BEECHCRAFT 1900D

PILOT TRAINING MANUAL

This training manual is for training purposes only.

The applicable FAA Aircraft Flight Manual contains more detailed

information and procedures which must be followed.
Contents
AIRCRAFT GENERAL............................................................................................................................................................... 10
Introduction ...................................................................................................................................................................... 10
General Description .......................................................................................................................................................... 10
Fuselage ............................................................................................................................................................................ 10
Aircraft Specifications & Dimensions ................................................................................................................................ 10
General Specifications................................................................................................................................................... 10
Exterior Dimensions ...................................................................................................................................................... 11
Cabin and Entry Dimensions ......................................................................................................................................... 11
Maximum Baggage Compartment Weights .................................................................................................................. 11
Aircraft Volumes ........................................................................................................................................................... 11
Minimum ground turning radius................................................................................................................................... 12
Aircraft Doors & Emergency Exits ..................................................................................................................................... 12
Airstair Door .................................................................................................................................................................. 12
Cargo Door .................................................................................................................................................................... 14
Cabin Emergency Exits .................................................................................................................................................. 15
Wing / Nacelle ................................................................................................................................................................... 16
Empennage ....................................................................................................................................................................... 16
Powerplants ...................................................................................................................................................................... 16
Electrical System ............................................................................................................................................................... 16
Hydraulic System............................................................................................................................................................... 17
Hydraulic System Capacity & Type................................................................................................................................ 17
Fuel and Oil System ........................................................................................................................................................... 17
Fuel Capacity & Type ..................................................................................................................................................... 17
Oil Capacity ................................................................................................................................................................... 17
Environmental Systems..................................................................................................................................................... 17
Flight Controls ................................................................................................................................................................... 17
Limitations......................................................................................................................................................................... 18
Airframe ........................................................................................................................................................................ 18
Engine............................................................................................................................................................................ 19
Instruments & Markings ................................................................................................................................................... 21
Radios & Related Equipment ............................................................................................................................................ 25
Lighting .............................................................................................................................................................................. 26
General Description .......................................................................................................................................................... 26
Interior Lighting................................................................................................................................................................. 26
Cockpit .......................................................................................................................................................................... 26
Cabin ............................................................................................................................................................................. 28
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 Entranceway lighting......................................................................................................................................................... 29
Airstairs ......................................................................................................................................................................... 29
Cargo Bay ...................................................................................................................................................................... 29
Exterior Lighting ................................................................................................................................................................ 29
ANNUNCIATOR SYSTEM ........................................................................................................................................................ 31
General Description .......................................................................................................................................................... 31
Master Flashers ................................................................................................................................................................. 31
Annunciators ................................................................................................................................................................. 32
Fluid Level Sensor.............................................................................................................................................................. 38
Annunciator Dimming Feature.......................................................................................................................................... 39
FIRE PROTECTION.................................................................................................................................................................. 40
General .............................................................................................................................................................................. 40
Engine Fire Protection....................................................................................................................................................... 40
Detection System Operation ......................................................................................................................................... 40
Interior Fire Protection (Cockpit/Cabin/Cargo/Lavatory) ................................................................................................. 41
Detection System Operation ......................................................................................................................................... 41
Extinguishing System Operation ................................................................................................................................... 41
Bleed Air System Fire Protection ...................................................................................................................................... 42
Detection System Operation ......................................................................................................................................... 42
Fire Prevention System Operation ................................................................................................................................ 42
System Checks................................................................................................................................................................... 43
Engine Fire Detection Test ............................................................................................................................................ 43
Engine Fire Extinguishing System Test .......................................................................................................................... 43
Bleed Air Detection System Test ................................................................................................................................... 44
Lavatory Smoke Detector System Test ......................................................................................................................... 44
ELECTRICAL SYSTEM .............................................................................................................................................................. 45
General Description .......................................................................................................................................................... 45
Electrical Terminology................................................................................................................................................... 45
Electrical Symbols ......................................................................................................................................................... 45
DC Power Distribution....................................................................................................................................................... 46
Battery............................................................................................................................................................................... 47
Thermal Runaway ......................................................................................................................................................... 48
External Power .................................................................................................................................................................. 48
Starter-Generators ............................................................................................................................................................ 49
Starter Operations ........................................................................................................................................................ 49
Starter Limitations......................................................................................................................................................... 50
Generator Operation .................................................................................................................................................... 50
Generator Control Unit (GCU)....................................................................................................................................... 51

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 Bus Tie System .................................................................................................................................................................. 52
Circuit Breakers & Fuses ................................................................................................................................................... 54
Breaker Panels .............................................................................................................................................................. 54
AC Power Distribution....................................................................................................................................................... 55
DC Bus System – Main Items ............................................................................................................................................ 58
Electric Malfunctions and Faults ....................................................................................................................................... 59
“L DC GEN” or “R DC GEN” ............................................................................................................................................ 59
“L GEN TIE OPEN” or “R GEN TIE OPEN” ....................................................................................................................... 59
“L GEN TIE OPEN” with “L DC GEN” or “R GEN TIE OPEN” with “R DC GEN” ................................................................ 59
“L GEN TIE OPEN” and “BATT TIE OPEN” and “R GEN TIE OPEN” ................................................................................. 59
“L DC GEN” and “L GEN TIE OPEN” and “R GEN TIE OPEN” and “R DC GEN” and “R ENIVRO FAIL” ............................. 59
“L PITOT HEAT” ............................................................................................................................................................. 59
“L AC BUS” or “R AC BUS” ............................................................................................................................................. 60
“L AC BUS” and “R AC BUS” .......................................................................................................................................... 60
Loss of all engine instruments, Comm #1, NAV #1 and RMI #1 without any annunciators ......................................... 60
System Checks................................................................................................................................................................... 60
Electrical System Test.................................................................................................................................................... 60
FUEL SYSTEM......................................................................................................................................................................... 62
General Description .......................................................................................................................................................... 62
Fuel Tank System .............................................................................................................................................................. 62
Fuel Storage .................................................................................................................................................................. 62
Fuel Tank Venting.......................................................................................................................................................... 62
Fuel Drains .................................................................................................................................................................... 63
Fuel System Operation ...................................................................................................................................................... 63
Main Tank Operation .................................................................................................................................................... 63
Auxiliary Tank Operation............................................................................................................................................... 63
Collector Tank Operation .............................................................................................................................................. 64
Fuel Heater .................................................................................................................................................................... 64
Fuel Pumps .................................................................................................................................................................... 65
Fuel Cross-Transfer Operation ...................................................................................................................................... 66
Firewall Fuel Valve ........................................................................................................................................................ 66
Fuel Indication Systems .................................................................................................................................................... 67
Fuel Quantity Gauges .................................................................................................................................................... 67
Low Fuel Quantity Warning System .............................................................................................................................. 68
Liquid Level Indicators .................................................................................................................................................. 68
Fuel Flow ....................................................................................................................................................................... 68
Fuel Purging System .......................................................................................................................................................... 68
Limitations......................................................................................................................................................................... 69

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 Fuel System Malfunctions and Faults ............................................................................................................................... 69
“L FUEL QTY” or “R FUEL QTY” ...................................................................................................................................... 69
“L COL TANK LOW” or “R COL TANK LOW” ................................................................................................................... 69
“XFR VALVE FAIL” .......................................................................................................................................................... 69
“L FUEL PRESS LO” or “R FUEL PRESS LO” ..................................................................................................................... 70
“L NO AUX XFR” or “R NO AUX XFR” ............................................................................................................................. 70
“L FW FUEL VALVE” or “R FW FUEL VALVE” .................................................................................................................. 70
Engine Driven High Pressure Fuel Pump Failure ........................................................................................................... 70
System Checks................................................................................................................................................................... 71
Fuel System Test ........................................................................................................................................................... 71
POWERPLANT........................................................................................................................................................................ 72
General Description .......................................................................................................................................................... 72
Engine Description & Operation ....................................................................................................................................... 72
Engine Stations.............................................................................................................................................................. 72
Engine Airflow ............................................................................................................................................................... 73
Gas Generator Section .................................................................................................................................................. 73
Accessory Section.......................................................................................................................................................... 74
Power Section ............................................................................................................................................................... 74
Lubrication System............................................................................................................................................................ 74
Oil Pressure System....................................................................................................................................................... 74
Oil Scavenge System ..................................................................................................................................................... 75
Fuel System ....................................................................................................................................................................... 76
Fuel Control Unit ........................................................................................................................................................... 76
Flow Divider and Purge Valve ....................................................................................................................................... 78
Fuel Manifold and Duplex Nozzles ................................................................................................................................ 79
Ignition System.................................................................................................................................................................. 79
Starting Mode ............................................................................................................................................................... 79
Auto Ignition Mode ....................................................................................................................................................... 79
Engine Instruments ........................................................................................................................................................... 80
Engine Controls ................................................................................................................................................................. 82
Power Lever .................................................................................................................................................................. 82
Condition Lever ............................................................................................................................................................. 83
Staring Requirements & Procedures................................................................................................................................. 83
Starting Requirements .................................................................................................................................................. 83
Limitations......................................................................................................................................................................... 85
Lubrication System Malfunctions & Faults ....................................................................................................................... 87
“L OIL PRESS LOW" or “R OIL PRESS LOW” ................................................................................................................... 87
Oil Pressure below 90 PSI.............................................................................................................................................. 87

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 “L FUEL PRESS LO” or “R FUEL PRESS LO” ..................................................................................................................... 87
PROPELLER ............................................................................................................................................................................ 89
General Description .......................................................................................................................................................... 89
Propeller Blade Angles ...................................................................................................................................................... 89
Normal Operation Angles ............................................................................................................................................. 89
Propeller Governing .......................................................................................................................................................... 89
Primary Governor.......................................................................................................................................................... 89
Primary Governor Operation ........................................................................................................................................ 90
Overspeed Governor ..................................................................................................................................................... 94
Fuel Topping governor .................................................................................................................................................. 95
Autofeather System .......................................................................................................................................................... 95
Engine Failure Autofeather Sequence .......................................................................................................................... 95
Propeller Synchrophaser................................................................................................................................................... 96
Limitations......................................................................................................................................................................... 96
Propeller Malfunctions ..................................................................................................................................................... 97
Propeller System Checks ................................................................................................................................................... 97
Propeller Autofeather System Test ............................................................................................................................... 97
Propeller Overspeed Test.............................................................................................................................................. 98
Propeller Low Pitch Stop Test ....................................................................................................................................... 99
PRESSURIZATION................................................................................................................................................................. 100
General Description ........................................................................................................................................................ 100
System Operation ........................................................................................................................................................... 100
Bleed Air ...................................................................................................................................................................... 100
Outflow Valves ............................................................................................................................................................ 101
Pressurization Control ................................................................................................................................................. 101
Pressurization Gauges ..................................................................................................................................................... 102
Cabin Differential Warning ......................................................................................................................................... 103
Cabin Altitude Warning ............................................................................................................................................... 103
Malfunctions & Faults ..................................................................................................................................................... 104
“L ENVIRO FAIL” and/or “R ENVIRO FAIL” ................................................................................................................... 104
“CABIN ALT HI” ............................................................................................................................................................ 104
“CABIN DIFF HI”........................................................................................................................................................... 104
Pressurization System Check .......................................................................................................................................... 105
Pressurization/Environmental Ducting ........................................................................................................................... 106
ENVIRONMENTAL SYSTEM.................................................................................................................................................. 107
General Description ........................................................................................................................................................ 107
Components .................................................................................................................................................................... 107
Bleed Air ...................................................................................................................................................................... 107

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 Air Cycle Machine (ACM) ............................................................................................................................................ 107
Vapour Cycle Machine (VCM) ..................................................................................................................................... 108
Environmental Controls .............................................................................................................................................. 109
Temperature Gauges .................................................................................................................................................. 110
Air Distribution ................................................................................................................................................................ 111
Air Distribution Controls ............................................................................................................................................. 111
System Operation ........................................................................................................................................................... 112
Heating ........................................................................................................................................................................ 112
Cooling (ACM Only) ..................................................................................................................................................... 112
Cooling (ACM and VCM).............................................................................................................................................. 112
Unpressurized Ventilation .............................................................................................................................................. 112
Failures & Malfunctions .................................................................................................................................................. 113
“L ENVIRO FAIL” and/or “R ENVIRO FAIL” ................................................................................................................... 113
Environmental System check .......................................................................................................................................... 114
T-Test (Over-Temperature Test) ................................................................................................................................. 114
Maintenance - P-Test (Over-Pressure Test) ................................................................................................................ 114
OXYGEN SYSTEM ................................................................................................................................................................. 115
General ............................................................................................................................................................................ 115
System Components ....................................................................................................................................................... 115
Oxygen Cylinders and Pressure Regulators ................................................................................................................ 115
Crew Oxygen System................................................................................................................................................... 116
Passenger Oxygen System .......................................................................................................................................... 117
Determining Available Oxygen........................................................................................................................................ 119
Effective Performance Time (EPT) or Time of Useful Conscientious (TUC) .................................................................... 119
Oxygen Use - Standard Operating Procedures (SOP’s) ................................................................................................... 120
Smoke or Fume Elimination ........................................................................................................................................ 120
Cabin Decompression ................................................................................................................................................. 120
System Checks................................................................................................................................................................. 120
Oxygen Duration Charts .................................................................................................................................................. 121
SCOTT-EROS & PURITAN BENNETT-SWEEP ON 2000 – Normal Mode ....................................................................... 121
PURITAN BENNETT-SWEEP ON 2000 (Inflatable Harness) - 100% Mode ................................................................... 122
SCOTT-EROS (Inflatable Harness) - 100% Mode ......................................................................................................... 123
Determining Available Oxygen Worksheet ..................................................................................................................... 124
PNEUMATICS ....................................................................................................................................................................... 125
General Description ........................................................................................................................................................ 125
Pneumatic Air Source ...................................................................................................................................................... 125
Pneumatic Air Systems.................................................................................................................................................... 126
Pressure Systems ........................................................................................................................................................ 126

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 Malfunctions & Faults ..................................................................................................................................................... 127
“L BL AIR FAIL” and “R BL AIR FAIL” and “L BK DI OVHT” and “R BK DI OVHT” ........................................................... 127
System Checks................................................................................................................................................................. 128
Pneumatic System....................................................................................................................................................... 128
LANDING GEAR & BRAKES................................................................................................................................................... 129
General Description ........................................................................................................................................................ 129
Landing Gear ................................................................................................................................................................... 129
Nose Landing Gear ...................................................................................................................................................... 129
Main Landing Gear ...................................................................................................................................................... 130
Landing Gear Controls................................................................................................................................................. 132
Hydraulic Power Pack.................................................................................................................................................. 133
Alternate Landing Gear Extension Handle .................................................................................................................. 134
Landing Gear Position/Warning System ..................................................................................................................... 134
Landing Gear Operation .............................................................................................................................................. 134
Landing Gear Troubleshooting .................................................................................................................................... 137
Landing Gear Switch Items.............................................................................................................................................. 138
Nose Landing Gear ...................................................................................................................................................... 138
Left Main Landing Gear ............................................................................................................................................... 138
Right Main Landing Gear............................................................................................................................................. 138
System Checks................................................................................................................................................................. 139
Landing Gear Control and Indication .......................................................................................................................... 139
Limitations....................................................................................................................................................................... 139
FLIGHT CONTROLS............................................................................................................................................................... 140
General Description ........................................................................................................................................................ 140
Primary Controls ............................................................................................................................................................. 140
Ailerons ....................................................................................................................................................................... 140
Elevator ....................................................................................................................................................................... 141
Rudder ......................................................................................................................................................................... 142
Flaps ............................................................................................................................................................................ 143
Stall Warning System ...................................................................................................................................................... 144
Limitations....................................................................................................................................................................... 144
System Checks................................................................................................................................................................. 144
Electric Elevator Trim Check ....................................................................................................................................... 144
Flap System Check....................................................................................................................................................... 145
Flight Controls Check .................................................................................................................................................. 145
Stall Warning System Test........................................................................................................................................... 145
Yaw Damp Check (includes Rudder Boost warning system) ....................................................................................... 145
Rudder Boost Check .................................................................................................................................................... 146

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ICE AND RAIN PROTECTION ................................................................................................................................................ 147
General Description ........................................................................................................................................................ 147
Rain Protection ............................................................................................................................................................... 147
Ice Protection .................................................................................................................................................................. 147
Anti-Ice Systems .......................................................................................................................................................... 147
Deice Systems ............................................................................................................................................................. 151
System Checks................................................................................................................................................................. 154
Windshield Anti-ice ..................................................................................................................................................... 154
Prop Deice ................................................................................................................................................................... 154
Surface Deice .............................................................................................................................................................. 155
Misc Deice ................................................................................................................................................................... 155
Brake Deice ................................................................................................................................................................. 156
Limitations....................................................................................................................................................................... 156
AVIONICS ............................................................................................................................................................................. 157
General Description ........................................................................................................................................................ 157
Avionics Master Switch ................................................................................................................................................... 157
Avionics Master Switch Operation .............................................................................................................................. 157
Primary Flight Instruments ............................................................................................................................................. 158
Airspeed Indicator ....................................................................................................................................................... 158
Altimeter ..................................................................................................................................................................... 158
Instantaneous Vertical Speed Indicator ...................................................................................................................... 158
EFIS-84 System ............................................................................................................................................................ 158
Backup Flight Instruments .............................................................................................................................................. 168
Standby Attitude indicator (AI-804AF) ........................................................................................................................ 168
Standby Magnetic Compass ........................................................................................................................................ 169
Environmental Inputs...................................................................................................................................................... 169
Pitot-Static System ...................................................................................................................................................... 169
Outside Air Temperature ............................................................................................................................................ 170
Communications ............................................................................................................................................................. 171
VHF Communication Receiver Control (CTL-22) ......................................................................................................... 171
Audio Control System ................................................................................................................................................. 171
Navigation ....................................................................................................................................................................... 172
Navigation Receiver Control (VIR-32) ......................................................................................................................... 172
Automatic Direction Finder Control (CTL-62) ............................................................................................................. 173
Radio Magnetic Indicator (RMI-36)............................................................................................................................. 173
Transponder (CTL-92).................................................................................................................................................. 174
Global Positioning System (KLN 90B) .......................................................................................................................... 175
Weather Avoidance ........................................................................................................................................................ 177

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 Weather Radar Control Panel ..................................................................................................................................... 177
Misc Systems ................................................................................................................................................................... 178
ST3400 Multifunction Display ..................................................................................................................................... 178
Traffic Collision and Avoidance System (TCAS II) ........................................................................................................ 180
Static Discharge Wicks ................................................................................................................................................ 181
System Tests.................................................................................................................................................................... 181
Overspeed Test ........................................................................................................................................................... 181
Standby Attitude Indicator.......................................................................................................................................... 181
EFIS Aux Power............................................................................................................................................................ 182
Ground Proximity Warning System............................................................................................................................. 182
Altitude Alerter ........................................................................................................................................................... 182
EFIS Test ...................................................................................................................................................................... 183
TAWS / TCAS ............................................................................................................................................................... 184
Transponder ................................................................................................................................................................ 184
EFIS reversionary ......................................................................................................................................................... 185
RADAR ......................................................................................................................................................................... 186
Briefer ......................................................................................................................................................................... 186
Cockpit Voice Recorder ............................................................................................................................................... 186
Flight Director ............................................................................................................................................................. 187

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AIRCRAFT GENERAL

Introduction
Beechcraft has developed 4 series of the 1900 Airliner (UA, UB, UC and UE) which were available in several
configurations and although there are many similarities between the different series, this manual has been developed
specifically for the Beechcraft 1900D Airliner (UE) and its available configurations.

This manual has been separated into 19 chapters that are designed to be independent of each other and referenced in
any order.

General Description
The Beechcraft 1900D is a high performance, pressurized, turboprop aircraft that has been certified for Day and Night
IFR as well as flight into known icing conditions. The 1900D was designed as a larger and improved version of the
Beechcraft King air 200 and retains many similarities.

Fuselage
The Beechcraft 1900D Airliner is a 19 seat (18 with optional Lavatory installed) commuter style aircraft. The fuselage is a
conventional semi-monocoque construction which is encased by a skin of high strength aluminum alloy. The fuselage
integrates a fold down airstair door at the front of the passenger cabin, ahead of the left wing and fold up cargo door at
the rear of the aircraft behind the passenger cabin on the left side of the aircraft. Also, there are 3 emergency exits; two
on the right side and one on the left side over the wings near the center of the passenger cabin.

Without Lavatory

With Lavatory

Aircraft Specifications & Dimensions

General Specifications
Aircraft Designation BE02
Minimum Crew One
Maximum Passengers 19 (18 with lavatory installed)
Engine Pratt & Whitney PT6A-67D
Propellers Hartzell HC-E4A-3J
Wing Area 310 ft2
Wing Loading 54.7 lbs per ft2
Power Loading 6.6 lbs per HP

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Exterior Dimensions

Note: Published ground clearance distances are based on properly inflated tires with oleo extension.

Cabin and Entry Dimensions

Cabin (Passenger/Cargo)
Length 473.5 inches
Width 54 inches
Height 71 inches
Airstair Entrance
Width 26.75 inches
Height 62 inches
Cargo Entrance
Width 52 inches
Height (fwd) 57 inches
Height (aft) 53.8 inches

Maximum Baggage Compartment Weights

Forward Closet
Total 250 lbs (includes hanger)
On Hanger 100 lbs
Cargo Bay
Fwd (w/o Lavatory) 1000 lbs
Fwd (w/ Lavatory) 800 lbs
Aft 630 lbs

Aircraft Volumes
Crew Station 103 ft3
Entrance & Closet 56 ft3 (Closet only: 17 ft3)
Main Cabin 584 ft3
Cargo Bay 175 ft3

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Minimum ground turning radius

Note: The published turning radius is predicated on the

use of partial braking and differential power as well as
the nose wheel being fully turned in the direction of
the turn.

Note: Due to the design of the aircraft, the tail of the

aircraft will clear any horizontal obstructions that the
wing tip has cleared as long as the turn radius is
maintained.

