Care Program Handbook should be thoroughly reviewed and

kept in the airplane at all times.

Contact a Cessna Authorized Service Facility for the first 100-hour or

annual inspection depending on the program chosen for the airplane.

These inspections can be performed by any Cessna Authorized

Service Facility.

CESCOM SYSTEM

CESCOM is Cessna’s Computerized Maintenance Records System.

This comprehensive system provides an accurate and simple method

of monitoring and scheduling inspections, Service Bulletins, Service

Kits, Airworthiness Directives as well as scheduled and unscheduled

maintenance activities. For detail information about CESCOM, refer to

the CESCOM Instruction Manual supplied with the airplane.

8-8

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

PILOT CONDUCTED PREVENTIVE MAINTENANCE

A certified pilot who owns or operates an airplane not used as an air

carrier is authorized by 14 CFR 43 to perform limited maintenance on

his airplane. Refer to 14 CFR 43 for a list of the specific maintenance

operations which are allowed.

NOTE

Pilots operating airplanes of other than U.S. registry should

refer to the regulations of the country of certification for

information on preventive maintenance that may be

performed by pilots.

A Maintenance Manual must be obtained prior to performing any

preventive maintenance to ensure that proper procedures are followed.

A Cessna Authorized Service Facility should be contacted for further

information or for required maintenance which must be accomplished

by appropriately licensed personnel.

ALTERATIONS OR REPAIRS

It is essential that the FAA be contacted prior to any alterations on the

airplane to ensure that airworthiness of the airplane is not violated.

Alterations or repairs to the airplane must be accomplished by licensed

personnel, utilizing only FAA Approved components and FAA Approved

data, such as Cessna Service Bulletins.

208BPHCUS-00

U.S.

8-9

SECTION 8

CESSNA

AIRPLANE HANDLING, SERVICE

MODEL 208B 867 SHP

AND MAINTENANCE

GARMIN G1000

GROUND HANDLING

TOWING

The airplane is most easily and safely maneuvered by hand with the

towbar attached to the nosewheel. The towbar may be stowed in Zone

6. Moving the airplane by hand will require that the individual steering

with the towbar be assisted by personnel pushing at the wing struts.

CAUTION

Do not push or pull the airplane using the propeller

blades or control surfaces.

Use extreme caution during towing operations, especially when towing

with a vehicle. Do not exceed the nose gear turning angle limit of 51.5°

either side of center as shown by the steering limit marks.

If excess force is exerted beyond the turning limit, a red over-travel

indicator block (frangible stop) will fracture and the block, attached to a

cable, will fall into view alongside the nose strut. This should be

checked routinely during preflight inspection to prevent operation with a

damaged nose gear.

CAUTION

UNLOCK the rudder lock and remove any external

rudder locks before towing.

If the airplane is towed or pushed over a rough surface during

hangaring, watch that the normal cushioning action of the nose gear

does not cause excessive vertical movement of the tail and the

resulting contact with low hangar doors or structure. A flat nose tire will

also increase tail height.

8-10

U.S.

208BPHCUS-00

CESSNA

SECTION 8

MODEL 208B 867 SHP

AIRPLANE HANDLING, SERVICE

GARMIN G1000

AND MAINTENANCE

GROUND HANDLING (Continued)

PARKING

When parking the airplane, head into the wind and set the parking

brakes. Do not set the parking brakes during cold weather when

accumulated moisture may freeze the brakes, or when the brakes are

overheated. Install the control wheel lock, engage the rudder lock, and

chock the wheels (if the brakes are not utilized) to prevent airplane

movement. In severe weather and high wind conditions, tie the airplane

down as outlined in the tiedown section.

CAUTION

Any time the airplane is loaded heavily, the footprint

pressure (pressure of the airplane wheels upon the

contact surface of the parking area or runway) will be

extremely high, and surfaces such as hot asphalt or

sod may not adequately support the weight of the

airplane. Precautions should be taken to avoid airplane

parking or movement on such surfaces.

208BPHCUS-00

U.S.

8-11

Cessna 208B Grand Caravan EX. Airplane Flight Manual (2012) - page 6