Aircraft Doors & Emergency Exits

Airstair Door
The airstair (cabin door) provides a built in stairway for entry and exit from the aircraft and can be opened from
inside or outside the aircraft. The airstair has 6 lighted steps for night operations and cable handhold on either
side of the airstair for passenger convenience. The airstair door opening is equipped with an inflatable rubber
seal, which seals the pressure vessel at the airstair door to allow the aircraft to pressurize during flight. Due to
the weight of the airstair door, it is equipped with a counterbalance assembly that utilizes a pressurized strut to
lower the airstair door gradually and to reduce the force required to raise and close the airstair. The correct
strut charge can be read from the gauge attached to the lower portion of the strut. The following value can be
observed on the Gauge connected to the cabin door snubber.

Airstair Open 1400 ± 50 PSI

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 Airstair Door Operation
Opening

To open the airstair door from the outside of the aircraft, push in and hold the button located above the
horizontal handle then rotate the handle clockwise as far as it will travel (nearly vertical). Move out from
under the airstair (toward the nose of the aircraft) and pull the door out and release it. Once the door
has been opened slightly, it will slowly fall under its own weight to the fully open position. Note: Always
move out from under the airstair when opening to avoid injury should the counterbalance assembly not
be charged correctly or should fail.

To open the airstair door from the inside of the aircraft, push in and hold the black button, located on
the left side of the chrome handle, then lift the handle up all the way. Slowly push the door out (about 1
foot) and visually check that nobody is below or near the airstair. Once the area is confirmed clear, push
the door out and release it. Once the door has been released, it will slowly fall under its own weight to
the fully open position. Note: Always release the airstair when opening to avoid injury should the
counterbalance assembly not be charged correctly or should fail.

Closing

To close the airstair door from the outside of the aircraft, lift the door up slowly (this allows the
counterbalance assembly to assist in supporting the weight of the door) and verify that the 4 support
cables (2 on each side of airstair) are not going to be caught in the door jam. Once the door is fully
closed, rotate the vertical handle counter-clockwise as far as it will travel (nearly horizontal). A click
should be heard and felt as it locks into position. At this point the door must be checked to ensure it is
secure by attempting to rotate the handle clockwise which should not move.

To close the airstair door from the inside of the aircraft, lift the door up slowly (this allows the
counterbalance assembly to assist in supporting the weight of the door) and verify that the 4 support
cables (2 on each side of the airstair) are not going to be caught in the door jam. Once the door is fully
closed, lower the chrome handle as far as it will travel. A click should be heard and felt as it locks into
position. At this point the door must be checked to ensure it is secure by; First, try and raise the handle
and it should not move, then push and hold the button on the right side of the 4th step which will
illuminate the locking mechanism within that step to verify that the latch has engaged, and lastly check
each of the 8 rotary camlocks, 4 on each side of the airstair, to ensure the orange strip on each camlock
aligns with the notch in its respective indicator pate.

Figure 1 Figure 1A

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Airstair Safety Systems
Monitoring the security of the airstair door, is accomplished through the 5 microswitches on the airstair
door (2x cam lock latches, 1x doorsill, 2x handle). If any of these microswitches indicate that the door in
not secure it will trigger the red “CABIN DOOR” annunciator.

The lock button mechanism prevents accidental movement of the handle and also incorporates a
differential pressure sensing diaphragm. The diaphragm compares atmospheric pressure to cabin
pressure and prevents operation of the lock button should the cabin pressure be greater.

Cargo Door
The cargo door provides a large opening to the cargo bay (or entire cabin if configured for cargo only) and is
normally opened from the outside of the aircraft, however, can be opened from the inside if required. The cargo
door opening is equipped with an inflatable rubber seal, which seals the pressure vessel at the cargo door to
allow the aircraft to pressurize during flight. Due to the weight of the cargo door, it is equipped with 2
pressurized struts that assist in raising and holding the door in the open position. A cable is attached to the
inside of the cargo door that assists in initially lowering the door.

Cargo Door Operation

Opening

To open the cargo door from the outside of the aircraft, push in and hold the button located above the
horizontal handle then rotate the handle clockwise as far as it will travel (nearly vertical) then pull on the
handle and open the door a couple feet. On the inside of the door remove the white cable from its
holder and then push the door upwards. Once the door is raised slightly, it will be raised automatically
by the pressurized struts.

To open the cargo door from the inside of the aircraft, the 5 quarter turn fasteners must be unlocked to
remove the protective shield from over the operating mechanism. Pull and hold the ring located to the
right of the handle then rotate the handle towards the nose of the aircraft (right) as far as it will travel
(nearly horizontal) and push the door out slightly. On the inside of the door remove the white cable
from its holder and then push the door upwards. Once the door is raised slightly, it will be raised
automatically by the pressurized struts.

Closing

To close the cargo door from the outside of the aircraft, pull the door down using the white assist cable
to within a couple of feet of closed. Fold the assist cable in half and secure it in the holder on the inside
of the door, then close the door. Once the door is fully closed, rotate the vertical handle counter-
clockwise as far as it will travel (nearly horizontal). A click should be heard and felt as it locks into
position. At this point the door must be checked to ensure it is secure by looking into the sight window,
located in front of the handle (towards the nose of the aircraft) and observing a red vertical line in the
center.

To close the cargo door from the inside of the aircraft, pull the door down using the white assist cable to
within a couple of feet of closed. Fold the assist cable in half and secure it in the holder on the inside of
the door, then close the door. Once the door is fully closed, rotate the horizontal handle towards the tail
of the aircraft (left) as far as it will travel (10 o’clock position). A click should be heard and felt as it locks
into position. At this point the door must be checked to ensure it is secure. At this point the door must
be checked to ensure it is secure by looking through the viewing window on the right lower side of the
door to verify that the orange line on the white background is lined up with the orange pointer.

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 Figure 2 Figure 2B

Cargo Door Safety Systems

Monitoring the security of the cargo door, is accomplished through the 5 microswitches on the door (2x
cam lock latches, 1x doorsill, 1x handle latch, 1x latch pin). If any of these microswitches indicate that
the door in not secure it will trigger the red “CARGO DOOR” annunciator.

The lock button mechanism prevents accidental movement of the handle and also incorporates a
differential pressure sensing diaphragm. The diaphragm compares atmospheric pressure to cabin
pressure and prevents operation of the lock button should the cabin pressure be greater.

Cabin Emergency Exits

Three emergency exits are available within the cabin at seats 4B, 6A and 6B. For aircraft configured for cargo
operations the emergency exits can be identified by window 3 right, 5 left and right. The door is a non-hinged
plug which removes completely from the frame once the latches have been released and has a quick release
electrical connector that allows for the safe disconnection of power from the plug when the door is removed.

Emergency Exit Operation

To open an emergency exits from the inside of the aircraft, pull the latch, located in the top center of
the exit, as far as it will travel out and up, then pull the plug into the cabin fully. This will disconnect the
electrical power from the plug and allow is not completely detached from the aircraft. To discard the
plug through the emergency exit opening, simply turn the plug on edge (see diagram).

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 Figure 3 Figure 3A

To open an emergency exit from the outside of the aircraft, push in the lock on the forward part of the
latch and pull the handle out and aft as far as it will travel, then push the plug into the cabin.

Note: Depending on the situation, the seat may be occupied and/or the seat back on the other side of
the exit may be fully upright and make opening of the exit more difficult to push into the aircraft.

Wing / Nacelle
Each wing is a single piece assembly that extends from the outboard edge of the left aileron to the outboard edge of the
right aileron and contains a one-piece spar. Each wingtip has an integrated winglet which enhances both climb
performance and aircraft stability while extending the effective wingspan to improve range. Integrated into each wing
are the main and auxiliary fuel tanks as well as the nacelle. Each aerodynamically designed nacelle contains the engine,
landing gear assembly and miscellaneous electrical systems as well as the air intake.

Empennage
The empennage is a conventional T-tail design which places horizontal stabilizer elevators out of the turbulent air
created by the propellers and exhaust. This design improves aircraft stability by minimizing the effects of power and flap
changes on the aircrafts pitch. Two stabilons (horizontal fins) have also been added to the rear of portion of the fuselage
to further increase the aircrafts pitch stability and expands the center of gravity envelope to an unusually wide and
beneficial 4 to 40% MAC. The stabilons also provide the aircraft with positive recovery characteristics from deep stall
regime. To increase the directional stability of the aircraft, 2 tailets (vertical fins) are positions on the bottom of the
horizontal stabilizer near each outboard tip.

Powerplants
The Beechcraft 1900D is equipped with two Pratt & Whitney PT6A-67D free-turbine reverse flow engines. Each engine is
rated at maximum of 1279 SHP and 1353 ESHP due to the jet thrust provided from the exhaust system. Each engine is
equipped with a composite four-blade, fully feathering, reversing, constant speed propeller.

Electrical System
The Beechcraft 1900D utilizes a multi-bus system that powers all the various electrical components. Power for these
systems can b e supplied by the aircraft’s battery, two generators, external power (ground power) and uses 2 inverters
to convert the DC power into AC for some of the avionics (and in UE-92 and lower for the torque gauges).

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Hydraulic System
The Beechcraft 1900D is equipped with hydraulically powered landing gear and brakes. The landing gear struts also use
the same type of hydraulic oil for damping landing forces.

Note: Optional power steering system is available for the 1900D which also utilizes hydraulic fluid from the brake
system.

Hydraulic System Capacity & Type

Landing Gear Capacity 2.75 Gallons

Brake System Capacity 0.5 Gallons
(1 Gallon with power steering)
Hydraulic Oil Type MIL-H-5606 (Red Colour)

Fuel and Oil System

The Beechcraft 1900D fuel system consists of an independent fuel system for each engine. Each system consists of 2 fuel
storage tanks, a collector tank, high and low pressure engine driven fuel pumps, standby electric fuel pump, jet transfer
pumps and an auxiliary electric fuel pump. The two systems are connected by a single cross-transfer line that allows fuel
to be transferred from one system to the other. The engine is lubricated by a closed oil system consisting of a pressure
pump, scavenge pumps, oil cooler, oil to fuel heater and a breather system.

Fuel Capacity & Type

Usable Fuel Pounds Litres Gallons
Total Fuel 4484 2518.8 665.4
Each Main Fuel Tank 1621 910.4 240.5
Each Auxiliary Fuel Tank 621 349.0 92.2
Standard Fuels: Jet A, A1, B

Oil Capacity
Quarts Gallons
Oil Tank 10 2.5
Oil Lines & Cooler 4.4 1.1
Total System 14.4 3.6

Environmental Systems
The aircraft bleed two air systems, one on each engine, to provide the aircraft with cabin pressurization, heating and
motive force to operating the air cycle machine (ACM) for cooling. The right engine is equipped with a vapour cycle
machine (VCM) to assist the ACM when extra cooling is required. The dual bleed air system is capable of maintaining a
maximum cabin differential of 5.1 psi (9000 ft cabin altitude when operating at the aircrafts service ceiling of 25,000 ft)
and has the capacity to produce 64,000 BTU of heat, 39,000 BTU of cooling on the ground, or 46,000 BTU of cooling
when in flight, with the assistance of the VCM. In the event of cabin depressurization or air contamination, the aircraft if
equipped with a supplemental oxygen system to provide the crew and/or passengers with oxygen if required.

Flight Controls
The ailerons, elevators and rudder are conventional surfaces, which are cable operated and do not require power to
assist in their movement. These control surfaces can be controlled from either pilot seat control wheel and each surface
has a cockpit adjustable trim tab. The flaps and optional trim system are electrically powered, along with the yaw
damp/rudder boost system (or autopilot system, if installed).

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Limitations
Airframe
Weight & Balance

Basic operating Weight 11,150 lbs (Fleet Average)

Includes
Maximum Zero Fuel Weight 15,165 lbs (Factory)
15,700 lbs (After Modification)
Maximum Ramp Weight 17,230 lbs
Maximum Takeoff Weight 17,120 lbs
Maximum Landing Weight 16,765 lbs
Note: The Takeoff and Landing Weight may be restricted lower due to
aircraft performance factors (see Performance Section).

Airspeeds

Speed (At 17,120 lbs) KIAS Remarks

Maximum Operating Speed (VMO) These speeds may not be deliberately
Sea Level to 13,200 ft 248 exceeded in any flight regime. Red and
13,200 ft to 25,000 ft 248 – 195 White Hash-Marked Pointer (Barber Pole)
(0.48 Mach) reflects VMO limit.
Manoeuvring Speed (VA) 178 Do not make full or abrupt control
movements above this speed.
Turbulent Air Penetration 170
Maximum Flap Extension/Extended Speed (VFE) Do not extend flaps or operate with flaps
Flaps 17 188 in the prescribed position above these
Flaps 35 154 speeds.
Maximum Landing Gear Operating Speed (VLO) Do not exceed this speed with the landing
Extension 180 gear extended.
Retraction 180
Emergency Descent 180
Best Glide (V) 140
Best Rate of Climb (VY & VYSE)
Two Engine (VY) 138 Speeds decrease 2 KIAS per 5000 ft
Single Engine (VYSE) 128
Best Angle of Climb (VX & VXSE)
Two Engine (VX) 122
Single Engine (VXSE) 122
Air Minimum Control Speed (VMCA) These are the lowest speeds at which the
Flaps 17 92 airplane is directionally controllable when
Flaps 35 92 one engine suddenly becomes inoperative
and the other engine is at takeoff power.
Conditions:
Stall Speed – Clean configuration (VSL)
17,120 lbs 100
Stall Speed – Landing Configuration (VSO)
17,120 lbs 84
Maximum Demonstrated Crosswind Component 22

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Engine
The following limitations shall be observed. Each Column presents limitations. The limits presented do not necessarily
occur simultaneously. Refer to the Pratt & Whitney Engine Maintenance Manual for specific actions required if limits are
exceeded.

Torque GAS Gen Prop Oil Temp

Operating Maximum Oil Press
SHP Ft-Lbs RPM RPM °C
Condition ITT °C PSI (2)
(1) % N1 N2 (3) (4)

Starting --- --- 1000 (5) --- --- 0 to 200 -40 (Min)

Idle --- --- 750 65 (Min) 950 (Min) 60 (Min) -40 to 110

Takeoff (8) 1279 3950 800 104 1700 90 to 135 10 to 110

Max 1214 3750 780 104 1700 90 to 135 10 to 105

Continuous

Cruise Climb 1105 3750 (5) 760 104 1700 90 to 135 10 to 105
and Max
Cruise

Max 900 --- 760 --- 1650 90 to 135 10 to 105

Reverse

Transient --- 5000 (7) 870 (7) 104 1870 (7) 40 to 200 -40 to 110

Footnotes:

(1) Torque limit applies within range of 1000-1700 propeller rpm (N2). Below 1000 rpm,
torque is limited to 2000 ft-lbs.
(2) Normal oil pressure is 90 to 135 psig at gas generator speeds above 72%. With
engine torque below 3000 ft-lbs, minimum oil pressure is 85 psig at normal oil
temperature (60 to 70C). Oil pressure under 90 psi is undesirable; it should be
tolerated only for the completion of the flight, and then only at reduced power
settings not exceeding 2000 ft-lbs torque. Oil pressure below 60 psi is unsafe; it
requires that either the engine be shut down, or that a landing be made at the
nearest suitable airport, using the minimum power required to sustain flight.
Fluctuations of plus or minus 10 psi are acceptable.
(3) A minimum oil temperature of 55C is recommended for fuel heater operation at
takeoff power
(4) Oil temperature limits are -40°C to 105°C. However, temperatures of up to 110°C
are permitted for a maximum time of 10 minutes.
(5) Starting ITT is time limited to 5 seconds.
(6) Cruise torque values vary with altitude and temperature
(7) These values are time limited to 20 seconds.
(8) Takeoff power is time limited to 5 minutes.

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 Engine Exhaust Heating of Cabin Windows and Airframe

1) Do not operate the engines with the propellers feathered, except during external power
starts and propeller feather checks, ex that the propellers may be operated in feather at
temperatures not to exceed +5°C for a maximum of 3 minutes for purposes of airframe
deicing.
2) Do not conduct static operations in ground fine when the OAT exceeds 38°C.
3) Do not back the airplane using reverse thrust for periods longer than 10 seconds.

Reduced Power Takeoffs

If the Static Reduced Takeoff Power setting specified in the AFM Performance section is utilized, the
following limitations must be observed:
The runways must not be wet or contaminated with snow, slush or ice.
Flaps must be set at 17° for takeoff.

The operator must establish a means to periodically check the availability of Static Takeoff Power, so
that the use of deteriorated engine(s) will not occur. One acceptable means is for the carrier or
operator to establish operating procedures which will require the use of Static Takeoff Power at
least once during each 25 hours of operating or 25 takeoffs, whichever occurs first. Other
equivalent procedures may be considered, based on the individual operational factors involved.

Starter Limits
If ignition occurs within 20 seconds of any start attempt, there is no limit on the time the starter is
engaged for that start.

For restarts following a normal start:

3 minute cool-down, beginning when the starter is turned off.

For an aborted start (no ignition) followed by an attempted restart:

Starter is limited to 20 seconds ON (aborted start), 30 seconds OFF, 20 seconds ON (seconds start
attempt), 60 seconds OFF, 20 seconds ON (third start attempt), 5 minutes OFF.

For continuous motoring without attempting to start the engine:

Starter is limited to 20 seconds ON, 5 minutes OFF. Repeat as required

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Instruments & Markings

Figure 4: Engine Instruments

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 AIRSPPED INDICATOR

92 KIAS AIR MINIMUM CONTROL SPEED

(VMCA)
84 to 154 KIAS FULL FLAP OPERATING
SPEED (VSO) AT MAXIMUM WEIGHT WITH
LANDING FLAPS AND IDLE POWER

248 KIAS / 0.48 MACH MAXIMUM SPEED

FOR ANY OPERATION
248-195 KIAS (13,000 TO 25,000 FEET) THIS
SPEED MAY NOT BE DELIBERATELY
EXCEEDED IN ANY FLIGHT REGIME

FUEL QUANTITY

0-363 LBS NO TAKE-OFF RANGE

PNUMATIC PRESUURE

17.5-20 PSI NORMAL OPERAYING RANGE

20 PSI MAXIMUM

GYRO SUCTION

HG 3.0-4.3 NORMAL OPERATING RANGE

35,000 - 15,000 FT.
HG 4.3-5.9 NORMAL OPERATING RANGE
15,000 – SEA LEVEL

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 CABIN ALTIMETER AND PRESSURE DIFFERENTIAL

0-5.2 PSI APPROVED OPERATING RANGE

5.2 PSI TO END OF SCALE (UNAPPROVED
OPERATING RANGE)

PROPELLER DEICER AMMETER

26-32 AMPERES NORMAL OPERATION

FLAP POSITION INDICATOR

FLAP 0 FLAP 17

FLAP 35

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BE02 Technical Manual
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Radios & Related Equipment

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Lighting

General Description
The Beechcraft 1900D is equipped 3 categories of aircraft lighting systems: Interior Lighting, which includes the cockpit
and passenger cabin; Entranceway lighting, for the airstairs and cargo door; and exterior lighting for navigation and anti-
collision.

Interior Lighting
Cockpit
The primary cockpit lighting is provided by several systems that are separated into light groups which are first
controlled by a Master Light Switch and then individually by 8 rheostats (rheostats allow for dimming control).
These Rheostats/light groups from left to right are:

1) PILOT FLIGHT LIGHTS

Controls the flight instrument backlights for the pilot’s Airspeed Indicator, Altimeter, RMI, VSI,
Clock, AHRS Control Panel, Standby Attitude Indicator Control Panel, EFIS Timer/Dimmer Panel,
EFIS Advisory/GPS Annunciator Panel and Flight Control Panel Buttons.

2) ENGINE INSTRUMENT LIGHTS

Controls the engine instrument backlights for the ITT, Torque, Propeller RPM, Turbine N1, Fuel
flow, Oil Temperature/Pressure and Propeller Sync Scope.

3) AVIONICS PANEL LIGHTS

Controls the avionics backlights in the Audio Control Panel, Communications and Navigation
Radios, ADF, Altitude Alerter, Transponder and GPS Button Lights

4) OVERHEAD FLOOD LIGHTS

Controls 2 overhead flood lights, which are located above each pilot’s position. Each light is
equipped with an additional switch which allows each pilot to turn ON/OFF their respective
flood light when the overhead flood lights switch is ON.

5) INSTRUMENT INDIRECT LIGHTS

Controls the 10 indirect lights, which are located under the glare shield. These lights have a blue
plastic lens on them and provide indirect light over most of the instrument panel.

6) EDGELIGHT PANEL LIGHTS

Controls the electroluminescent lighting systems for panel backlighting on the overhead panel,
upper and lower center consoles, left and right subpanels and circuit breaker panels as well as
the post lighting for the Trim Wheels.

7) OVERHEAD, SIDE PANEL, SUBPANEL & CONSOLE IND LIGHTS

Controls the instrument backlights on the Overhead Panel, Upper and Lower Center Consoles,
Right Subpanel and Fuel Gauges.

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 8) CO-PILOT FLIGHT LIGHTS

Controls the flight instrument backlights for the Co-pilot’s Airspeed Indicator, Altimeter, RMI,
VSI, Clock, AHRS Control Panel, EFIS Timer/Dimmer Panel, EFIS Advisory/GPS Annunciator Panel
and Flight Control Panel Buttons.

Other aftermarket installations may be connected into one of the above lighting controls or may be
controlled independently by the unit.

Figure 15: Cockpit Lighting

The rest of the cockpit lighting systems operate independently from the Master Light Switch and are listed
below:

Communications Panel Dim Knob – Controls the brightness of both FCP’s (Flight Director Control Panel),
Enables the avionics auto dim mode and brightness of the following supplemental annunciators:

(Both Altitude Alert, Drive Transfer, GPS LEG/OBS, GPS ARM/ACTV Panels), (TCAS/TAWS control Panel),
(Both AHRS Control Panel), (EFIS Aux Power Panel), (Standby Attitude Indicator Control Panel)

Figure 16

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 Map Light - Operates one light in each the pilot and co-pilot’s yokes and illuminates the area below the
yoke. The map light is controlled by an ON/OFF switch located on the top, inboard side of each yoke.

Emergency Cockpit lighting - Operates 4 of the indirect instrument lights, 2 over each pilot’s main flight
instruments. The switch for the emergency cockpit lighting is located on the bottom right side of the
overhead panel.

Figure 17: Instrument

Emergency Lights

Flashlights – The cockpit is equipped with 2 removable flashlights that are mounted on the lower right
side of the center console.

Cabin
Cabin lighting is provided by overhead panels located above each passenger seat and in the lavatory (if
installed). Control for the cabin lights are located on the lower right side of the overhead panel in the cockpit
and consist of 3 switches. These switches, from left to right are:

Figure 18: Cabin Lights

1) Cabin Lights Switch (3 Positions – FULL/ PARTIAL/OFF)

Full – All passenger overhead panel flood lights ON along with the under seat aisle lights

Partial – If the Battery is ON, passenger overhead panel flood lights at seats 1B, 3A, 5B, 7A and
9B are ON, along with the under seat aisle lights. However, if the Battery is OFF, only the
passenger overhead panel flood lights at seats 1B and 7A are ON.

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 Off – All passenger overhead panel flood lights OFF

2) Reading Lights (2 Position – ON/OFF)

This controls the reading light on each passenger’s overhead panel. When selected ON, it
enables the reading light switch on each overhead panel to control is respective reading light.

3) FSB – Fasten Seat Belt (2 Position – ON/OFF)

Controls the “No Smoking/Fasten Seat Belt” sign in lavatory.

In addition, the hump in the aisle between seats 4A and 4B is equipped with glow tape and the emergency exit
signs made with T-lights (glass tubes filled with tritium gas), both systems require no power to function.

Figure 18: Cabin Lights

Entranceway lighting
Airstairs
The main airstair lighting consists of a light in the vertical section of each step and a threshold flood light
attached the bulkhead on the aft side of the cabin entrance. The 2 position switch (ON/OFF) for controlling the
airstair lights is a located on the left side of the third step of the airstair door. When the airstair lights are
selected ON, there is a microswitch located at the base of the door that will turn off the lights when the door is
closed.

The airstairs also contain an internal inspection light that is used to view the internal locking mechanism through
an inspection window located in the vertical section of the step above the 3rdstep. The light is controlled by a
momentary pushbutton switch that must be held on as long as desired.

Cargo Bay
The cargo bay lighting consists of two overhead flood lights, one above the lower cargo area and one above the
upper cargo area and are controlled by an ON/OFF switch that is located on the inside of the fuselage, on the
wall just behind the cargo door opening. When the cargo lights are selected ON, there is a microswitch located
at the base of the cargo door that will turn off the light when the door is closed.

Exterior Lighting
All the various exterior lighting systems are controlled from the cockpit by switches that are located on lower
section of the overhead panel.

1) LANDING - LEFT (2 Position – ON/OFF)

Switch controls the Left Landing light, which is located in the leading edge of the left wing, outboard of
the engine nacelle.

2) LANDING – RIGHT (2 Position – ON/OFF)

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 Switch controls the Right Landing light, which is located in the leading edge of the right wing, outboard
of the engine nacelle.

3) TAXI (2 Position – ON/OFF)

Switch controls the single taxi light, which is located on the upper portion of the nose gear.

4) ICE (2 Position – ON/OFF)

Switch controls the ice inspection lights, which are located on the outboard side of each engine neccelle,
ahead of the leading edge of the wing.

5) NAV (2 Position – ON/OFF)

Switch controls the Left, Right and Tail Navigation Lights, which are located on the left and right wingtip
light assemblies and on the rear top of the tail. Each location contains 2 bulbs with the applicable
coloured lens: Left Wing - Red, Right Wing - Green, Tail – White (No Lens), and are dual powered. The
Forward Left, Forward Right, Left Tail Navigation Lights are powered by the Center Bus and the Rear
Left, Rear Right and Right Tail Navigations Lights are powered by the Right Generator Bus.

6) ANTI-COLLISION – BEACON (3 Position – FLT/GND/OFF)

Switch controls the Anti-Collision beacons, which are located on underside of the fuselage and on the
top of the horizontal stabilizer.

FLT – Beacon is ON and powered by the Left Generator bus. This position may be selected once
the aircraft is operating under its own power.

GND – Beacon is ON and powered by the Center Bus. This position is selected prior to start and
allowed the Beacon to be ON when starting on the aircraft’s internal battery

OFF – Beacon is OFF

7) ANTI-COLLISION – STROBE (2 Position – ON/OFF)

Switch Controls the Anti-Collision Strobe lights, which are located on each wingtip on the outboard side
of the winglets.

8) RECOG (2 Position – ON/OFF)

Switch controls the Recognition Lights, which are located in the left and right wingtip light assemblies
beside the navigation lights.

WARNING: Heat from the recognition light could damage the wing tip lens if operated on the ground.

9) TAIL FLOOD (2 Position – ON/OFF)

Switch controls the Tail Flood lights (also called logo lights), which are located on the underside of the
horizontal stabilizer and contains 2 bulbs.

Figure 19: Exterior Lights

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ANNUNCIATOR SYSTEM

General Description
The annunciators are separated into 2 main panels. The upper panel located on the glareshield contains all the red -
warning annunciators (except for the fire T-Handles) and the lower panel located in the upper center console contains
all the yellow - caution, Green - advisory and white - advisory annunciators. All the annunciators bulbs may be tested by
holding the PRESS TO TEST button located on the right side of the upper annunciator panel (Note: The Press To Test
button must be held longer than 5 seconds so that the annunciators that are triggered by the Fluid Level Sensors will
activate after their built in time delay is passed)

Master Flashers
The Master Flashers consist of a Red - Master Warning and Yellow - Master Caution located on the glareshield directly in
front of each pilot seat. The Master Warning or Caution flasher will activate when their respective Red or Yellow
annunciator illuminates to direct the pilot’s attention to the problem. The Master Warning and/or Caution will continue
to flash until it is pressed. However, the annunciator that triggered the Master Flasher will remain illuminated until the
fault is corrected. Note: There is NO flasher associated with either the green or white advisory annunciators.

Figure 20: Master Warning / Master

Caution Flashers

Figure 21

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Annunciators

Master Warning “Red Annunciators”

L FUEL PRES LO Trigger: Pressure sensed in the Collector Tank fuel line is < 10PSI

R FUEL PRES LO Possible cause: Failure of engine driven low pressure fuel pump
Note: Engine driven low pressure fuel pump is installed on the same drive
shaft the 2 external oil scavenger pumps. Monitor oil temperature/pressure
after illumination of this annunciator.

CABIN ALT HI Trigger: Cabin altitude sensed at 10,000’ +0 / -500

Possible cause: Cabin altitude exceeds 10,000ft
Note: Sensor activates at 9,500’ to 10,000’ when cabin altitude is climbing
and deactivates at a maximum of 8,500’ when cabin altitude is descending.

CABIN DIFF HI Trigger: Cabin differential pressure sensed at > 5.25 ± 0.15 PSI
Possible cause: Cabin differential pressure high
Note: Nominal cabin pressure is 5.0 ± 0.1 PSI and overpressure relief valve
(Poppet) activates at 5.1 PSI

L OIL PRES LO Trigger: Oil pressure sensed at < 60 PSI

R OIL PRES LO Possible cause: Oil pressure failure

Note: The L/R OIL PRES LO annunciators and oil pressure gauge receive their
data from separate sensors; both should to be used to confirm a problem.

L ENVIRO FAIL Trigger: Disagreement between switch position and valve position or valve is
not powered
R ENVIRO FAIL
Possible cause: Environment air duct over pressure (Activates at > 44 ± 1 PSI)
and/or over temperature (Activates at > 500 ˚F ± 10˚F)
Note: Nominal air duct pressure of 37 ± 3 PSI and temperature of 450˚F ±
25˚F. Both annunciators also illuminate as part of the T-Test system check.

Note: The L ENVIRO FAIL annunciator will illuminate whenever the left
generator bus is unpowered.

CABIN DOOR Trigger: Any of the 5 internal microswitches is indicating that the door is NOT
secure.
Possible cause: Cabin door open or not secure
Note: The 5 microswitches are: 2x cam lock latches, 1x doorsill, 2x handle
latches

CARGO DOOR Trigger: Any of the 5 internal microswitches is indicating that the door is NOT
secure.
Possible cause: Cargo door open or not secure
Note: The 5 microswitches are: 2x cam lock latches, 1x doorsill, 1x handle
latch, 1x latch pin

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L BL AIR FAIL Trigger: Warning line tubing pressure sensed at < 1.5 ± 0.5 PSI

R BL AIR FAIL Possible cause: Melted or failed bleed air failure warning line or system off
Note: Nominal pressure in warning line (EVA tubing) is 17.5 – 19.0 PSI. Both
annunciators also illuminate if both ENVIRO switches are placed in the INST &
ENVIRO OFF position.

L AC BUS Trigger: The 115v AC Bus is unpowered

R AC BUS Possible cause: AC Bus has an inoperative inverter or either 115v bus or 26v
bus has a short
Note: When the L/R 115v AC Bus is unpowered the same side 26v AC Bus is
unpowered. Either annunciator will also illuminate when the respective AC
Bus switch is turned off.
A/P TRIM FAIL Trigger: Autopilot trim servo has reached its maximum operational limit

(If Autopilot Installed) Possible cause: Improper trim or no trim from autopilot trim command
Note: Pilot flying should hold yoke with both hands for autopilot disconnect
as force required to maintain the aircraft's attitude may be higher than
normal.

A/P DISC Trigger: Autopilot data computer failure

(If Autopilot Installed) Possible cause: Autopilot is disconnected by means other than pilot’s switch

Note: Crew may re-engage autopilot after selecting the opposite computer.
However, the pilot flying must be on the same side as the autopilot data
computer being used.

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 Figure 22

Master caution “Yellow Annunciators”

L DC GEN Trigger: Generator line contactor is open

R DC GEN Possible Cause: Generator is offline

Note: See electrical section for line contactor control details
L FUEL QTY Trigger: Fluid Level Sensor is detecting less than 324 lbs of usable fuel

R FUEL QTY Possible Cause: Fuel quantity is below 324 lbs of usable fuel
Note: Fluid Level Sensor has a built-in 5 second delay prior to activation.
L COL TANK LOW Trigger: Fluid Level Sensor is detecting less than 53 lbs of usable fuel in the
collector tank
R COL TANK LOW
Possible Cause: Fuel system collector tank below 53 lbs of usable fuel
Note: Remaining fuel is approximately 8 minutes if configured for normal
cruise at 160 knots and 10,000’
STALL HEAT Trigger: Insufficient current being supplied to the stall heat transducer
Possible Cause: Insufficient current to provide enough heat on stall warning
transducer to prevent icing
Note: System only detects if there is insufficient heat to prevent ice in ground
mode. There is NO warning system for insufficient heat during flight mode
and a failure can only be identified by visual means. Annunciator also
illuminates if the Stall Heat switch is in the OFF position.
BATTERY CHARGE Trigger: Battery charge rate is > 7 amps for > 6 seconds
Possible Cause: Excessive charge rate on battery, monitor for thermal
runaway (see Electrical Section for details)
Note: This annunciator normally illuminates after engine starts and must be
monitored for a decreasing charge rate. Annunciator extinguishers when
charge rate is < 6 amps.

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L FW VALVE Trigger: Firewall fuel valve has not reached its selected position

R FW VALVE Possible Cause: Firewall fuel valve has not reached its selected position

Note: When the Firewall Fuel Valve is opened or closed the respective L/R FW
VALVE annunciator should blink quickly (not long enough however to trigger
the Master Caution Flasher).

L GEN TIE OPEN Trigger: Generator bus tie is open

R GEN TIE OPEN Possible Cause: L/ R Generator bus tie HED has isolated the L/ R generator
bus from the center bus
Note: The respective Generator and Generator Bus have been isolated from
the Center Bus, Battery Bus and opposite Generator Bus. Either triggered by
the respective HED(s) or the generator ties have been manually opened. See
Electrical System and Abnormal Procedures checklist for details.

BAT TIE OPEN Trigger: Battery bus tie is open

Possible Cause: Battery tie HED has isolated the battery from the center bus
Note: The Battery is no longer being charged. The battery is still powering any
items on the battery bus and hot battery bus.
L ENG ICE FAIL Trigger: Disagreement between switch position and ice vane position.

R ENG ICE FAIL Possible Cause: Ice vane has not reached its selected position within 30
seconds.
Note: If neither actuator will correct the problem, the crew must consider
flight conditions for potential icing as well as the effect on aircraft
performance.
L BK DI OVHT Trigger: Warning line tubing pressure sensed at < 1.5 ± 0.5 PSI

R BK DI OVHT Possible Cause: Melted or failed brake de-ice failure warning line or system is
turned off.
Note: Nominal pressure in warning line (EVA tubing) is 17.5 – 19.0 PSI. Both
annunciators also illuminate if both ENVIRO switches are placed in the INST &
ENVIRO OFF position.
HYD FLUID LOW Trigger: Fluid Level Sensor is detecting a low quantity of hydraulic fluid in the
reservoir.
Possible Cause: Landing gear hydraulic fluid level low.
Note: Fluid Level Sensor has a built-in 5 second delay prior to activation.
ANTI SKID FAIL Trigger: Anti-skid Computer Failure.
Possible Cause: Anti-skid computer electrical fault or low hydraulic oil
pressure in the antiskid brake system.
Note: Anti-skid should be selected OFF and normal performance charts from
the AFM shall be used for all calculations (These charts assume the anti-skid
system is not installed).

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ANN PWR SOURCE Trigger: Loss of power to one or both of the power sources that supply the
sensors that trigger certain annunciators.
Possible Cause: Annunciator sensors power supply failure.
Note: The annunciator bulbs are not affected by this failure and will test
successfully, but the sensors that trigger certain annunciators may not work.
(See Annunciators Section and B1900D MEL for details).
L FIRE LOOP Trigger: Fire detection loop continuity monitoring has failed.

R FIRE LOOP Possible Cause: Fire detection loop is broken in one or more places.
Note: The fire detection system may no longer be able to detect an engine
fire (See Fire Protection for Details).
L PITOT HEAT Trigger: Insufficient current being supplied to the pitot-static mast

R PITOT HEAT Possible Cause: Insufficient current to provide enough heat on pitot-static
mast to prevent icing.
Note: Annunciator also illuminates if the pitot switch is in the OFF position.
XFR VALVE FAIL Trigger: Disagreement between switch position and valve position
Possible Cause: Fuel transfer valve has not reached the selected position
within 2 seconds
Note: See Fuel System and Abnormal Procedures Checklist: Transfer Valve
Failure for details.
MAN STEER FAIL Trigger: Disagreement between switch position and power steering actuator
position
(If power steering installed)
Possible Cause: Power steering system actuator has not returned to free
caster mode
Note: Nose wheel won’t free caster with power steering dis-engaged
L NO AUX XFR Trigger: Pressure sensed in fuel line between Aux tank to collector tank is < 5
PSI
R NO AUX TFR
Possible Cause: Switch in AUTO position; indicates a failure of the float valve
to shutoff the aux pump. Switch in ON position; indicates that the tank is
either empty or the aux pump has failed.
AUTOFTHER OFF Trigger: Autofeather switch is in the OFF position and the left main landing
gear is fully not retracted.
Possible Cause: Autofeather system turned OFF with landing gear selected
down
Note: Autofeather must be armed for take-off, climb, approach and landing.
The exception to this is when on a single-engine approach this annunciator
will illuminate and should be acknowledged and left OFF.
PITCH TRIM OFF Trigger: Trim system has been de-energized by either yoke disconnect switch
and the pitch trim switch is ON.
Possible Cause: Trim system has been de-energized by either yoke disconnect
switch and the pitch trim switch is ON.

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AFX DISABLE Trigger: Autofeather system logic has disabled the autofeather capability
Possible Cause: Autofeather system is not capable of feathering the
propellers
Note: System logic prevents feathering both propellers at the same time
and/or prevents feathering the propeller on the good engine following an
engine failure.
YD/RB FAIL Trigger: Failure of the Yaw damper/Rudder Boost computer

(If autopilot NOT installed) Possible Cause: Yaw Damp/Rudder Boost computer has failed
Note: Crew may re-engage Yaw damper after selecting the opposite
computer.
RUD BOOST OFF Trigger: Rudder boost system is off

Possible Cause: Rudder boost switch is in the OFF position

Advisory “Green Annunciators”

INBD WG DEICE Trigger: Pressure sensed in respective de-ice boots is > 12 PSI

OUTBD WG DEICE Significance: Pressure in the respective de-ice boots is sufficient for de-icing
Note: Sensor only detects whether enough pressure is present for de-icing
TAIL DEICE and not actual boot inflation. It is possible to receive the correct annunciation
without boot inflation, always verify inflation visually.
L AUTOFEATHER Trigger: Autofeather switch armed and power leavers advanced above 85-
90%
R AUTOFEATHER
Significance: Autofeather system logic has enabled the autofeather capability
Note: The annunciator indication is duplicated for ease of viewing above the
top left of the torque gauges and illuminates “AFX”
L IGNITION ON Trigger: Engine ignitor is powered

R IGNITION ON Significance: Engine Ignitor is powered

Note: during start/ignition or with auto-ignition armed and torque < 700 ft-
lbs
L ENG ANTI-ICE Trigger: Engine anti-ice vane is fully extended

R ENG ANTI-ICE Significance: Engine anti-ice vane has reached its fully extended position

L BK DE-ICE ON Trigger: Brake de-ice bleed air valve is in the open position

R BK DE-ICE ON Significance: Use of brake de-ice will lower aircraft performance and raise ITT
Note: The ITT will raise approx 18°C when selected on during flight and
torque will decrease slightly.

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MAN TIES CLOSE Trigger: Generator tie switch has selected to manually closed
Significance: Load shedding will NOT automatically occur and/or battery is
supplying all the electrical power for the system
Note: Crew must be aware of discharge rate and power remaining (See
Emergency Procedures Checklist: Load Management)

Advisory “White Annunciators”

L ENVIRO OFF Trigger: Environmental bleed air valve is closed

R ENVIRO OFF

RDR PWR ON Trigger: Radar power is on and aircraft is on the ground

(If annunciator system installed)

FUEL TRANSFER Trigger: Fuel transfer valve is in the fully open position

TAXI LIGHT Trigger: Taxi light switch is on and the landing gear is fully
retracted.
Note: Taxi light will not be cooled when
EXTERNAL POWER Trigger: External power plug has been connected to the
aircraft
Note: Only indicates a plug is connected, not if power is
available

Fluid Level Sensor

The Fluid Level Sensors are used for triggering the following annunciators: “L FUEL QTY”, “R FUEL QTY”, “L COL
TANK LOW”, “R COL TANK LOW” and “HYD FLUID LOW”. The sensor operates by shining light from the top of the
device onto a prism at the bottom and detecting the amount of light which returns to a light sensor the top.

There are 3 modes of operation:

Fluid is Sufficient: When there is fluid above the prism level, the fluid deflects the
light away and not enough light return to the light sensor to trigger the
annunciator.
Fluid is Insufficient: When the fluid is below the prism level, enough light is
returned from the prism to the light sensor to trigger the annunciator. Note: The
light sensor has a built-in 5 second delay to allow for temporary movement of the
fluid (turbulence, turns during taxiing) prior to triggering the annunciator.
System Test: When the Press To Test button is held on the upper annunciator
panel, it illuminates a light in the Fluid Level Sensor located right beside the light
sensor for testing purposes. Note: The 5 second built-in delay still must be met
prior to triggering the annunciator.

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Annunciator Dimming Feature
The annunciators will automatically dim to the appropriate level if the following conditions are met:

1) At least 1 generator is online

2) Master Panel light switch is ON

3) Pilot Flight lights are ON

4) Overhead Flood Light switch is OFF

5) Photocell must detect light below a preset level (See below)

The exception to this is that the following annunciators will never dim: Master Warning Flasher and Firewall Fuel
Valve (T-Handle).

Figure 22: Photo cell

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FIRE PROTECTION

General
The Beechcraft 1900D fire protection systems are designed to provide detection and prevention and/or
extinguishing of a fire with the following systems: Engine Fire Protection, Interior Fire Protection and Bleed Air
Fire Protection.

Engine Fire Protection

Detection System Operation
Each Engine is independently protected 2 thermally sensitive fire-zone cables which are connected at one end
and routed around the engine in a continuous loop. Principle of operation: The loop is constructed with a centre
wire surrounded in a semi-conductive material (thermistor) within a stainless steel sheath. The thermistor
restricts the flow of electricity from the center wire to the outer sheath; and as temperature increases, the
resistance of the thermistor decreases and allows more electricity to flow.

Cables (fire loops) from both engines are then connected to a control amplifier, located in the nose of the
aircraft. The control amplifier activates its respective T-Handle when the resistance in the cable drops below
approx 100 Ohms (which is approx 315°C to 483°C (600°F to 900°F) on the forward cable and 248°C to 382°C
(480°F to 720°F) on the aft cable). To verify that the fire loops are operational, the control amplifier continuously
monitors the fire cable for continuity (uninterrupted path from beginning to end). Should the fire cable ever
break, the respective “L/R FIRE LOOP” annunciator will illuminate to alert the crew to a failure of the fire
detection system. The control amplifier also employs short circuit discriminating and time delay circuitry to
prevent false indications.

NOTE: If a single break in the fire loop where to occur, the fire protection would remain fully functional.
However, if the fire loop where to break in multiple locations, the fire protection would be degraded as
the segment of wire between the breaks would no longer be monitored and a fire would have to spread
to a functional portion of the fire cable before an annunciation would be received. There is no way for
the crew to know which of the above faults have occurred and must be repaired prior to flight.

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 Extinguishing System Operation
Each engine is equipped with a single fire extinguisher bottle, located in the outboard side of the gear bay. The
extinguisher is filled with 2.10 lbs of Bromotrifloromethane (Commonly known as HALON 1211) charged to 360
psi at approx 21°C (70°F).

The left or right extinguisher is activated by lifting respective the clear protective cover (breaking the witness
wire) and pushing the button. This activates the explosive squib in the extinguisher bottle and releases the
extinguishing agent into the engine nacelle. Once all the extinguishing agent is used the amber “DISCH” (“D” UE-
92 and lower) illuminates to indicate a discharged extinguisher bottle.

Note: The fire extinguisher squib is powered by both the Center Bus and the Hot Battery Bus, which
means that even with the Battery Master switch OFF, the extinguisher can still be activated if the
“EXTINGUISHER PUSH” button is pressed.

Extinguisher Discharge

Interior Fire Protection (Cockpit/Cabin/Cargo/Lavatory)

Detection System Operation
The Cockpit, Cabin and Cargo areas have no fire detection systems. The lavatory (If Installed) is protected by
either a common household style smoke detector or a build-in Lavatory Smoke Detection System (Powered by:
Triple Fed Bus). The Lavatory Smoke Detection System is located beside the mirror on the aft bulkhead of the
lavatory and when activated trigger’s the red “LAVATORY SMOKE” annunciator.

Extinguishing System Operation

The only extinguishing systems available in the Interior of the aircraft are 3 portable handheld fire extinguishers,
which are located; under the co-pilot seat, on the aft bulkhead of the closet and under seat 9A. The fire
extinguishers are filled with Bromotrifloromethane (Commonly known as HALON 1211).

NOTE: HALON 1211 is a colorless gas that leaves no residue and works by chemically inhibiting the
combustion process as well as cooling the surface. After discharging HALON 1211 the area should be
adequately ventilated. From HALON 1211 MSDS Sheet: “Exposure to concentrations of this material
above 4% for longer than one minute can cause toxic side effects. These can include dizziness, impaired
coordination, reduced mental acuity and cardiac effects. Higher concentrations with longer exposures
can cause unconsciousness or even death.”

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Bleed Air System Fire Protection
Detection System Operation
In close proximity to each of the 4 main bleed air lines there is a small EVA (ethylene vinyl acetate) tube that is
pressurized from the pneumatic system to 17.5 – 19 PSI. In event a bleed air line should rupture, the escaping air
will melt the EVA tube and the pressure in that EVA tube will drop below 1.5 ± 0.5 PSI which causes its
respective annunciator will illuminate: “L BL AIR FAIL”, “R BL AIR FAIL”, “L BK DI OVHT”, and “R BK DI OVHT”.

Note: Any time there is less than 1.5 ± 0.5 PSI in the EVA tube its respective annunciator will illuminate.
For example: If both pneumatic sources were turned off, all 4 of the annunciators would illuminate.

Fire Prevention System Operation

After receiving an “L BL AIR FAIL” or “R BL AIR FAIL” annunciator, the only way to prevent a fire or heat damage
from occurring is to turn off the entire bleed air system on the effected side. As per the Emergency Procedures
Checklist, this is done by placing the ENVIRO switch, located on the right lower subpanel, to the “INSTR &
ENVIRO OFF position. This will turn the respective Left or Right bleed air line off at a point just past the pre-
cooler. To verify that the bleed air valve has closed and no more air is escaping from the bleed air line, the crew
monitor the engine instruments for a slight increase in torque and decrease in ITT on the same side as the closed
bleed air valve.

Note: This will not extinguish the annunciator as the sensor will still detect less than the nominal
pressure. The cause of the rupture and repair of the EVA tube must both be repaired prior to the
annunciator extinguishing.

To prevent a “L BK DI OVHT” or “R BK DI OVHT” annunciation, if the brake de-ice has been selected ON then the
gear was retracted, a timing circuit is activated which will automatically turn off the brake de-ice system after 10
minutes to prevent any possible damage to the wheel well or adjacent components.

Note: If the brake de-ice system has been turned off by the automatically, the system cannot be
activated again until the gear as been cycled from the retracted position.

After receiving a “L BK DI OVHT” or “R BK DI OVHT” annunciator, as there is no checklist procedure provided by

Beechcraft; it is recommended that the brake de-ice switch be turned OFF and, if the gear was retracted after
the use of the brake de-ice system, that the gear be lowered so that the hot bleed air that is flowing from the
rupture can exit the aircraft without causing damage and provide the maximum amount of airflow in the main
gear wheel wells and around the main landing gear to cool the area where the bleed air line is routed.

Note: This will not extinguish the annunciator as the sensor will still not detect the 18 PSI minimum
pressure. The cause of the rupture and repair of the EVA tube must both be repaired prior to the
annunciator extinguishing.

Brake Deice

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System Checks
Engine Fire Detection Test
The engine fire detection system is tested though a single switch that has 3 positions (1 switch for each engine)
and is located in the cockpit, on the lower right subpanel. Pull both Firewall Fuel Valves out to the CLOSED
position. Hold the L Detect Switch to the AMP position and verify that the left firewall fuel valve red lights and
the left amber “Extinguisher Push” annunciator illuminate to indicate that the respective amplifier can
successfully detect a fire. Now hold the L Detect Switch to the LOOP position and verify that the amber “L FIRE
LOOP” annunciator illuminates to indicate that the fire loop has continuity (no breaks). Release the switch and
complete the same procedure for the right side.

SOP – Engine Fire Detection Test

Engine Fire Extinguishing System Test

The Fire Extinguishing System is tested to through a single switch that has 3 positions (1 switch for each engine)
and is located in the cockpit, on the lower right subpanel. Hold the L Ext Switch to the TEST A position and verify
that the left amber “DISCH” (or “D” on UE-92 and lower) and the green “OK” annunciators, located on the
respective glareshield illuminate to indicate a successful test of the squib and diaphragm functionality. Now hold
the L Exit Switch to the TEST B position and verify that the green “OK” annunciator illuminates to indicate a
successful test of the secondary wiring to the squib. Release the switch and complete the same procedure for
the right side.

Also, during the pre-flight inspection, the crew must look into the gear bay and verify that the extinguisher
bottle contains the correct charge as per the chart below.

SOP – Engine Fire Extinguishing System Test

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 Bleed Air Detection System Test
The bleed air system fire protection is tested by triggering the respective annunciators and verifying correct
function. While both engines are running, verify that the Vacuum Gauge is in the green range (4.3 - 5.9 HG), the
Pneumatic Pressure gauge is in the green range (17.5 - 20 PSI). Select the L Bleed Air Valve to the INST & ENVIRO
OFF position and verify that the vacuum and pressure gauges remain in the green and no annunciators
illuminate which confirms that the R Bleed Air system can adequately supply the system. Select the R Bleed Air
Valve to the INST & ENVIRO OFF position and verify that the Vacuum Gauge and Pneumatic Pressure gauge both
indicate “0” and the “L BL AIR FAIL”, “R BL AIR FAIL”, “L BK DI OVHT” and “R BK DI OVHT” annunciators illuminate
to confirm that the warning systems activate when a loss of pressure in its respective tubing is lost. Select the L
Bleed Air Valve to the OPEN position and verify that the Vacuum Gauge and Pneumatic Pressure gauge return to
the green range and the above annunciators extinguish to confirm that the L Bleed Air system can adequately
supply the system. Select the R Bleed Air Switch to the OPEN position.

SOP – Bleed Air Detection System Test

Lavatory Smoke Detector System Test

The smoke detector system is tested through the self test features on the unit in the lavatory. There are two
different systems that may be installed on the aircraft. The test procedure for each system is described below.

If the household style smoke detector is installed:

1) Press and hold the test button and observe aural warning
2) Release the test button

The Lavatory Smoke Detection System is tested by:

1) Observe the green “Power Indicator” light is illuminated

2) Push and hold the “self test switch” and observe the “LAVATORY SMOKE” annunciator in the cockpit
illuminates and the red “alarm indicator” light illuminates,
3) While holding the “self test switch”, momentarily press the “interrupt switch” and observe the
“LAVATORY SMOKE” annunciator extinguishers, but the red “alarm indicator” light remains
illuminated.
4) Release the “Self test switch” and observe the red “alarm indicator” light extinguishes

SOP – Lavatory Smoke Detector System Test

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ELECTRICAL SYSTEM
General Description
The Beechcraft 1900D utilizes a multi-bus system that powers all the various electrical components. Power for these
systems can b e supplied by the aircraft’s battery, two generators, external power (ground power) and uses 2 inverters
to convert the DC power into AC for some of the avionics (and in UE-92 and lower for the torque gauges).

Electrical Terminology
DC –Direct Current – Electricity flows in a constant direction through the conductor.

AC –Alternating Current – Electricity flow back and forth through the conductor on a set frequency. The
frequency is measured in hertz (Hz).

V – Voltage – Measure of electrical force.

A or Amps – Amperage – Measure of electrical flow.

Ω - Resistance – Measure of opposition to electrical flow.

Electrical Symbols
Battery Power Supply

Fuse Excessive current protection

device that is single use
Circuit Breaker Excessive current protection
device that is resetable

Switch On/Off switch (multiple types:

toggle, gated, spring loaded,
momentary)
Switch with Circuit Breaker On/Off Switch with built-in
(Identified by stamped excessive current protection
amperage rating on end of that is resetable
switch)
Diode One-way valve for electricity
(Current can only flow in the
direction of the arrow)
Current Limiter Allows only a pre-set amount
of current to pass through. If
an extremely high amount of
current tries to pass, the
device will fail like a fuse and
open the connection.
Relay Relays are used to allow one
circuit to control another
circuit without connecting
the two.

A normally open (N.O.)

Normally Open contacts connect the circuit
when the relay is activated.

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 A normally closed (N.C.)
contacts disconnect the
circuit when the relay is
activated.
Normally Closed
Bus Tie Sensor
(Hall Effect Device – HED)

Ground Common connection for

neutral (negative) lines.

DC Power Distribution
With the engines running, DC power is normally available from the left and right generators, along with the aircraft’s
battery. Without the engines running power is available from the battery, and if connected, the external power. The
power is distributed to various electrical systems though 6 main buses: Left Generator Bus, Right Generator Bus, Center
Bus (Left and Right are always connected), Battery Bus, Hot Battery Bus and the Triple Fed Bus. During normal
operations all the buses are tied together and protected by; bus ties, diodes or current limiters which all provide the
ability to isolate a bus from the rest of the system.

The hot battery bus is always powered directly from the battery regardless of the position of the battery master switch.
This means that any systems powered from this bus, if left on, will deplete the charge battery over time if there is no
other power source to recharge the battery. When the battery master switch is turned ON, the battery relay and battery
bus tie both will close. Power can now flow through the battery relay to the battery bus and triple fed bus as well as
through the battery bus tie to the center buses and starter systems. Since only the battery is powering the system at this
point, both generator bus ties will remain open. Once either engine is started and is respective generator is brought
online both generator ties will close automatically and power will flow from the online generator’s bus to: the triple fed
bus, through the center bus to the opposite generator bus and through to the triple fed bus, as well as, through the
center bus and battery bus tie to the battery to allow the battery to recharge and through to the triple fed bus. At this
point all the aircraft systems are being powered. Once the other engine is started and its respective generator is brought
online, both generators will share the aircraft’s electrical requirements.

The triple fed bus is fed from both generator buses and the battery bus. Each of these connections contain a one way
diode that allows power to only flow into the triple fed bus and a 60 Amp current limiter, which will isolate that
connection from the triple fed bus in the event of excessive current flow. Once a current limiter has opened the circuit,
the connection can only be restored by maintenance. As the triple fed bus is powered from 3 sources, the bus would still
remain fully functional with any 2 of the connections isolated. However, should a short occur on the triple fed bus; all 3
current limiters would completely isolate the bus and all associated electrical items would not be available. Note:
Whenever power flows through a one way diode, there is a loss of approximately 1 volt from the total supplying voltage.
Therefore, when checking the different bus using the volt-meter, the triple fed bus will indicate about 1 volt less than
the other buses.

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For monitoring the aircraft’s electrical system, it is equipped with 2 loadmeters for monitoring each generator, a bus
selectable volt-meter for monitoring individual bus voltage and battery ammeter to monitor the rate of charge or
discharge. The loadmeters, volt-meter and battery ammeter are located on the bottom-center portion of the overhead
panel.

Overhead panel: load gauges

Battery
The aircraft is equipped with a 24 Volt, 20-cell, nickel-cadmium (Ni-Cad) battery rated at 36 Amp/hours. This rating
means that a battery in good condition, with a full charge, should supply 36 Amps at 24 VDC for 1 hour prior to the
voltage (and therefore power) dropping rapidly to nearly zero. For example, if the aircraft is operated on the battery
alone with a current discharge of 72 Amps, the battery should provide power for 30 minutes of operation. However, if
the discharge rate were 12 Amps, the battery should provide power for 3 hours of operation. The battery ammeter is
located on the lower portion of the overhead panel and in the same gauge as the selectable voltmeter. The chart below
is taken from the Emergency Procedures checklist and provides recommended load management information when
operating on the battery alone. The checklist assumes that only 75% of the rated 36 Amp/hours is available to
compensate for the battery’s state of charge and loss of capacity due to battery aging.

The Battery box is located inside the right wing, between the fuselage and nacelle and directly in front of the main spar.
The battery box is vented through the bottom of the battery box and in the battery box cover, which forms parts of the
surface of the wing. These vents are designed to provide cooling for the battery during normal operations and to allow
the escape of dangerous gases and heat in the event of a thermal runaway (described below). Also located in the battery
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box is the battery relay, which is controlled by the Battery Master switch in the cockpit, and the battery charge rate
monitoring shunt which constantly monitors the current flow (Amps) entering the battery. If the battery charge monitor
detects a current flow that exceeds 7 Amps for 6 seconds will cause the “BATT CHARGE” annunciator to illuminate. The
annunciator will remain illuminated until the current flow decreases below 6 Amps. See Abnormal Procedures Checklist
for details in dealing with a “BATT CHARGE” annunciator.

Thermal Runaway
A thermal runway, which is also known as an overcharge runaway results in the self-destruction of the battery.
The primary factors involved in a thermal runaway are heat, resistance and current flow.

If the battery is exposed to frequent battery only starts with constantly high electrical loads during the flight, the
battery is constantly being discharged then charged over and over and causing excessively high battery
temperatures. Combine this high temperature with as little as 1 bad plate in a cell, and that cell will begin to
overheat. This in turn, causes the adjacent cells to overheat, raising the battery’s overall internal temperature
even more. Shortly, the material between the plates in the bad cell will deteriorate, which lowers the resistance
of the cell. The lower the cell’s resistance, the more current the bad cell will absorb and the hotter the cell, and
in turn, the battery will get.

If the problem is identified at this point, turning off the battery master will isolated the battery (more
importantly, the bad cells) from all charging sources and the cycle should be interrupted.

However, if the problem in not identified and the cycle is allowed to continue; eventually the resistance of the
overheating cells decreases sufficiently that the good cells in the battery start to supply current to the bad cells,
and battery is now internally self-destructing. At this point, turning off the battery master switch will have no
effect and the only thing the crew can do is land the aircraft as soon as possible and have the ARFF services
(Aircraft Rescue and Firefighting) respond to deal with the battery. The longer the battery is allowed to
thermally runaway the greater the chance of the battery catching on fire (Note: the right auxiliary fuel tank is
located directly behind the battery box) and/or overheat to such an extreme that the battery could melt straight
through the bottom of the wing.

External Power
The aircraft is equipped with an external power (ground power) receptacle that can be used to supply all the various
electrical systems and for assisting the battery during engine starts. The receptacle accepts a standard “AN” type plug
and depending on the aircraft serial number, is located in either the bottom aft section of left nacelle (under the inboard
flap), or on the left side of the nose (near the ground communication jacks). The switch for connecting external power to
the aircraft’s electrical system is located on the top-left side of the left subpanel (2 Positions – ON / OFF). Prior to
connecting a Ground Power Unit (GPU), the crew must ensure that the GPU meets the following requirements:

- Voltage: 28 ± 0.5
To protect the aircraft’s electrical system, circuitry will not allow the connection between the GPU and
aircraft’s electrical system in the case of overvoltage (> 32 ± 0.5 VDC) or reverse polarity.
- Power Rating: 300 Amps continuous & 1000 Amps transient (For up to 1 second)

When an external power plug is connected to the aircraft, it completes a circuit and illuminates the white “EXTERNAL
PWR” annunciator. Note: This annunciator only indicates the GPU plug is connected, not if power is available; as the
circuitry that triggers this annunciator is powered by the Hot Battery Bus.

To connect the GPU to the aircraft electrical system, the procedure is:

1) Select BAT on the Voltmeter and verify the battery is supplying a minimum of 20 VDC
Note: If the battery voltage is below 20V do NOT use the external power to charge the battery.
Maintenance must either remove the battery for proper recharging or be replaced

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 2) Check that the “EXTERNAL PWR” annunciator is illuminated
3) Select EXT PWR on the Voltmeter and verify that the GPU is on and supplying the correct voltage
4) Turn ON the Battery Master
5) Turn ON the EXT PWR switch
6) Observe that the “L GEN TIE“ and “R GEN TIE” annunciators extinguish and check that the battery ammeter
does not show a discharge

Note: It is important to always turn ON the battery master switch prior to turning ON the external power switch
as power is used to remove power from the avionics buses. The avionics buses must be unpowered prior to
introducing an external power source to prevent damage to the individual components. Also, the battery will
protect the aircraft electrical systems by absorbing any transient power surges and by providing a backup power
source during engine starts.

Starter-Generators
The aircraft is equipped with 2 Starter-Generators, one per engine, and are located on the accessory gearbox. The
starter-generators are dual purpose; a starter winding is used for starter operation and a shunt field is used during
generator operation.

Starter Operations
Each starter is controlled by a 3 position switch labelled IGNITION AND ENGINE START, located on the bottom-
left side of the left subpanel. Each switch controls its respective starter other interconnected systems. The
positions are:

ON –In this position, the following occurs: (Gated position)

1) Starter relay is CLOSED and the starter is engaged

2) Ignitors energized
3) Standby Fuel Pump energized
4) All 3 Bus Tie sensors are de-sensitized (Left Bus Tie, Right Bus Tie & Battery Bus Tie)

OFF – Starter relay is OPEN and the starter is unpowered (Gated to this position)

STARTER ONLY – In this position, the following occurs: (Momentary position)

1) Starter relay is CLOSED and the starter is engaged

2) All 3 Bus Tie sensors are de-sensitized (Left Bus Tie, Right Bus Tie & Battery Bus Tie)

When the starter relay is closed, the starter is engaged and through the accessory gearbox, will start to rotate
the engine’s compressor section.

During the start sequence, the Compressor Section RPM (N1) must reach a minimum of 12% prior to the addition
of fuel. Battery only starts will typically have enough power to reach 13.5-14.5% N1 (based on a charged battery
in good condition), while external power assisted starts, typically will reach 18-21% N1 (based upon the
performance of the GPU)
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 Ignition and Engine Start

Starter Limitations
Prior to starting an engine the battery must be checked for the following minimum battery voltages, so that in
the event of a fire, the fire bottle explosive squib that releases the extinguishing agent will have sufficient power
to be activated.

Battery Only Start - 23 VDC

External Power Assisted Start - 20 VDC

Once engine ignition (light-off) occurs, there is no limit for the time the starter may be engaged for that start.
However, to prevent damage to the starters, the following cycle limitations must be observed for aborted starts:

20 Seconds ON – 30 Seconds OFF (Aborted Start)

20 Seconds ON – 60 Seconds OFF (Engine Clearing or 2nd attempted start)

20 Seconds ON – 5 Minutes OFF (2nd attempted Start or 3rd attempted start)

Note: A clearing start is only required when fuel has been introduced, but no ignition has occurred. The
clearing start necessary to purge any internally trapped fuel and fuel vapour.

Note: For continuous motoring without attempting an engine start, the starter is limited to 20 Seconds
ON – 5 Minutes OFF.

Note: For restarts, following a normal engine start, a 3 minute cool down is required (time starts from
when the starter is turned off).

Generator Operation
Each generator is capable of supplying 30 VDC (regulated lower) and 300 Amps and can be monitored by the 2
generator load meters, located on lower-center portion of the overhead panel and display generator load in
percent of maximum. When a generator’s respective line contactor is open, the respective amber “L DC GEN” or
“R DC GEN” annunciator will illuminate.

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 The maximum sustainable load that a generator may provide is limited by the N1 (Gas Generator) speed as
follows:

Allowable Generator Load % Minimum Gas Generator RPM (% N1)

0 to 75 65

0 to 100 72

Each generator is controlled by a 3 position switch located on the left side of the left subpanel, under the gang
bar. The positions are:

RESET – To bring a generator online, momentarily hold the switch to this position. This allows the
generator to build up voltage and when the switch is released, and as long as all the GCU monitoring
parameters are met, the GCU will close the line contactor and connect the generator to the aircraft
system. (Momentary position)

NORM – The GCU controls its respective generator. (Spring loaded to this position)

OFF – The GCU keeps line contactor open. (Toggle position)

Note: If the starter is engaged on a generator that is online, the GCU will disable the generator to allow the
starter to operate.

L/R Gen

Generator Control Unit (GCU)

Each generator is controlled by its own GCU and are both located in the cabin floorboards. The GCU provides 2
normal functions; voltage regulation and Load Sharing, as well as 4 protection functions: overvoltage protection,
over-excitation protection, reverse current protection and cross-start current limiting.

Normal Functions

Voltage Regulation
The GCU regulates the voltage supplied by the generator down to 28.25 ± 0.25 VDC. To regulate the
voltage, the GCU first uses residual voltage (0.3 to 2.0 VDC DC that is present whenever the generator is
spinning) to power the generator shunt field. The GCU then varies the voltage in the shunt field to
maintain a constant voltage supply to the electrical system under varying conditions such as: generator
speed, electrical load and temperature.

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 Load Sharing (Paralleling)
The GCU load sharing circuitry monitors the output of each generator and maintains the electrical loads
within 10% of each other for their entire operating range. Once both generators are brought online, the
GCU compares the load of each generator and utilizes the voltage regulation function to make small
voltage adjustments which results in an equally shared load.

Protection Functions

Overvoltage Protection
The GCU monitors each generators output voltage and should it exceed 32.5 VDC will open the
respective line contactor (described below). Also, a separate circuit will open the respective line
contactor, should the generator’s output voltage exceed 40 VDC.

If a failure causes a generator’s voltage to increase above the normally regulated amount, that
generator will attempt to assume the entire electrical load and force current flow into the opposite
(properly regulated) generator. This will trigger the properly regulated generator’s reverse current
protection and its GCU will open the line contactor and isolate it from the system. If the bad generator
voltage continues to increase above 32.5 Volts, its GCU will open the line contactor and isolate it from
the system; while at the same time, the properly regulate generator will be automatically reconnected.

Over-excitation Protection
The GCU monitors each generator’s output voltage and current and if it should exceed 28.5 VDC and the
current differential between generators is > 15% for 5 seconds, the GCU will open the respective line
contactor and isolate the generator. Note: This protection only functions with both generators online
and does not require overvoltage to function.

Reverse Current Protection

The GCU monitors the generators for positive current flow (flow out of each generator). Should a
generator‘s voltage output drop or the generator has slowed down enough that it can no longer
maintain positive current flow (like during engine shutdown), it will begin drawing current in from the
rest of the aircraft’s electrical system. At this point, the GCU will open the respective line contactor and
isolate that generator from the system. Note: The generator does not need to be reset when reverse
current protection has been triggered. The GCU will automatically correct itself should the situation be
corrected.

Cross-Start Current Limiting

The GCU limits the current that the online generator can output during engine cross-starts. This
prevents the online generator from providing excess current to the starter on the other engine. This
protection is activated on the opposite generator, when either ignition and engine start switch is placed
in the ON or the STARTER ONLY position.

Bus Tie System

The aircraft’s electrical system utilizes bus tie’s that are protected from excessively high current flow by 3 current
sensors attached to their respective bus tie relays and are also known as Hall Effect Devices (HED’s). The 3 HEDs
automatically open their associated bus ties if the preset current level is exceeded after a time delay. The left and right
generator bus tie and HED are located between their respective left/right generator bus and the center buses and
illuminates the “L GEN TIE OPEN” and/or “R GEN TIE OPEN” whenever the associated bus tie is open. Both generator
bus HEDs are set to open their bus tie when the current flow exceeds 325 Amps. The battery bus tie is located between
the battery and the right center bus and illuminates the “BATT TIE OPEN” whenever the battery bus tie is open and is set
to open the battery bus tie when the current flow exceeds 275 Amps. All the ties have a built in time delay to allow for
transient loads. The generator bus ties reaction time of 0.01 seconds and the battery tie’s reaction time is 0.12 seconds,
which is longer than the other bus ties to avoid all 3 bus ties ever opening all at the same time during transient loads.
The Battery Bus Tie will automatically close when it senses a minimum of 20 VDC and each Generator Bus Tie will

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automatically close when the respective tie senses a minimum of 28 VDC (If the GEN TIES switch is in the NORM
position).

Manual control of all of the bus ties is provided by a 3 position switch located on the left subpanel. The positions are:

RESET - Once an HED has opened a bus tie, that bus tie will remain open until the circuit is reset. To reset the
HED and allow the bus ties to close, momentarily press the BUS SENSE switch to RESET (Momentary position)

NORM – Allows HEDs to function normally. (Spring loaded to this position)

TEST – To test the bus tie system, start with the aircraft engine running and both generators online.
Momentarily press the BUS SENSE switch to TEST, this will apply voltage directly to the HEDs and simulate and
excessive current situation, which will open the 3 bus ties and illuminate all 3 of the bus tie annunciators: “L GEN
TIE OPEN”, “R GEN TIE OPEN” and “BATT TIE OPEN”. (Momentary position)

Note: Voltage is continuously applied to the HEDs for as long as the switch is held in the TEST position.
The longest the switch should need to be held is approx 0.12 seconds (longest reaction time for the
battery bus) and prolonged application of the test voltage will damage the HED. The switch is spring
loaded to the center off position.

Bus Sense Reset

The GEN TIES switch, which is located on the right side of the bus sense switch, is used to control the generator bus tie
relays. The positions are:

MAN CLOSE – The generator bus tie relays are manually closed. This allows the battery to supply power to all of
the aircraft’s electrical system. (Momentary position)

Note: With the battery being the only source of power, it will discharge quickly and should only be used
for the minimum amount of time required.

NORM – The generator bus ties function normally. When operating with only the battery master on, both
generator bus ties will remain open as part of the automatic load shedding function (see abnormal indications
below) and will automatically close once either generator is brought online or external power is connected to
the aircraft’s electrical system and the voltage is at or above 28 VDC. (Spring loaded to this gated position)

OPEN –Both generator bus ties are forced to stay open and isolate the generator buses from the center buses.
(Gated position)

Note: If the generator bus tie was opened by its respective HED, using the GEN TIES switch will have no effect on the bus
tie position and the BUS SENSE switch must be used as described above.

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 Gen Ties Man Close

Circuit Breakers & Fuses

The aircraft is equipped with circuit breakers and fuses to protect the electrical system. Both systems are used to detect
excessive current flow and to isolate the faulty system from the rest of the electrical system. Fuses are one time use only
and once blown, must be removed and replaced. Circuit Breakers come in two types:

Regular Circuit Breakers

Regular circuit breakers are stamped with their rated

amperage in the center and when tripped, slightly pop out
of the panel slightly to display a white coloured shaft to
indicate its activation.

Switch Type Circuit Breakers Switch type circuit breakers display their rated amperage
on the end of the switch and when tripped, will return to
its OFF position. No other indication of its activation is
available.

Once a circuit breaker has tripped, it should not be reset, except the following cases; and in any case, if the circuit
breaker trips again after being reset, it must not be reset again.

1) The flight crew has determined that the system is required for the safe continuation of the flight and careful
consideration should be given to the timing and duration of the reset.

2) The aircraft checklist and/or operating manual directs the reset a circuit breaker as part of a procedure.

Breaker Panels
The aircraft is equipped with many circuit breaker panels; however, many are under the cabin floorboards,
engine nacelles, and avionics bay that are not accessible by the flight crew and will not be covered in this
manual. The 2 panels that are accessible to the crew contain the circuit breakers for most of the important
systems. The fuel system circuit breakers are located on the left side of the cockpit, integrated into the fuel
management panel. While the main circuit breaker panel is located on the cockpit’s right sidewall.

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AC Power Distribution
Some of the aircraft avionics and engine instruments operate on AC power (see AC Bus Electrical Items chart below). The
dual output inverters convert the nominal 28.25 VDC to both 115VAC and 26VAC with an operating at a frequency of
400 Hz. The DC power is supplied to the inverters from the respective center bus or generator bus. When the respective
generator bus is powered the inverter will be powered from that bus. However, if the generator bus should become
unpowered, the inverter will automatically switch to being powered from the respective center bus as shown below.

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The inverters can be monitored on the voltage/frequency meter located on the lower portion of the overhead panel.
The meter only displays one piece of data at a time. The switch on the left side of the meter is used to select which
inverter to monitor. By default the meter will displays the frequency of the selected inverter and to display the voltage,
the button on the bottom left of the meter must be held down. The range displayed on the meter is also the operational
limits of the inverter.

Inverter

Each inverter is controlled by a 3 position switch, which is located on the top-left of the left subpanel. The positions are:

ON – Connects DC power to the inverter. Inverter powers AC buses. (Gated position)

OFF – Removes DC power from the inverter. Both AC buses are unpowered. (Gated position)

TXFR – Removes DC power from the inverter and connects the opposite inverter to both of the AC buses on the
side that was selected to TXFR. (Gated position)

Left AC Bus Electrical Items Right AC Bus Electrical Items

115 VAC Bus 115 VAC Bus
AHRS #1 AHRS #2
RADAR
Flight Recorder 26 VAC Bus
RH Torque Pressure Indicator
26 VAC Bus (Only UE-92 and lower)
LH Torque Pressure Indicator RMI #1
(Only UE-92 and lower) RMI #2 VOR Pointer
RMI #2
RMI #1 VOR Pointer
Flight Recorder

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DC Bus System – Main Items
L GENERATOR BUS R GENERATOR BUS L CENTER BUS:
***FLAP MOTOR PITCH TRIM (OPT) ***LANDING GEAR MOTOR
***FLAP CONTROL & INDICATOR PROP SYNCHROPHASER R FUEL BOOST PUMP
NOSE WHEEL STEERING CONTROL (OPT) SUBPANEL, OVHD & INSTR LIGHTS NO. 1 INVERTER POWER
***AUTOFEATHER CABIN INDIRECT (AISLE) LIGHTS – FULL L MANUAL PROP DEICE
FLIGHT INSTRUMENT LIGHTS READING LIGHTS
ENGINE & AVIONICS INSTR LIGHTS EDGE LIGHT PANELS R CENTER BUS:
***R BLEED AIR CONTROL RIGHT GEN BUS TIE POWER CONDENSER BLOWER MOTOR
VENT BLOWER CONTROL L ENG ANTI-ICE CONTROL (STANDBY) ***PNEUMATIC SURFACE DEICE
L FUEL VENT HEAT STALL WARNING HEAT MANUAL PROP DEICE CONTROL
BRAKE DEICE (OPT) R ENG ANTI-ICE CONTROL (STANDBY) R MANUAL PROP DEICE
L ENG ANTI-ICE CONTROL (MAIN) R FUEL VENT HEAT WINDSHIELD WIPER MOTOR
R ENG ANTI-ICE CONTROL (MAIN) AIR CONDITIONER CLUTCH GENERATOR RESET
L GEN BUS TIE POWER ALTERNATE STATIC HEAT NAV LIGHTS (PRI)
L LANDING LIGHT R LANDING LIGHTS ICE LIGHTS
TAIL FLOOD LIGHT (OPT) RECOGNITION LIGHTS TAXI LIGHTS
L ENGINE AUTOMATIC PROP DEICE R ENGINE AUTOMATIC PROP DEICE L FUEL BOOST PUMP
R FIREWALL VALVE R WINDSHIELD ANTI-ICE NO. 2 INVERTER POWER
INVERTER NO. 1 L FIREWALL VALVE GROUND WARNING LIGHT
POWER STEERING PUMP MOTOR (OPT) INVERTER NO. 2 CABIN INDIRECT LIGHTS - PARTIAL
L GEN AVIONICS BUS R GEN AVIONICS BUS L & R FIRE EXTINGUISHER MONITOR
FWD VENT BLOWER AFT VENT BLOWER
ANTI-SKID PUMP MOTOR (OPT) PROP GOV TEST HOT BATTERY BUS AND BATTERY BUS:
L WINDSHIELD ANTI-ICE R AUX FUEL TRANSFER PUMP STEREO (OPT) DOOR ENTRY LIGHTS
FURNISHING CONTROL (OPT) MOD/FURNISHINGS POWER RNAV MEMORY (OPT)
L AUX FUEL TRANSFER PUMP EMERGENCY EXIT LIGHTS CARGO COMPARTMENT LIGHTS
ANTI-SKID (OPT) ANTI-COLLISION LIGHT - STROBE (OPT) ***L ENG FIRE EXTINGUISHER
ANTI-COLLISION LIGHT - FLASHING PA & COPILOT PHONE ***R ENG FIRE EXTINGUISHER
PILOT PHONE SECONDARY NAVIGATION LIGHT COCKPIT EMER LIGHTS
GPWS (IF INSTALLED) ***COPILOT’S EFIS ***L F.W. SHUTOFF VALVE
TCAS (IF INSTALLED) COPILOT’S ENCODING ALTIMETER R PITOT HEAT (BATTERY BUS)
***PILOT’S EFIS PROP GND SOL * ***R F.W. SHUTOFF VALVE
***STANDBY ATTITUDE IND. POWER CABIN LOADING LIGHTS
* Airplanes Modified by Kit No. 129-9011-1. EXTERNAL POWER OVER-VOLTAGE SENSOR &
ADVISORY LIGHT
TRIPLE FED BUS:
L FUEL FLOW INDICATOR R FUEL FLOW INDICATOR L OIL TEMP INDICATOR
R OIL TEMP INDICATOR L OIL PRESS INDICATOR R OIL PRESS INDICATOR
L FIRE DETECTION R FIRE DETECTION L IGNITION POWER
R IGNITION POWER L START CONTROL R START CONTROL
L BLEED AIR WARNING R BLEED AIR WARNING STALL WARNING
AURAL WARNING ***ANNUNCIATOR POWER ***ANNUNCIATOR INDICATOR
LANDING GEAR WARNING HORN SILENCE ***LANDING GEAR POSITION CONTROL CABIN TEMP CONTROL
L BLEED AIR CONTROL BATTERY BUS TIE POWER CURRENT SENSE TEST
AVIONICS MASTER CONTROL CABIN PRESS CONTROL AUTO PROP DEICE CONTROL
*PILOT’S ENCODING ALTIMETER AVIONICS ANNUNCIATOR OVERSPEED SENSOR
*** TRIPLE FED AVIONICS BUS (NAV1/COM1) L PITOT HEAT LANDING GEAR CONTROL
FUEL TRANSFER VALVE L FUEL QUANTITY R FUEL QUANTITY
L FUEL PRESSURE WARNING R FUEL PRESSURE WARNING L FUEL LOW LEVEL WARNING
R FUEL LOW LEVEL WARNING L ITT INDICATOR (UE-93 & After) R ITT INDICATOR (UE-93 & After)
L TORQUEMETER (UE-93 & After) R TORQUEMETER (UE-93 & After) L PROPELLER TACHOMETER
R PROPELLER TACHOMETER L TURBINE TACHOMETER R TURBINE TACHOMETER
RUDDER BOOST INSTRUMENT INDIRECT LIGHTS PILOT SPEAKER
COPILOT SPEAKER L GEN BUS TIE POWER & CONTROL COCKPIT VOICE RECORDER
**LAVATORY SMOKE DETECTOR ***ENGINE INSTRUMENTS
*On Airplanes Equipped With a Flight Director, But No Autopilot.
** Airplanes Modified by Kit No. 129-5031-1.
*** THESE ARE THE MAIN ITEMS ASSOCIATED WITH EACH DC BUS AND SHALL BE COMMITTED TO MEMORY

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Electric Malfunctions and Faults
“L DC GEN” or “R DC GEN”
Abnormal Procedures – “Generator Inoperative Checklist”

Generator line contactor is open

1) Generator switch is OFF

2) The GCU has detected a fault and opened the line contactor

“L GEN TIE OPEN” or “R GEN TIE OPEN”

Abnormal Procedures – Generator Tie Open Checklist

1) Associated loadmeter is less than 100% - Possible transient excess current caused HED to open bus Tie. Bus
Sense Reset may be attempted.
2) Associated loadmeter is greater than 100% - Generator bus has a short, respective generator must be turned
OFF to remove power from being supplied to the short. Affected generator bus electrical systems are no
longer available.

“L GEN TIE OPEN” with “L DC GEN” or “R GEN TIE OPEN” with “R DC GEN”
Abnormal Procedures – “Generator Inoperative Checklist and Generator Tie Open Checklist”

No specific checklist for this situation. First perform the Generator Inoperative checklist so that if a short exists
on the associated generator bus, the fault is isolated to that bus and will not affect the rest of the system.

1) If the checklist results in the loadmeter indicating less than 100% (regardless if the generator is online or
offline), continue with the Generator Tie Open checklist.
2) If when performing the Generator Inoperative checklist, the generator loadmeter indicates greater than
100%; the generator bus has a short and the respective generator must be turned OFF to remove power
from being supplied to the short. Affected generator bus electrical systems no longer available.

“L GEN TIE OPEN” and “BATT TIE OPEN” and “R GEN TIE OPEN”
Abnormal Procedures – “Both Generator Ties and Battery Tie Open Checklist”

Center Bus Short – Center buses has been isolated from the aircraft’s electrical system. Both center bus
electrical items are no longer available.

“L DC GEN” and “L GEN TIE OPEN” and “R GEN TIE OPEN” and “R DC GEN” and “R ENIVRO FAIL”
Emergency Procedures – “Dual Generator Failure Checklist”

Dual generator failure – Both generator line contactors are open and as the voltage has dropped below 28 VDC,
both generator bus ties have opened (load shedding). The R ENVIRO FAIL has illuminated because the valve is
powered from the left generator bus, which is now unpowered. The left environmental valve is powered from
the center buses so it will remain powered from the battery. (See Emergency Procedures Checklist)

Note: To aid the crew in identifying which systems will remain functional after load shedding occurs, the systems
that are powered from either the center buses or the triple fed bus are identified by having a white circle around
the switches that control them.

“L PITOT HEAT”
Abnormal Procedures – “Pitot-Static Mast Heat Failure”

1) Left pitot heat switch is OFF

2) Insufficient current being supplied to the left pitot heat to prevent icing
3) Batter Master switch is off. Left pitot heat is powered by the battery bus.

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“L AC BUS” or “R AC BUS”
First illumination of annunciator: Emergency Procedures – “Single Inverter or AC Bus Failure”

Annunciator indicates a failure of either the AC inverter or one of the associated AC Buses. If after completing
the first portion of the checklist, if the annunciator does not extinguish the crew must decide if the associated
loss electrical items are required for safe completion of the flight. At this point the crew still does not know
whether the failure was caused by a failed inverter, or a short on one of the AC Buses.

If the rest of the checklist procedure is followed and the fault was a failed inverter, full functionality of all
electrical items will be restored. However, if the fault was a short on one of the AC Buses then the checklist
procedure has a remote possibility of causing the operative AC inverter to fail which results in all AC powered
electrical items will no longer be available (See AC Power Distribution for AC Bus Electrical Items).

If the decision is made not to continue with the remainder of the checklist procedure, the crew member on
same side of the failed AC Inverter/Bus is left without useable data being displayed on their EFIS screens.
However, if one ignores the fact that it is the AC Inverter/Bus failure and just reacts to the red ATT flag on the
EADI and the red HDG flag on the EHSI, the fault could be identified as a failure of an AHRS (Attitude and
Heading Reference System. If the abnormal procedures – Failure of an Attitude or Heading Reference System
checklist is used, both crew members will once again have fully functioning EFIS screens and allow for proper
application of the pilot flying and pilot monitoring roles.

“L AC BUS” and “R AC BUS”

If both annunciators illuminate with no previous failures - Emergency Procedures – “Dual Inverter Failure”

With both annunciators illuminate simultaneously, the left and right AC inverters are offline, resulting is
the loss of all AC powered items. When following the checklist procedure, if one AC inverter successfully
reset, all systems are transferred to the operative inverter. If neither will reset, than all AC powered
electrical items will no longer be available (See AC Power Distribution for AC Bus Electrical Items)

If both annunciators illuminate after transferring an AC bus during the “Single Inverter Or AC Bus Failure” –
“Single Inverter Failure Followed By A Second Inverter Or AC Bus failure”

If both annunciators illuminate simultaneously, both AC inverters are offline, resulting is the loss of all
AC powered items. When following the checklist procedure, the only previously operative inverter is
reset in an attempt to regain partial power to some of the AC bus items. If the inverter will not reset,
than all AC powered electrical items will no longer be available (See AC Power Distribution for AC Bus
Electrical Items)

Loss of all engine instruments, Comm #1, NAV #1 and RMI #1 without any annunciators
No Procedures or Checklist available

Triple Fed Bus short – All items on the triple fed bus are no longer available, including the items on avionics bus
#1. There are no actions the crew can take to resolve the problem and can only be corrected by maintenance.

System Checks
Electrical System Test
The electrical system is tested to ensue normal and abnormal functions function correctly. With both engines
operating and the electrical system operating normally, select the R Generator Switch to the OFF position and
verify that the amber “R DC GEN” annunciator illuminates to confirm that the warning system is operating
correctly, and check the voltmeter to confirm that the L Generator is adequately supplying the system. Select
the R Generator Switch to RESET then ON, and verify that the respective annunciator extinguishes.

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 Note: Only the L Generator (turning off the R Generator) is tested as the R Generator was tested during
the generator cross-start of the #1 engine.

Momentarily select the Bus Sense Switch to the TEST position and verify that the amber “L GEN TIE OPEN”,
“BATT TIE OPEN” and “R GEN TIE OPEN” annunciators illuminate and check the center bus on the voltmeter to
confirm that the 3 bus ties have opened and that the center bus is NOT powered. Momentarily select the Bus
Sense Switch to the RESET position and verify that the respective annunciators extinguish and check for proper
voltage on the center bus.

Note: Voltage is continuously applied to the HEDs for as long as the switch is held in the TEST position.
The longest the switch should need to be held is approx 0.12 seconds (longest reaction time for the
battery bus) and prolonged application of the test voltage will damage the HED. The switch is spring
loaded to the center off position.

Select the Gen Tie Switch to the OPEN position and verify that the amber “L GEN TIE OPEN” and “R GEN TIE
OPEN” annunciators illuminate and check the triple fed bus on the voltmeter to confirm that the both the Gen
Ties have opened and the voltage on the triple fed bus is maintained by the generators. Return the Gen Tie
Switch to the NORM position and verify that the respective annunciators extinguish and check that the
loadmeters are sharing the electrical load.

SOP – Electrical System Test

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FUEL SYSTEM

General Description
The Beechcraft 1900D fuel system consists of an independent fuel system for each engine. Each system consists of 2 fuel
storage tanks, a collector tank, high and low pressure engine driven fuel pumps, standby electric fuel pump, jet transfer
pumps and an auxiliary electric fuel pump. The two systems are connected by a single cross-transfer line that allows fuel
to be transferred from one system to the other.

Fuel Tank System

Fuel Storage
Fuel is stored in 2 main and 2 auxiliary wet-wing tanks. Each main tank is located in the wing, outboard of the
nacelle and is filled through a single port near the wingtip. Each auxiliary tank is located in the wing root,
inboard of the nacelle and is filled through a single port beside the nacelle. As wing structure forms the fuel
tank, the structural ribs inside the tank also function as baffles to prevent the fuel from sloshing. The filling ports
are equipped with an anti-siphon valve that prevents fuel from exiting the tank in the event the filler cap should
become dislodged.

Note: The main tanks must be filled to capacity prior to adding any fuel to the auxiliary tanks. Also, the auxiliary
fuel tanks must be used first during normal operations.

Usable Fuel Pounds Litres Gallons

Total Fuel 4484 2518.8 665.4
Each Main Fuel Tank 1621 910.4 240.5
Each Auxiliary Fuel Tank 621 349.0 92.2

Fuel Tank Venting

The vent system for the main wing tanks is located in each wing tip. Air enters the system through a recessed air
scoop that is designed to be ice resistant, a protruding ram vent that is heated to prevent ice formation and a
flush mount vent that is equipped with a flame arrester. The flame arrester allows air to enter the vent system
while preventing flame from propagating from an external source, such as a lighting strike or static discharge,
from entering the fuel tank.

Inside the main tank, the vent system connects to two vents. The primary vent is equipped with a float valve
that allows air to enter and maintain slight positive pressure in the tank, while at the same time, designed to
shut off if fuel is present at the float, to prevent fuel from exiting the vent. The backup vent is a pressure
activated valve that allows the fuel tank to vent out to the atmosphere should the pressure inside the tank
exceed 1.5 PSI.

The auxiliary fuel tanks are vented through the main tank system. A tube connects the auxiliary tank with its
respective main tank which maintains equal pressure. This tube is always open and contains no valves which can
result in fuel flowing from the main tank to its respective auxiliary tank under certain conditions.

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 Condition 1 – Operations at low power settings, with full fuel in the main tank and the auxiliary fuel
pump is on. The auxiliary fuel pump transfers fuel into the collector tank and is then delivered to the
engine. However, during operations at low power settings, the demand for fuel is less than the amount
being supplied and the excess fuel is returned into the main tank. As the main tank is already full, the
excess fuel will return to the auxiliary fuel tank through the vent tube. This cycle will continue until the
engine’s fuel demand increases, the auxiliary tank in empty or the auxiliary pump is turned off.

Condition 2 – If the main fuel tank is full or close to full, thermal expansion of fuel in the main tank may
cause fuel to flow into the auxiliary tank.

Note: If the main fuel tank and respective auxiliary fuel tank are both full, fuel may vent out of the
backup vent in the event of thermal expansion.

Fuel Drains
During pre-flight inspections, the fuel tank sumps, pumps and filters should be drained to check for fuel
contamination. The locations of the drains are listed below:

Drain Location
Main Tank (2) Bottom of wing, outboard of nacelle
Aux Tank (1) Bottom of wing, Inboard of nacelle
Collector Tank (1) Outboard side of nacelle
Fuel Filter (1) Through access door, bottom of the
wing, outboard of nacelle

Note: Water does not settle out of jet fuel as easily as avgas. To properly drain water contamination
from the system, the aircraft must sit undisturbed for several hours.

Fuel System Operation

The fuel system is designed to be almost fully automatic. Crew inputs are only required during use of the auxiliary tank,
out of balance conditions and/or abnormal operations.

Main Tank Operation

Fuel is added to the main tanks using the filling port near the wing tip. When filling the main tanks completely,
the tank must be “topped up” several times as the fuel takes a little bit of time to flow through ribs and the air
must travel back up to the filling port.

Auxiliary Tank Operation

Fuel may be added to the auxiliary tanks only after the main tanks have been filled to capacity. Fuel is added to
the auxiliary tanks using the filling port just inboard of the nacelles. To use fuel in the auxiliary tanks, it must first
be pumped into the collector tank. Each auxiliary pump is controlled by a 3 position switch, located outside of its
respective fuel quantity gauge and when needed is normally operated in the AUTO position and a float switch
attached to the pump. Also, the fuel transfer line between the auxiliary fuel pump and collector tank is

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 monitored by a pressure switch that will trigger its respective “L NO AUX TFR” and/or “R NO AUX TFR”
annunciator when the pressure drops below 5 PSI, and the auxiliary pump switch is in the ON or AUTO positions.

The auxiliary pump has the following modes of operation.

ON – Auxiliary pump is ON, regardless of the position of the float switch. (Gated position)

Note: Once the tank is empty, the respective “L NO AUX TFR” or “R NO AUX TFR” annunciator
will illuminate to alert the crew that the tank is empty. However, the pump will continue to
operate and must be turned OFF manually. As the pump is cooled by the fuel it pumps, if the
pump continues to operate it will overheat.

AUTO – Auxiliary pump float controls the pump. If the float detects the presence of fuel, the auxiliary
pump will turn ON and once the tank is empty, the pump will turn OFF. (Gated position)

OFF – Auxiliary pump is OFF (Gated Position)

Aux Pump

As a safety precaution, the auxiliary pump contains a thermal protection circuit. Should the pump
temperature increase beyond a preset limit, power to the pump is interrupted and only once it cools
sufficiently will the power be reconnected.

Collector Tank Operation

The collector tanks are the source of fuel for each engine. Fuel from each main tank is transferred into its
respective collector tank through 3 sources:

Tank cut-outs – The dividing stringer between the main tank and collector tank has 3 cut-outs in the
upper portion of the stringer, which allows fuel to gravity feed into the collector tank.

Flapper Valves – The dividing stringer between the main tank and collector tank has 3 flapper valves in
the lower portion of the stringer, which allows fuel to gravity feed into the collector tank, but not exit
back into the main tank

Jet Transfer Pumps – The main tank contains 2 jet transfer pumps (fwd and aft) which utilize the motive
flow of fuel from the high pressure and low pressure engine driven fuel pumps and when used, the
standby fuel pump

Once the fuel is in the collector tank, the engine driven low pressure pump provides fuel (at a raised fuel
pressure) for the engine’s high pressure fuel pump and motive flow for the 2 jet transfer pumps in the main tank
and the main jet transfer pump in the collector tank.

Fuel Heater
The oil-to-fuel heat exchanger uses the heat from the engine’s oil system to pre-heat the fuel and prevent icing
in the other fuel system components. A temperature sensing bypass valve constantly adjusts the flow of oil
through the heater core to maintain a fuel temperature of 21-32°C (70-90°F).

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 Fuel Pumps

Jet Transfer Pumps

The jet transfer pumps are built using a venturi and operate with no moving parts. Motive flow provided
by the engine driven low pressure fuel pump or standby electric pump is used as the motive flow for the
operation of the jet pumps. As the motive flow fuel under pressure passes through the jet pump, the
venturi causes the motive flow fuel to increase in velocity and creates a vacuum inside the chamber.
This draws fuel from the vacuum line and combines it with the motive flow fuel and delivers it to the
required location.

Low Pressure Engine Driven Fuel Pump

The engine driven low pressure fuel pump powered from the engine’s accessory section and therefore
operates whenever the engine is operating. The pump provides fuel at approximately 45 PSI for the
engine’s high pressure fuel pump and motive flow for the jet pumps. Should a low pressure engine
driven fuel pump fail, the loss of motive flow pressure (10 PSI switch) will trigger the respective “L FUEL
PRESS LOW” or “R FUEL PRESS LOW” annunciator. The engine will continue to operate, however, if the
annunciator remains illuminated, the engine is limited to 10 hours of operation before overhaul or
replacement of the engine driven high pressure fuel pump is required.

High Pressure Engine Driven Fuel Pump

The engine driven high pressure fuel pump raises the fuel pressure to approximately 800-850 PSI. There
is no backup system for the high pressure fuel pump and its failure, will result in an engine failure.

Standby Electric Fuel Pump

The standby electric fuel pump is used for engine starts, cross-transferring fuel and backup for engine
driven fuel pump and raises the fuel pressure to greater than 11 +/- 2 PSI. The only time that the
STANDBY PUMP switch used, is after the failure of the engine driven low pressure fuel pump and is
located on the top corners of the fuel control panel.

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 Fuel Cross-Transfer Operation
Normally each fuel system (Left and Right) operate independently. However, in the event of an imbalance
between fuel systems, a manually controlled cross-transfer line connects the two systems. This line contains a
normally closed solenoid valve. With the valve closed, the fuel system draws fuel from their respective tanks. To
transfer fuel, the gated TRANSFER FLOW switch in placed in the direction that fuel is to be transferred. Once a
transfer is selected, the cross-transfer solenoid valve is opened and the white “FUEL TRANSFER” annunciator will
illuminate to indicate the valve is fully open, as well, the opposite standby electric fuel pump is turned ON. To
stop transferring fuel, the switch is returned to the center position.

LEFT – Cross-transfer solenoid valve is OPEN and the RIGHT standby fuel pump is turned ON

OFF – Cross-transfer solenoid valve is CLOSED

RIGHT - Cross-transfer solenoid valve is OPEN and the LEFT standby fuel pump is turned ON

Fuel Transfer

If a cross-transfer is selected and the cross-transfer solenoid valve does not reach the selected position (Open or
Closed) within 2 seconds, the “XFR VALVE FAIL” annunciator will illuminate. (See Fuel System Malfunctions &
Faults)

Maximum fuel imbalance between fuel systems is 200 lbs and is defined as the Left Main + Left Aux compared to
Right Main + Right Aux.

Note: Never initiate a cross-transfer to a fuel system that is thought to have a fuel leak.

Firewall Fuel Valve

Each engine is equipped with an electric motor controlled firewall fuel valve that is used to isolate the fuel
system from the engine. The firewall fuel valves are operated by its respective T-handle, which are located in the
top center of the main instrument panel. Once a T-handle is pulled out, the associated firewall fuel valve
powered closed and respective fire extinguisher is armed.

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Fuel Indication Systems
Fuel Quantity Gauges
Fuel quantity is displayed on the fuel control panel, located on the left side panel and displays the quantity of
fuel remaining in pounds. Each fuel system has an independent gauge that display the quantity of fuel in the
main tanks or auxiliary tanks. The fuel quantity selection switch, located on the lower right corner of the fuel
control panel, is used to select if the fuel gauges display the fuel quantity in the main tanks or the auxiliary tanks.
The switch is spring loaded to the MAIN position, which displays the main fuel tank quantity, and must be held
down into the AUX position to display the auxiliary fuel tank quantity.

The fuel quantity is derived by a capacitance multi-probe system that compensates for changes in the specific
gravity (fuel temperature and fuel type) of the fuel and allows the quantity to be displayed in pounds on a linear
scale. The main tanks contain 5 fuel probes, the auxiliary tanks contain 2 probes and the collector tank contains
1 probe.

Note: The maximum error of the capacitance system is 3%

Fuel Control Panel

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 Low Fuel Quantity Warning System
In addition to the fuel quantity gauges, the aircraft is equipped with a low fuel quantity warning system. Each
main fuel tank and collector tank contains a fluid level sensor (See Annunciator Section for Description) that
illuminates an annunciator once the fuel level has been reduced to a preset level. The “L FUEL QTY” and “R FUEL
QTY” annunciators illuminate once the fuel in its respective main fuel tank drops below 324 lbs. The “L COL
TANK LOW” and “R COL TANK LOW” annunciators illuminate once the fuel quantity in its respective collector
tank has been reduced below 53 lbs.

Note: During taxiing the “L FUEL QTY” and/or “R FUEL QTY” annunciators may illuminate during sustained turns
if the fuel quantity is sufficiently low.

Liquid Level Indicators

Each main tank is equipped with 2 liquid level indicators located under each wing, which alert the crew during
pre-flight inspections, to the minimum amount of fuel that is in the tank at two preset quantities. The sight
window of the indicator shows black to indicate the presence of fuel and red with a black dot to indicate the
absence of fuel. The outboard indicator on each wing will be coloured black when the fuel in the tank is 1150
lbs or greater and the inboard indicator will be coloured black when the fuel in the tank is 745 lbs or greater.

Fuel Flow
Fuel flow gauge displays the fuel being delivered to each engine, located in the engine instrument cluster and
displays the fuel flow in pounds per hour. The fuel flow gauges receive their data from fuel flow transducers
(sending units) located downstream of the fuel control unit.

Fuel Flow Gauges

Fuel Purging System

The aircraft is equipped with an EPA (Environmental Protection Agency) fuel Purging System. The system ensures that
any residual fuel in the fuel manifolds is consumed during engine shutdown. While an engine is in operation, P3 air (See
Pneumatic Section) is routed through a filter and pressurizes a small air tank. The P3 air cannot exit the tank during
engine operation as a check valve is held closed by fuel pressure on the other side. Once the engine is shutdown, the
fuel pressure in the manifold drops below the pressure held within the air tank. This allows the pressurized air to pass
through the check valve and force all the residual fuel in the manifolds out though the fuel nozzles and into the
combustion chamber. During shutdown this can be observed by a momentary surge in N1 as this residual fuel is burned.

Note: If during engine shutdown, smoke continues to flow from the exhaust system for several minutes, the EPA Purge
system may not be functioning correctly and the residual fuel is slowly dripping from the fuel nozzles into the still hot
combustion chamber.
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Limitations
The approved fuels for the aircraft are JET A, JET A1 and JET B. The aircraft may be operated on AVGAS, however, the
following limitations apply:

- Operation is limited to 150 hours between engine overhauls

- Operation is limited if either standby pump is inoperative
- Operation is prohibited above 18,000 ft
- Standby pumps must be on for takeoff and landing

Takeoffs are prohibited with the fuel quantity in the main tanks is in the yellow arc (less than 363 lbs)

Fuel System Malfunctions and Faults

“L FUEL QTY” or “R FUEL QTY”
Abnormal Procedures – “Fuel Level Low”

Fuel quantity remaining in the respective main tank is at or below 324 lbs. Verify actual fuel quantity remaining
with appropriate fuel gauge.

Note: Approximately 30 minutes of usable fuel remains when this annunciator first illuminates.

“L COL TANK LOW” or “R COL TANK LOW”

Abnormal Procedures – “Collector Tank Low”

Fuel remaining in the respective collector tank is at or below 53 lbs. Verify that the associated main tank is
empty, and that if fuel was in the auxiliary tank, has transferred successfully. Depending on the root cause of the
fuel status, consider transferring fuel to continue engine operation.

Note: Approximately 8 minutes of usable fuel remains when this annunciator first illuminates, if configured for
normal cruise at 160 knots and 10,000’.

“XFR VALVE FAIL”

Abnormal Procedures – “Transfer Valve Failure”

The cross-transfer solenoid valve has not reached the selected position.

1) If annunciator illuminates when initiating fuel transfer – cross-transfer valve is either closed or partially
closed. Fuel cross-transfer may not be available.
2) If annunciator illuminates when terminating fuel transfer – cross-transfer valve is either fully open or
partially open. Constant monitoring of fuel systems is required to maintain proper balance.

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“L FUEL PRESS LO” or “R FUEL PRESS LO”
Emergency Procedures – Fuel Pressure Low

1) Low Pressure Engine Driven fuel Pump Failure – Only the high pressure engine driven fuel pump is operating
and the standby electric fuel pump should be turned ON to compensate for the loss which should extinguish
the annunciator.
2) Accessory Driveshaft failure - The low pressure engine driven fuel pump is mounted on the same accessory
driveshaft as the external savage oil pump and if the driveshaft has failed, the indication would be a
constantly increasing oil temperature. The external scavenge pumps are responsible for the return of oil
from engine bearings #3 and #4 as well as the reduction gearbox.

Note: The engine is limited to 10 hours of operation before overhaul or replacement of the engine driven high
pressure fuel pump is required

Note: Fuel cross-transfer is not possible to the side of the failed engine driven low pressure fuel pump. However,
fuel can still be transferred to opposite side.

“L NO AUX XFR” or “R NO AUX XFR”

Abnormal Procedures – Auxiliary Fuel Transfer failure

1) Aux pump switch in the AUTO position

a. If the tank is empty – The automatic float shutoff switch has failed to turn off the auxiliary pump and
the fuel line pressure has dropped below 10 PSI as the tank is empty.
b. If fuel remains in tank – Possible system failure in AUTO mode, attempt operation is ON mode
2) Aux pump switch in the ON position
a. If the tank is empty – The fuel line pressure has dropped below 10 PSI as the tank is empty.
b. If fuel remains in tank – System has failed and fuel remaining is not available for flight.

Note: The auxiliary pump is cooled by fuel; if the pump continues to operate dry, it may overheat and damage
the pump.

“L FW FUEL VALVE” or “R FW FUEL VALVE”

Abnormal Procedures – Firewall Fuel Valve Failure

The firewall fuel valve is not in the T-handle selected position.

Engine Driven High Pressure Fuel Pump Failure

Emergency Procedure – Engine Failure Or Fire In Flight

The failure of an engine’s high pressure fuel pump results in an engine flameout. There is no backup for this
pump and the engine will not be available for flight. The following annunciators will illuminate following an
engine failure:

Left Engine Failure Right Engine Failure

“L FUEL PRESS LO” “R FUEL PRESS LO”
“L OIL PRESS LO” “R OIL PRESS LO”
“L DC GEN” “R DC GEN”
“AFX DISABLE”1 “AFX DISABLE”1

1
If the autofeather system was armed and active, AFX DISABLE will illuminate once the failing engine torque drops below approximately 750 ftlb
advising the pilot that the system has disabled the autofeather function on the operative engine. The failed engine will feather as the torque drops
below approximately 350 ftlb providing the system is functioning normally. During single engine operation, it is imperative to keep both power
levers together and move both levers when changing power as if both engines are operative. The autofeather system is also looking to see where
the power levers are set to ensure that decreasing torque is a result of a malfunction and not requested by the pilot (DO NOT RETARD THE POWER
LEVER OF THE FAILED ENGINE).

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System Checks
Fuel System Test
The Fuel System is tested in the first flight checks and uses multiple systems to perform the tests. Pull CLOSED
both Firewall Fuel Valves to cut off the flow of fuel to the FCU (Fuel Control Unit). Turn ON both Standby Fuel
Pumps and with the Battery Master ON, verify that the red “L FUEL PRESS LO” and “R FUEL PRESS LO”
annunciators are illuminated to indicate no fuel pressure (or flow) through the respective Firewall Fuel Valve.
Push OPEN the both Firewall Fuel Valves (one at a time to check each side properly) and verify that the
respective red “L FUEL PRESS LO” and “R FUEL PRESS LO” annunciators extinguish to indicate fuel pressure
through the respective Firewall Fuel Valve.

Select the Fuel Transfer Switch to LEFT and verify that the red “R FUEL PRESS LO” annunciator extinguishes and
white “FUEL TRANSFER” annunciator illuminates to indicate that the RIGHT Standby Fuel pump is activated and
that the Fuel Transfer Valve has OPENED. Select the Fuel Transfer Switch to OFF and verify that the red “R FUEL
PRESS LO” annunciator illuminates and the white “FUEL TRANSFER” annunciator extinguishes to indicate the
Fuel Transfer Valve is CLOSED. Select the Fuel Transfer Switch to RIGHT and verify that the red “L FUEL PRESS
LO” annunciator extinguishes and white “FUEL TRANSFER” annunciator illuminates to indicate that the LEFT
Standby Fuel pump is activated and that the Fuel Transfer Valve has OPENED. Select the Fuel Transfer Switch to
OFF and verify that the red “L FUEL PRESS LO” annunciator illuminates and the white “FUEL TRANSFER”
annunciator extinguishes to indicate the Fuel Transfer Valve is CLOSED.

Lastly, check the Fuel Quantity (both the Main and Aux Tanks) for the required fuel and verify that all the circuit
breakers on the Fuel Control Panel are in (not tripped).

SOP – Fuel System Test (Multiple Sections)

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POWERPLANT

General Description
The Beechcraft 1900D is equipped with two Pratt & Whitney PT6A-67D free-turbine reverse flow engines. Each engine is
rated at maximum of 1279 SHP and 1353 ESHP (due to the jet thrust provided from the exhaust system). Being a free-
turbine engine, the engine uses two independent turbine sections; a single stage power turbine driving the compressor
section and a two stage power turbine driving the propeller through a reduction gearbox. The engine is elf-sufficient
since its gas generator driven oil system provides lubrication for all areas of the engine and propeller. The reverse flow
design means that air enters the intake at rear of the engine, travels forward through the engine and exits the exhaust at
the front of the engine.

Engine Description & Operation

The PT6A-67D is a modular design that splits into two major sections; the gas generator section and the power section.
This modular design requires minimal maintenance downtime, and when hot sections inspections are required, they can
be accomplished simply by splitting the engine sections and does not require the entire engine to be removed from the
aircraft.

Engine Stations
To describe positions on and within a turbine engine, engineers assign specific “station numbers” which identify
these points. Furthermore, the letter “P” for pressure, “T” for temperature or “N” for driveshaft’s can be added
to the station number when describing conditions at that station. The below shows the station numbers in
round circles, bearing numbers in squares, and engine flange points as letters in circles.

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 Engine Airflow

Gas Generator Section

Air enters at the rear of the engine through an annular plenum chamber and passes through the inlet screen and
directed forward to the compressor section. The compressor section consists of four axial stages followed by
single centrifugal stage; which operate as a single unit and mounted on the compressor turbine shaft. A row of
stator vanes are located between each compressor stage, which diffuse the air, raises its static pressure and
directs it onto the next compressor stage. The air exits the last compressor stage and passes through a diffuser
tube and straightening vanes before entering the area around the combustion chamber liner.

The air turns rearward 180° and approximately 25% of the air passes through perforations in the combustion
chamber and is mixed together with fuel injected through 14 duplex nozzles. The remaining 75% is used to
center the flame in the combustion chamber and for internal engine cooling. Once the fuel/air mixture is ignited,
the gases expand and exit the combustion chamber while turning forward 180° towards the turbine section.

As the expanding gases enter the power section, they pass over a row of compressor turbine stator vanes prior
to the compressor turbine to insure that they impact the turbine blades at the correct angle and optimize
energy transfer. Approximately 60% of the energy is extracted from the expanding gasses and is used to power
the compressor section through the compressor turbine shaft.

Compressor Bleed Valve

At low N1 rpm, the 4 axial compressor stages produce more compressed air than the centrifugal
compressor can utilize. This excess compressed air must be released to reduce back pressure on the
axial stages for smoother engine operation and help prevent an axial compressor stall. A compressor
bleed valve uses pressure differential pressure between the axial compressor (P2.5) and centrifugal
compressor (P3) to monitor for excess compressed air.

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 Operation
At low N1 rpm, as the axial compressor stages are producing more compressed air than the
centrifugal compressor (P2.5 much greater than P3) and the compressor bleed valve is forced
fully open. As the power is increased above approximately 72% N1 (Standard Atmosphere), the
centrifugal compressor can utilize more of the compressed air from the axial compressors and
the compressor bleed valve begins to close (P2.5 still greater than P3, with a smaller margin). At
approximately 90% N1 the centrifugal compressor can utilize all the compressed air produced by
the axial compressors and the compressor bleed valve is now fully closed (P2.5 now less than
P3).

Bleed Air Valve Malfunctions

If the compressor bleed valve should fail in the fully open position, the engine will operate
normally at lower power settings (at or below approximately 72% N1). However, at higher
power settings, the engine will not develop full power and the ITT will be much higher for a
given power setting. If the compressor bleed valve should fail in the fully closed position, the
engine will operate normally at higher power settings (at or above approximately 90% N1).
However, at lower power settings, the engine may start suffering from compressor stalls.

Should the compressor bleed valve fail in some intermediate position, some combination of the
above scenarios would occur depending on the actual position of the valve.

Accessory Section
Most of the engine driven accessories are mounted on the accessory gearbox at the rear of the engine and are
driven by the compressor turbine shaft. The accessories include: oil pressure pump, oil scavenge pumps, starter-
generator, high pressure fuel pump, N1 tachometer generator, air conditioning compressor (right engine only).
The propeller governor, overspeed governor and tachometer generator are mounted on the front accessory
section and driven by a shaft within the reduction gearbox.

Power Section
The expanding gases enter the two-stage power turbine section and most of the remaining energy is extracted.
The power turbine section consists of the first-stage stator vanes and turbine, followed by the interstage stator
vanes and turbine which operate as a single unit driving the power turbine shaft. The power turbine shaft then
drives the propeller shaft through a reduction gearbox.

Lubrication System
The engine is lubricated by a closed oil system consisting of a pressure pump, scavenge pumps, oil cooler, oil to fuel
heater and a breather system. The lubrication system is comprised of 2 main systems; the pressure system and the
scavenge system. The pressure and temperature control are mounted to the upper section of the oil tank and are
externally adjustable.

Oil Pressure System

The oil pressure system provides a constant supply of oil from an internal oil tank, through a filter, then to
internal and external passageways that distribute the oil to the engine bearings, reduction gears, accessory
drives, torquemeter and propeller governor (after having the oil pressure boosted). The oil filter is equipped
with a bypass valve so that in the event that it should become blocked, the engine will continue to operate with
unfiltered oil.

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 Oil Scavenge System
The oil scavenge system utilizes two double-element gear-type pumps driven by the accessory driveshafts to
collect the oil from the various components. One pump is mounted within the accessory gearbox and the other
is mounted externally on the same driveshaft as the low pressure engine driven fuel pump. Prior to oil being
returned to the oil tank by the rear element of the external scavenge pump, it passes through a thermostatic ally
bypass valve. When the oil is cold, it is returned directly to the oil tank; when the oil is hot, it is routed through
the airframe mounted cooler, located below the engine air intake, and then returned to the oil tank to maintain
a temperature of approximately 71°C (160°F). Prior to oil being returned to the oil tank by the front element of
the internal scavenge pump, it is routed through the oil-to-fuel heater then through the same thermostatically
controlled bypass valve as described above.

Oil Capacity
The total capacity of the oil system is shown in the chart below. However, most PT6A engine will seek an
oil level of 1-2 quarts lower than full and can has a normal oil consumption of 1 Quart every 10 hours. To
check the engine oil level, a dipstick is located on the upper portion at the rear of each engine and is
accessible through a quick access panel on the nacelle. The engine dipstick has two separate scales
printed on it that indicate quarts less than full; one for cold and one for hot. To check the oil level when
the engine is cold, it must be equal to ambient temperature. Otherwise, to check the oil level when the
engine it hot, it must be checked within 5 minutes after shutdown. When required oil is added through
the dipstick tube and must be identical brand and type of oil currently in the engine.

Oil Capacity
Quarts Gallons
Oil Tank 10 2.5
Oil Lines & Cooler 4.4 1.1
Total System 14.4 3.6

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Fuel System
The engine fuel delivery system consists of: low pressure fuel pump, an oil-to-fuel heat exchanger, high pressure fuel
pump, fuel control unit, flow divider with purge valve, dual fuel manifold and 14 duplex fuel nozzles. Also, some aspects
of the primary propeller governor are included, due its interaction with the engine fuel control system.

The low pressure engine driven fuel pump is driven from the engine’s accessory section and operates whenever the
engine is running. This fuel pump deliver’s fuel at approximately 45 PSI to the engine’s high pressure fuel through an oil-
to-fuel heat exchanger and fuel to the airframe fuel system jet pumps (See Fuel System chapter). Between the low
pressure engine driven fuel pump and heat exchanger is the firewall fuel shutoff valve and transducer that triggers the
“L/R Fuel Press Low” annunciator any time 10psi is not being sensed. The oil-to-fuel heat exchanger uses the heat from
the engine’s oil system to pre-heat the fuel to 21-32°C (70-90°F) which prevents ice from forming in the fuel lines. Fuel
exits the heat exchanger and enters the high pressure fuel pump, which is also driven from the accessory section, and
raises the fuel pressure to approximately 850 PSI. The fuel then enters the fuel control unit where the fuel pressure and
flow rate are metered according to engine operating conditions (see below). The high pressure engine driven fuel pump
delivers more fuel to the FCU than is required during all power settings and excess fuel is returned to the respective
collector tank through a purge line. Once the fuel has been metered appropriately, it passes by the fuel flow transducer,
which provides the data to the fuel flow gauge in the flight deck. Next, the fuel must pass through the minimum
pressure and shutdown valve which prevents fuel from continuing past this valve, whenever the fuel pressure falls
below 100 PSI. Once the fuel pressure has increased past this level, it continues to the flow divider and purge valve (see
below) through the dual fuel manifold and 14 duplex nozzles (see below) into the combustion chamber.

Fuel Control Unit

The fuel control unit meters the fuel for all modes of engine operation. The FCU operates normally by
interpreting inputs from the power lever, condition lever and P3 air. Also, within the FCU, the PY line connects
the primary propeller governor to the bellows assembly to control engine operation in the event of a propeller
malfunction.

FCU Inputs

Power Lever
The angular movement of the power lever is transformed, through a gear, into rotational movement of
the N1 governor. This movement is transferred through the 3D cam follower to fuel rotor. The rotor
subsequently rotates to increase or decrease the size of the metering port and therefore, the fuel flow
to the engine and as a result, the gas generator (N1) speed.

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 Condition Lever
The condition lever controls the minimum N1 idle speed and fuel cut-off selections. During shutdown
the condition lever actuates the cut-off valve and stops all fuel flow from within the FCU. The selection
of low idle or high idle, through a linkage, moves the power lever gear that is connection to the N1
governor to reset the minimum idle speed. Essentially the condition lever changes what N1 speed the
power lever is calling for, while at the idle gate. The minimum N1 idle speed to approximately 72% for
high idle, and a minimum of 65% for low idle.

P3 air
The P3 air input controls all accelerations (including engine starts), decelerations and environmental
changes. The P3 air enters the bellows assembly, which is located at the bottom of the fuel rotor. The
bellows assembly consists of the P3 air input, a compressor discharge pressure (CDP) sensor (bellows set
to a preset pressure) and PY vent line (see below). During changes in air pressure, caused by engine
accelerations, decelerations or changes to air density due to environmental conditions, the pressure of
the P3 air discharged by the compressor section is sensed by the bellows assembly and moves the fuel
rotor vertically to increase or decrease the size of the metering port and therefore, the fuel flow to the
engine, and as a result, the gas generator (N1) speed.

P3 air operation
The P3 air source is used to ensure the correct fuel/air mixture is maintained within the
combustion chamber during all modes of flight and environmental conditions.

Acceleration: When power lever is moved to a higher setting, the N1 governor moves the fuel
rotor to its new position. However, this will not allow all the fuel required for the new N1 speed
through the metering port, only enough to start the acceleration. As the engine accelerates, the
amount of P3 air increases and this causes the bellows assembly to move the fuel rotor vertically
and allow more fuel through the metering port. This loop continues until the engine has reached
equilibrium at the selected N1 speed.

Deceleration: When power lever is moved to a lower setting, the N1 governor moves the fuel
rotor to its new position. However, this will not immediately limit the fuel required for the new
N1 speed to be reduced completely though the metering port, only enough to start the
deceleration. As the engine decelerate the amount of P3 air decreases and this causes the
bellows assembly to move the fuel rotor vertically and reduce the fuel through the metering
port. This loop continues until the engine has reached equilibrium at the selected N1 speed.

Environmental Changes: Changes to air density will result in a change to the amount of P3 air.
This change is sensed by the bellows assembly and the metering port is change accordingly.
Examples: altitude, temperature and intake ram air

PY Air
In the event of a stuck propeller blade, the propeller governor would not be able to maintain the
propeller rpm selection or possibly rpm limitations. Should this occur, an extreme overspeed condition
would be sensed within the governor. The governor flyweights would press against a series of levers
which results in PY air being vented. This venting PY air would reduce the amount of pressure within the
fuel rotor bellows assembly and reduce the amount of fuel through the metering port. This will limit the
propeller rpm to 106% (1802 RPM with the propeller lever full forward) of the current setting. This
same system is used during operations in reverse, except that the propeller rpm limit is reset to 96%
(1650 RPM with the propeller lever full forward) of the current propeller rpm setting.

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 Flow Divider and Purge Valve

Flow Divider
The flow divider controls the flow of fuel to the primary and secondary ports on each fuel nozzle. While
the engine is not running, or during shutdown, the flow divider is in the purge position (see purge valve
below). During engine start, once the fuel has passed through the minimum pressuring and shutdown
valve, the flow divider is in the primary flow position and allows fuel through the primary manifold only.
Once the gas generator speed has increased to approximately 44 -46% N1, the fuel pressure forces the
fuel divider into the primary and secondary flow position and allows fuel through both the primary and
secondary manifolds.

Purge Valve
The aircraft is equipped with an EPA (Environmental Protection Agency) fuel Purging System. The system
ensures that any residual fuel in the fuel manifolds is consumed during engine shutdown. While an
engine is in operation, P3 air (See Pneumatic Section) is routed through a filter and pressurizes a small air
tank. The P3 air cannot exit the tank during engine operation as a check valve is held closed by fuel
pressure on the other side. Once the engine is shutdown, the fuel pressure in the manifold drops below
the pressure held within the air tank. This allows the pressurized air to pass through the check valve and
force all the residual fuel in the manifolds out though the fuel nozzles and into the combustion chamber.
During shutdown this can be observed by a momentary surge in N1 as this residual fuel is burned.

Note: If during engine shutdown, smoke continues to flow from the exhaust system for several minutes,
the EPA Purge system may not be functioning correctly and the residual fuel is slowly dripping from the
fuel nozzles into the still hot combustion chamber.

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 Fuel Manifold and Duplex Nozzles
The Fuel manifold has two separate tubes which connect the primary and secondary fuel supplies from the flow
divider to their respective ports on each of the 14 duplex fuel nozzles.

Fuel manifold (Partial) Fuel Nozzle

Ignition System
The engine is equipped with two spark-type igniters that are located within the combustion chamber at the 4 and 9
o’clock positions on the gas generator module. The ignition system consists of a solid state ignition exciter box mounted
on the airframe and a pair of high tension cables and spark igniters. The entire system is powered from the aircrafts 28
VDC system and will operate between 9V to 30V. The igniters are capable of quick light-offs over a wide temperature
range.

Starting Mode
Whenever the “Ignition and Engine Start” switch is moved to the ON position the respective igniter will be
activated and its green “Ignition” annunciator will be illuminated. The ignition system will continue to operate
until the switch is returned to the OFF position. As the engine is self sustaining, the ignition system is only
required for initial light-off during normal operations. Note: The STARTER ONLY position is for engine dry
motoring only and therefore the igniters are not activated.

Ignition and Engine Start

Auto Ignition Mode

The auto ignition system is designed to automatically engage the ignition system in the event that self sustained
combustion is lost. Each auto ignition system is controlled by the “Auto Ignition” switch in the pilots left lower
subpanel. When in the OFF position, the system is disabled. When placed in the ARM position, an oil pressure

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 sensor monitors the pressure in the torque manifold of the respective engine, and whenever the oil pressure is
below approximately 750 ft-lbs of torque, the igniters on that engine are activated.

Engine Auto Ignition

Engine Instruments

Gauges Description
ITT (Interstage Turbine Temperature)
The ITT gauge measures T5 (Temperature at
station 5) which is located between the
compressor turbine and two-stage power
turbines. Temperature is displayed in °C on
400 – Normal Operating Range an analog needle.
780°C
780-800°C Caution Range
----- 800°C Maximum Continuous
Limit
◊ 1000°C Maximum Starting Limit
Only

Torque Gauge
The Torque gauge utilizes oil pressure
developed within the reduction gearbox to
display the torque being produced on the
propeller shaft. Torque is displayed in ft-lbs
on an analog needle.
0-3750 ft- Normal Operating Range
lbs
3750-3950 Caution Range
ft-lbs
----- 3950 ft-lbs Maximum Limit

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 N2 Gauge / Propeller Gauge
The propeller gauge displays the rotation
speed of the propeller shaft. Speed is
displayed in RPM on an analog needle.

400-950 Prohibited Range (Ground)

RPM
950-1250 Normal Operating Range
RPM (Ground)
1250 – Prohibited Range (Ground)
1395 RPM
1395 – Normal Operating Range
1700 RPM (Flight)
----- 1700 RPM Maximum Limit
N1 Gauge (Gas Generator)
The N1 gauge displays the rotational speed
of the gas generator section of the engine.
Speed is displayed in % on an analog needle
with a digital readout.
65-104% Normal Operating Range
----- 104% Maximum Limit
Fuel Flow Gauge
The fuel flow gauges display the rate of fuel
being injected into the combustion
chamber and is measured downstream of
the FCU. Fuel flow is displayed in lbs/hour
on and analog needle.
No Markings
Oil Temperature
The oil temperature portion displays the
temperature of the oil immediately after
exiting the oil pressure pump. Temperature
is displayed in °C on an analog needle on
Oil Temperature the left side of gauge.
10-105C Normal Operating
Range Oil Pressure Gauge
105- Caution Range The oil pressure portion displays the
110C pressure of the oil immediately after exiting
----- 110C Maximum Limit the oil pressure pump. Pressure is displayed
in PSI on an analog needle on the right side
Oil Pressure of gauge.
----- 60 PSI Minimum Limit
60-90 PSI Caution Range
90-135 PSI Normal Operating
Range
----- 135 PSI Normal Maximum
Operating Range
◊ 200 PSI Maximum Limit

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Engine Controls
Power Lever
The power levers control the engine in one flight range and 2 ground ranges. Each range is gated to prevent the
power lever to be moved aft to the next range until the lever is lifted. The gates do not effect forward
movement of the power lever.

Flight Range
In the flight range, each power lever controls the amount of power each engine delivers to its respective
propeller by controlling the N1 governor (See Fuel Control above) to provide forward thrust. The
amount of power developed is a function of the engine’s N1 speed and decreases the further aft the
power lever is moved.

Ground Fine Range

In the ground fine range, each power lever only controls its respective propeller’s blade angle to provide
efficient aerodynamic braking during rollout and taxiing. Throughout this range the N1 speed should
remain stable at approximately 68% ± 1% and the amount of propeller braking increases the further aft
the power lever is moved.

Reverse Range
In the reverse range, each power lever controls its respective propeller’s blade angle as well as power
output to provide reverse thrust. The amount of power provided and reverse blade angle produced is
increased the further aft the power lever is moved.

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 Condition Lever
The condition lever controls the minimum idle speed and fuel cut-off selections.

Note: Although the lever has a range for each position, the selections are not variable and only actuate at the
detent stops.

High Idle / Low Idle

In these positions, a lever is adjusted on the FCU to set the minimum N1 idle speed to approximately
72% for high idle, and a minimum of 65% for low idle.

Fuel Cut-off Position

In this position, the lever actuates the cut-off valve and stops all fuel flow from within the FCU.

Staring Requirements & Procedures

Starting Requirements
The engines can be started with just the battery or with of external power assistance. The use of external power
is recommended when available. This not only to reduce the wear on the battery, but also to allow the gas
generator to rotate at a much higher speed prior to introducing fuel and results in a lower ITT during start.

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 Prior to starting an engine the battery must be checked for the following minimum battery voltages, so that in
the event of a fire, the fire bottle explosive squib that releases the extinguishing agent will have sufficient power
to be activated.

Battery Only Start - 23 VDC

External Power Assisted Start - 20 VDC

During the start sequence, the Compressor Section RPM (N1) must reach a minimum of 12% (4495 RPM) prior to
the addition of fuel. Battery only starts will typically have enough power to reach 13.5-14.5% N1 (based on a
charged battery in good condition), while external power assisted starts, typically will reach 18-21% N1 (based
upon the performance of the GPU).

Once engine ignition (light-off) occurs, there is no limit for the time the starter may be engaged for that start.
However, to prevent damage to the starters, the following cycle limitations must be observed for aborted starts:

20 Seconds ON – 30 Seconds OFF (Aborted Start)

20 Seconds ON – 60 Seconds OFF (Engine Clearing or 2nd attempted start)

20 Seconds ON – 5 Minutes OFF (2nd attempted Start or 3rd attempted start)

Note: A clearing start is only required when fuel has been introduced, but no ignition has occurred. The clearing
start necessary to purge any internally trapped fuel and fuel vapour.

Note: For continuous motoring without attempting an engine start, the starter is limited to 20 Seconds ON – 5
Minutes OFF.

Note: For restarts, following a normal engine start, a 3 minute cool down is required (time starts from when the
starter is turned off).

Starting Procedures

See current version of company Standard Operating Procedures for complete starting procedures.

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Limitations
The following limitations shall be observed. Each Column presents limitations. The limits presented do not necessarily
occur simultaneously. Refer to the Pratt & Whitney Engine Maintenance Manual for specific actions required if limits are
exceeded.

Torque GAS Gen Prop Oil Temp

Operating Maximum Oil Press
SHP Ft-Lbs RPM RPM °C
Condition ITT °C PSI (2)
(1) % N1 N2 (3) (4)

Starting --- --- 1000 (5) --- --- 0 to 200 -40 (Min)

Idle --- --- 750 65 (Min) 950 (Min) 60 (Min) -40 to 110

Takeoff (8) 1279 3950 800 104 1700 90 to 135 10 to 110

Max 1214 3750 780 104 1700 90 to 135 10 to 105

Continuous

Cruise Climb 1105 3750 (5) 760 104 1700 90 to 135 10 to 105
and Max
Cruise

Max 900 --- 760 --- 1650 90 to 135 10 to 105

Reverse

Transient --- 5000 (7) 870 (7) 104 1870 (7) 40 to 200 -40 to 110

Footnotes:

(9) Torque limit applies within range of 1000-1700 propeller rpm (N2). Below 1000 rpm,
torque is limited to 2000 ft-lbs.
(10)Normal oil pressure is 90 to 135 psig at gas generator speeds above 72%. With
engine torque below 3000 ft-lbs, minimum oil pressure is 85 psig at normal oil
temperature (60 to 70C). Oil pressure under 90 psi is undesirable; it should be
tolerated only for the completion of the flight, and then only at reduced power
settings not exceeding 2000 ft-lbs torque. Oil pressure below 60 psi is unsafe; it
requires that either the engine be shut down, or that a landing be made at the
nearest suitable airport, using the minimum power required to sustain flight.
Fluctuations of plus or minus 10 psi are acceptable.
(11)A minimum oil temperature of 55C is recommended for fuel heater operation at
takeoff power
(12)Oil temperature limits are -40°C to 105°C. However, temperatures of up to 110°C
are permitted for a maximum time of 10 minutes.
(13)Starting ITT is time limited to 5 seconds.
(14)Cruise torque values vary with altitude and temperature
(15)These values are time limited to 20 seconds.
(16)Takeoff power is time limited to 5 minutes.

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 Engine Exhaust Heating of Cabin Windows and Airframe

4) Do not operate the engines with the propellers feathered, except during external power
starts and propeller feather checks, ex that the propellers may be operated in feather at
temperatures not to exceed +5°C for a maximum of 3 minutes for purposes of airframe
deicing.
5) Do not conduct static operations in ground fine when the OAT exceeds 38°C.
6) Do not back the airplane using reverse thrust for periods longer than 10 seconds.

Reduced Power Takeoffs

If the Static Reduced Takeoff Power setting specified in the AFM Performance section is utilized, the
following limitations must be observed:

1) The runways must not be wet or contaminated with snow, slush or ice.
2) Flaps must be set at 17° for takeoff.
3) The operator must establish a means to periodically check the availability of Static
Takeoff Power, so that the use of deteriorated engine(s) will not occur. One acceptable
means is for the carrier or operator to establish operating procedures which will require
the use of Static Takeoff Power at least once during each 25 hours of operating or 25
takeoffs, whichever occurs first. Other equivalent procedures may be considered, based
on the individual operational factors involved.

Starter Limits

If ignition occurs within 20 seconds of any start attempt, there is no limit on the time the starter is
engaged for that start.

For restarts following a normal start:

3 minute cool-down, beginning when the starter is turned off.

For an aborted start (no ignition) followed by an attempted restart:

Starter is limited to 20 seconds ON (aborted start), 30 seconds OFF, 20 seconds ON (seconds start
attempt), 60 seconds OFF, 20 seconds ON (third start attempt), 5 minutes OFF.

For continuous motoring without attempting to start the engine:

Starter is limited to 20 seconds ON, 5 minutes OFF. Repeat as required

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 ITT Limits & Tolerance Torque Limits & Tolerance

Lubrication System Malfunctions & Faults

“L OIL PRESS LOW" or “R OIL PRESS LOW”
Abnormal Procedures – Low Oil Pressure Indication Checklist

Crosscheck the “L/R Oil PRESS LOW” indication against its respective oil pressure gauge to verify the condition as
the annunciator and gauge receive their data from separate sensors within the same oil pressure line.

If the oil pressure is below 60 PSI, this condition is unsafe and as per the checklist, the crew must either secure
the engine or land at the nearest suitable airport using minimum power required.

Oil Pressure below 90 PSI

Note: There is no annunciator associated with this malfunction. Two indications of this condition are:

1) Oil Pressure Gauge – Indicating below 90 PSI but above 60 PSI

2) Power Levers – A large split between power lever positions is required to match requested torque is
a possible indication and should be verified against the oil pressure gauge

Abnormal Procedures – Low Oil Pressure Indication Checklist

Oil Pressure between 60-90 PSI – This condition is undesirable and crew must not exceed 2000 ft-lbs of torque
on the respective engine for the completion of the flight.

Note: The limitation is due to the fact that the torque gauge receives its information from the engine oil
pressure and therefore will not read accurately with low oil pressure. A maximum torque setting of 2000 ft-lbs
will ensure that the engine will not exceed and actual torque of 3750 ft-lbs. The crew should be able to match
the power levers to achieve asymmetric power and use the functional engine’s torque gauge to estimate the
torque being produced on the malfunction engine’s torque gauge; however, in no way does this permit override
the maximum allowable torque setting of 2000 ft-lbs of the malfunctioning engine.

“L FUEL PRESS LO” or “R FUEL PRESS LO”

Emergency Procedures – Fuel Pressure Low

Note: This annunciation is due to a malfunction in the fuel system, not the engine. However, the failure directly
impacts on the operation of the engine. (See Fuel System for more details)

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 3) Low Pressure Engine Driven fuel Pump Failure – Only the high pressure engine driven fuel pump is operating
and the standby electric fuel pump should be turned ON to compensate for the loss which should extinguish
the annunciator.
4) Accessory Driveshaft failure - The low pressure engine driven fuel pump is mounted on the same accessory
driveshaft as the external savage oil pump and if the driveshaft has failed, the indication would be a
constantly increasing oil temperature. The external scavenge pumps are responsible for the return of oil
from engine bearings #3 and #4 as well as the reduction gearbox.

Note: The engine is limited to 10 hours of operation before overhaul or replacement of the engine driven high
pressure fuel pump is required.

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PROPELLER

General Description
The Beechcraft 1900D is equipped with two Hartzell (HC-E4A-3J), four bladed composite propellers. The propeller is
mounted on the output shaft (propeller shaft) of its respective engine’s reduction gearbox. Each constant speed, full
feathering, reversible propeller is controlled by engine oil from a single action engine driven governor. Each propeller is
equipped with a primary governor, overspeed governor and fuel topping governor (which is a component of the primary
governor). As the power turbine section of the engine has no physical connection to the gas generator section, the
propeller can rotate freely while the engine is shutdown. To avoid wear on internal components, the propeller tie should
be installed whenever the aircraft is not parking in a hanger.

Propeller Blade Angles

The propeller blade angle is the angle between the chord of each blade and the propellers plane of rotation. As the
design of the blade, the blade angle is greater at the root than at the tip. The blade angle for the Beechcraft 1900D is
measured at the 42” blade station (42 inches from the blade root).

Normal Operation Angles

Operational Mode Resultant Angle

Feathered 79° ± 0.5°
Cruise 30°-40°
Flight Idle 12.7°
Ground Idle 4.9°
Maximum Ground Fine -5.9°
Maximum Reverse -14.5° ± 0.5°

Propeller Governing
The propeller pitch is controlled during all normal operations by centrifugal counterweights on each propeller blade and
feathering spring within the propeller servo as well as the primary governor (and fuel topping governor in the reverse
range). The overspeed governor and fuel topping governor (in the forward range) are used to control the propeller pitch
if a malfunction of the normal propeller control system should occur.

Primary Governor
The primary governor controls varies the pitch of the propeller by ensuring that the load torque on the propeller
is equal to the engine torque during changing flight conditions. The primary governor is located at the 12 o’clock
position on the front of the reduction gearbox.

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 Primary Governor Operation
Engine oil is used to control the propeller pitch; however, the engine oil pressure is insufficient and is
raised to approximately 800-850 PSI by the governor oil pump. When oil reaches the maximum pressure
relief valve opens and allows the oil to recalculate through the pump. Flyweights within the governor
react to the propeller lever position and actual the propeller speed. These flyweights move a pitot valve
plunger to control the propeller pitch, depending on the conditions below.

On Speed Condition
When the propeller lever selection and the propeller speed are in equilibrium (the force of the
speeder spring on the flyweights and the centrifugal force of the flyweights are equal), the pitot
valve plunger is centered. This covers the oil discharge port and no oil flows in or out of the
propeller servo.

Underspeed Condition
When the propeller lever selection is set to a higher speed than the propeller currently at (the
force of the speeder spring on the flyweights is greater than the centrifugal force of the
flyweights), the pitot valve plunger moves down. This opens the oil discharge port to the
governor pump and oil is forced into the propeller servo. This increases the force on the
propeller servo and results in less of a propeller pitch (higher speed, lower blade angle).

Overspeed Condition
When the propeller lever selection is set to a lower speed than the propeller is currently at (the
force of the speeder spring on the flyweights is less than the centrifugal force of the flyweights),
the pitot valve plunger moves up. This opens the oil discharge port to the engine oil sump and
allows oil out of the propeller servo. This decreases the force on the propeller servo and results
in a more course propeller pitch (lower speed, higher blade angle).

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 Feather Condition
When the propeller lever is pulled back past the feathering detent (regardless of the speeder
spring or flyweight forces), the pitot valve plunger is moves fully up. This opens the oil discharge
port to the engine oil sump and allows all the oil out of the propeller servo. To allow the
propeller to rapidly move to the fully feathered position, a propeller dump valve is incorporated
within the governor and opens a poppet valve when the propeller lever reaches the feathering
stop. Both these paths remove all the force on the propeller servo and results in the propeller
pitch increasing (lower speed, higher blade angle) to the fully feathered position of 79° ± 0.5°.

Beta Range Operation

To allow the propeller to enter the beta range (ground fine and reverse), the primary governor is
equipped with a beta valve that controls the low pitch stop (see below). Movement of the
power lever into the beta range causes the beta valve to reset the propeller low pitch stop.

Ground Fine Range

When the power lever is moved into the ground fine range, each power lever only
controls its respective propeller’s blade angle by moving the beta valve, to provide
efficient aerodynamic braking. Throughout this range the N1 speed should remain
stable at approximately 68% ± 1% and the amount of propeller braking increases the
further aft the power lever is moved.

Reverse Range
When the power lever is moved into the reverse range, each power lever controls its
respective propeller’s blade angle by moving the beta valve, as well as engine power
output to provide reverse thrust. The amount of power provided and reverse blade
angle produced is increased the further aft the power lever is moved. Also, when the
power lever is moved into the reverse range, the Fuel Topping Governor setting is
changed to 96% (1650 RPM with propeller levers full forward) of selected RPM. This will
limit the propeller speed in reverse by limiting engine power output.

Note: Since oil pressure is forced into the propeller servo to increase the propeller pitch
into the reverse range, if the primary governor flyweights where allowed to govern the
propeller speed at 1700 RPM (with the propeller lever full forward). It would dump the
oil pressure from the propeller servo and allow the propeller pitch to exit the reverse
range. Therefore the propeller speed must be limited below the operational range of
the primary governor when operating in the reverse range.

Low Pitch Stops

Propeller systems that do not include utilize ground fine or reverse ranges contain a mechanical
stop that prevents the propeller pitch from decreasing (lower blade angle, higher RPM) past a
preset level. As the PT6-67D has been designed to provide both these ranges, the propeller has
been equipped with hydraulic low pitch stops that utilize the primary governor oil and Beta
valve to limit the propeller pitch.

Note: Whenever a sustained propeller speed is below the speed selected by the propeller lever,
the propeller pitch is on the low pitch stop (not including prop feathering).

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 Low Pitch Stop Operating

Flight Idle Stop

When the power lever is reduced to the idle stop (while airborne), the primary governor
attempts to maintain the selected propeller speed by forcing oil into the propeller servo.
As the propeller speed decreases and is no longer being governed by the primary
governor flyweights, the propeller pitch decreases and moves the low pitch stop collar
forward. Once the propeller pitch has decreased to 12.7° the beta valve is pulled
outwards and blocks the oil from the governor pump from entering the propeller servo
and prevents the propeller pitch from decreasing any further.

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 When the power lever is increased again, the propeller speed will increase and the beta
vale moves rearward at the same time that the primary governor flyweights are able to
govern the propeller speed.

Ground Idle Stop

Upon touchdown, the right main landing gear squat switch (or lifting of either power
lever) activates an electric solenoid (powered). This solenoid extends the power lever
linkage slightly which changes the position at which the beta valve is pulled out. Once
the propeller pitch has decreased to 4.9° the beta valve is pulled outwards and blocks
the oil from the governor pump from entering the propeller servo and prevents the
propeller pitch from decreasing any further.

Once the aircraft has taken off (weight off wheels) or the Low Pitch Test switch is held
down, the electric solenoid is deactivated (unpowered) and the power lever linkage
retracts slightly and resets the position at which the beta valve is pushed in, back to the
flight idle stop position.

Ground Fine Stop

When the power lever is lifted over the idle gate, the beta valve is moved rearwards and
allows oil to enter the propeller servo and further decrease the propeller pitch. As the
propeller pitch decreases, it moves the low pitch stop collar forward. Once the propeller
pitch has decreased to -5.9° the beta valve is pulled outwards and blocks the oil from
the governor pump from entering the propeller servo and prevents the propeller pitch
from decreasing any further.

As the power lever is moved forward, the beta valve is pulled outwards further. This
allows oil to exit the propeller servo into the engine oil sump which allows the propeller
pitch to increase. As the propeller pitch increases, it moves the low pitch stop collar

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 rearward. Once the propeller pitch has increased to the requested setting, the beta
valve is pushed in and blocks any more oil from exiting the propeller servo.

Reverse Stop
When the power lever is lifted over the ground fine gate, the beta valve is moved
rearwards and allows oil to enter the propeller servo and further decrease the propeller
pitch. As the propeller pitch decreases, it moves the low pitch stop collar forward. Once
the propeller pitch has decreased to -14.5° the beta valve is pulled outwards and blocks
the oil from the governor pump from entering the propeller servo and prevents the
propeller pitch from decreasing any further. However, even though the propeller pitch is
being controlled, during reverse operation, the engine powers up to provide reverse
thrust and results in an increased propeller speed (RPM). The propeller speed is
therefore limited to 1650 RPM by the fuel topping governor (see beta range operation
above).

As the power lever is moved forward, the beta valve is pulled outwards further. This
allows oil to exit the propeller servo into the engine oil sump which allows the propeller
pitch to increase. As the propeller pitch increases, it moves the low pitch stop collar
rearward. Once the propeller pitch has increased to the requested setting, the beta
valve is pushed in and blocks any more oil from exiting the propeller servo.

Overspeed Governor
In the event of a failure of the primary governor to control the propeller pitch, the overspeed governor
is a safeguard to prevent against a propeller overspeed. If the propeller was not equipped with an
overspeed safeguard and the primary governor were to fail to limit the maximum propeller speed; the
propeller pitch would be forced to decrease into full reverse. The overspeed governor operates on the
same principles as the primary governor, except that the propeller speed setting preset to 106% of the
maximum propeller speed or 1802 RPM. Since the overspeed governor has no mechanical controls, a
testing solenoid resets the overspeed governor to 1565 ± 30 RPM for ground testing. So that maximum
engine torque is not exceeded, torque is limited to 3538 ft-lbs with propellers overspeeds of up to 1802
RPM.

Overspeed Governor Operation

Primary Governor Operating Normally

During all propeller speeds below 1802 RPM, (the force of the preset spring on the
flyweights is greater than the centrifugal force of the flyweights), the overspeed
governor pitot valve plunger blocks the oil dump port and allows the oil to enter and
exit the propeller servo as controlled by the primary governor.

Primary Governor Malfunction

If the primary governor should fail in such a way that it is unable to stop the flow of oil
into the propeller servo, at 1802 RPM (the force of the preset spring on the flyweights is

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 less than the centrifugal force of the flyweights), the overspeed governor pitot valve
plunger opens the oil dump port and allows oil to exit from the propeller servo. The
overspeed governor will repetitive open and close the dump port to maintain a
maximum propeller speed at 1802 RPM.

Fuel Topping governor

In the event of a stuck propeller blade, the propeller governor would not be able to maintain the
propeller rpm selection or possibly rpm limitations. Should this occur, an extreme overspeed condition
would be sensed within the governor. The governor flyweights would press against a series of levers
which results in PY air being vented. This venting PY air would reduce the amount of pressure within the
fuel rotor bellows assembly and reduce the amount of fuel through the metering port. This will limit the
propeller rpm to 106% (1802 RPM with the propeller lever full forward) of the current setting. This
same system is used during operations in reverse, except that the propeller rpm limit is reset to 96%
(1650 RPM with the propeller lever full forward) of the current propeller rpm setting.

Autofeather System
Due to the great amount of drag that is created from a failed engine, the Beechcraft 1900D is equipped with an
autofeather system that automatically detects an engine failure and feather the respective propeller. The system has
several conditions that must be met to arm the system, the autofeather switch is in the “ARM” position and the power
levers must set above 85-90% (positions switches in pedestal). With these conditions met, the green “L/R
AUTOFEATHER” annunciators on the lower annunciator panel and “AFX” annunciators beside the torque gauges will
illuminate. While armed, the autofeather system continually monitor the oil pressure within the torque manifold for a
loss of torque, and therefore, engine failure .Should the amber “AFX DISABLE” annunciator illuminate, the autofeather
system is not capable of feathering the propellers (System logic prevents feathering both propellers at the same time
and/or prevents feathering the propeller on the good engine following an engine failure).

Note: As the propeller autofeather system must be operable for all flights and must be armed for takeoff, climb,
approach, and landing. As a warning, if the landing gear is not fully retracted (triggered by left landing gear uplock
switch), and the autofeather switch is OFF, the amber “AUTOFTHER OFF” annunciator will illuminate.

Engine Failure Autofeather Sequence

With both engines operating and the above conditions met, the autofeather system is armed. Should an engine
fail, the respective torque will decrease quickly. As the torque decreases on the inoperative engine below
approximately 750 ft-lbs, the autofeather high pressure switch is activated, this causes the inoperative engine’s
green “L/R AUTOFEATHER” annunciator to extinguish and the amber “AFX DISABLE” annunciator to illuminate.
This indicates that the system has sensed an impending engine failure and has disabled autofeathering of the
operative engine. If the torque on the inoperative engine continues to decrease below approximately 350 ft-lbs,
the autofeather low pressure switch activates, this powers the autofeather dump solenoid which immediately
dumps the oil pressure from within the propeller governor and feathers the propeller.

Note: During single engine operation, it is imperative to keep both power levers together and move
both levers when changing power as if both engines are operative. The autofeather system is also
looking for the microswitches in the power quadrant to see where the power levers are set to ensure
that decreasing torque is a result of a malfunction and not requested by the pilot (DO NOT RETARD THE
POWER LEVER OF THE FAILED ENGINE as it could result in the propeller having to be manually
feathered).

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Propeller Synchrophaser
The Beechcraft 2900D has a propeller synchrophaser system installed to decrease cabin noise created from the
propellers. When activated, the propeller synchrophaser automatically matches the left and right propeller speeds and
positions at a preset phase relationship.

Signal pulses are received from magnetic pickups (one pulse per propeller revolution) within each propeller hub and fed
to a controller box. This controller box increases the propeller RPM of the slower propeller to match the faster
propeller’s RPM by changing the voltage supply to the electromagnetic coil (speed biasing coil) within each primary
governor. The synchrophaser had been designed so that it can only increase the propeller RPM so that in event of an
engine failure (while the synchrophaser is on), that the operative engine propeller RPM cannot fall below the RPM set by
its respective propeller lever. Also, the synchrophaser can only increase the propeller RPM over a limited range of
approximately 25 ± 2 RPM.

Note: If the synchrophaser is ON, however does not properly match the propeller RPMs, then the synchrophaser may
have reached its range limit. To correct this, turn OFF the synchrophaser, manually synchronize the propeller RPMs, then
turn ON the synchrophaser. Also, when the synchrophaser is turned ON it can take a few seconds, up to 30 seconds for
the input signals to drift within the phasing range.

Limitations
Propeller Rotational Speed Limits
Transients not exceeding 20 seconds 1870 RPM
Reverse 1650 RPM
Minimum idle speed 950 RPM
Red arc ground operation prohibited ranges 400-950 RPM
1250-1395 RPM
All other conditions 1700 RPM

Propeller Rotational Overspeed Limits

The maximum propeller overspeed limit is 1870 RPM and is time limited
to 20 seconds. Sustained propeller overspeed faster than 1700 RPM
indicate failure of the primary governor. Flight may be continued at
propeller overspeeds up to 1802 RPM provided torque is limited to 3538
ft-lbs. Sustained propeller overspeeds above 1802 RPM indicate failure
of both the primary governor and the hydraulic overspeed governor.

Propeller Autofeather
The propeller autofeather system must be operable for all flights and
must be armed for takeoff, climb, approach, and landing.

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Propeller Malfunctions
Propeller RPM goes to 1700 when a lower RPM was selected

Broken propeller control linkage (governor is spring loaded to return to 1700 RPM). It is recommended that the
autofeather be armed and remain on during the rest of the flight in case of an engine failure as the propeller
control is now unable to feather the propeller.

Troubleshooting

a) Attempt to adjust propeller lever results no change to RPM, N1, Fuel Flow, ITT or Torque
b) Pair propeller levers at 1700 RPM for duration of flight

Propeller RPM goes to 1802 when a lower RPM was selected (*RPM stays at 1802 regardless of propeller
control movement)

Primary governor has failed and the high overpeed governor is active. Torque now limited to 3538 ft-lbs to
prevent engine over-torque.

Troubleshooting

a) Attempt to adjust propeller lever results in no change to RPM, N1, Fuel Flow, ITT or Torque
b) Pair Propeller Levers (as close as possible)
c) Torque limited to 3538 on the engine with propeller at 1802 RPM

Propeller RPM changes with changes in airspeed and/or climbs/descents

Primary governor not able to control propeller RPM due to a stuck blade and fuel is being limited to prevent
RPM from increasing beyond 106% of selected. Propeller control movement will cause changes in N1, fuel flow,
ITT and torque.

Troubleshooting

a) Attempt to adjust propeller lever results in changes to RPM, N1, Fuel Flow, ITT and Torque
b) Pair Propeller RPM, constant adjustments will be required. Maximum power available from
malfunction propeller and therefore engine will be limited depending on the pitch position (blade
angle) the blade is stuck at.

Propeller System Checks

Propeller Autofeather System Test
The propeller autofeather system test switch is used to override the power lever switches within the pedestal to
allow the system to be tested at lower power settings.

To test the “AUTOFTHER OFF” annunciator warning system, move the autofeather switch to the “ARM” position
and verify the green “AUTOFTHER OFF” annunciator extinguishes, then return the autofeather switch to the
“OFF” position and verify the “AUTOFTHER OFF” annunciator illuminates.

To test the autofeathering system, with both propeller levers in the taxi range, hold the autofeather switch in
the “TEST” position and observe the amber “AFX DISABLE” annunciator illuminates. Increase the power on both
power levers over 1000 ft-lbs and verify the amber “AFX DISABLE” annunciator extinguishes and both green “L/R

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 AUTOFEATHER” and green “AFX” annunciators illuminate. Slowly reduce the power on one engine (simulated
failing engine) and verify that through 750 ft-lbs the amber “AFX DISABLE” illuminates and the opposite engine’s
(simulated good engine) green “L/R AUTOFEATHER” annunciator extinguishes. Continue slowly reducing the
power below 500 ft-lbs and verify that the respective (simulated failing engine’s) green “L/R AUTOFEATHER”
annunciator extinguishes as the propeller feathers momentarily. Return the power lever to above 1000 ft-lbs
and repeat the procedure with the other power lever. With both power levers above 1000 ft-lbs, reduce the
power on both power levers to idle and verify that through 750 ft-lbs the amber “AFX DISABLE” illuminates and
both green “L/R AUTOFEATHER” annunciators extinguish and that neither propellers feather. Finally, release the
autofeather switch and it will return to the “OFF” position.

Autofeather ARM/OFF/TEST

SOP – Propeller Autofeather System Test Procedure

Propeller Overspeed Test

The overspeed test switch used to change the RPM setting of the overspeed governor to test its operation.
When the switch is held to “Prop Test - Overspeed” position, a solenoid within the overspeed governor resets it
operational speed from 1802 RPM down t 1565 ± 30 RPM, which sets it below the speed at which the primary
governor will operate (with the propeller lever full forward). With the switch held and the propeller levers full
forward, each power lever is increased sufficiently to observe the propeller RPM governing by overspeed
governor (within the range above) to successfully pass this test. Also, while governing at 1535-1595, continue to
increase torque to 2500 ftlb or more and the rudder boost will begin to deflect the rudder pedals to assist the
pilot with increased control (rudder pedal) forces. Pressing the control wheel disconnect to the first level will
momentarily interrupt the rudder boost and will be felt as a kick on the rudder pedals.

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 Prop Test OVERSPEED

SOP – Propeller Overspeed Test Procedure (Includes Rudder Boost Test)

Propeller Low Pitch Stop Test

The low pitch test switch is used to test the solenoids that control the low pitch stops between flight idle and
ground idle. The switch is located in on the bottom of left subpanel. When the switch is held down in the “Prop
Test - Low Pitch” position, the solenoid believes the aircraft is airborne and switches to the flight idle position
(overrides the squat switch position). This will cause both propellers to decrease by approximately 200 RPM.
Now to test the power lever override switches, each power lever is lifted individually and the both propellers
should change to the ground idle low pitch stop, this is identified by an increase of approximately 200 RPM.
Once the power lever is released, both propellers should return to the flight idle low pitch stop, identified by
both propellers decreasing by approximately 200 RPM. Finally, the low pitch test switch is released and both
propellers should return to its normal operating mode with the squat switch controlling the low pitch stop
solenoid. This should return the system to the ground idle low pitch stop position, identified by an increase of
200 RPM to both propellers.

Prop Test OVERSPEED

SOP - Low Pitch Test Procedure

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PRESSURIZATION

General Description
The Beechcraft 1900D can be pressurized to reduce the physiological effects on both the crew and passengers and
decreases or all together removes the need for supplemental oxygen during normal operations. Compressed air is used
from both engines to allow the cabin to be pressurized and a pressurization controller connected to a pair of outflow
valves to maintain and control the cabin pressurization. The design of the aircraft pressurization system will maintain a
cabin altitude of approximately 9,500’ with the aircraft operating at its maximum operating altitude of 25,000’.

System Operation
Bleed Air
Bleed air from the P3 compressor stage of both engines is used to provide air for pressurization, heating and
cooling of the cabin. Once the P3 air exits the engine, it passes through a set of precooler valves (directs air into
or to bypass the heat exchanger) that regulates the bleed air temperature to 450˚F ± 25˚F, and through a
pressure regulator/shutoff valve that maintains a bleed air pressure of 37 ± 3 PSI before it is ducted into the Air
Machine Cycle or directly into the cabin.

To open the bleed air system on an engine, the respective bleed air valve switch, which is located on the left side
of the co-pilot’s subpanel, must be moved to the “OPEN” position. This powers the bleed air control circuitry
that controls the bleed air temperature and pressure.

To protect the pressurization and environmental systems, each bleed air line is equipped with an over-pressure
sensor (44 ± 1 PSI) and an over-temperature sensor (500˚F ± 10˚F) that will automatically close the precooler
valves and shutoff valve to isolate the malfunctioning bleed air source from the rest of the system. For
additional protection a secondary set of over-pressure and over-temperature sensors (same values) are installed
in the common bleed air line (after the left and right bleed air lines connect together). Should any sensor in the
left or right bleed air lines be triggered, the respective red “L ENVIRO FAIL” or “R ENVIRO FAIL” annunciators will
illuminate. Should either of the secondary sensors be triggered, both red “L ENVIRO FAIL” or “R ENVIRO FAIL”
annunciators will illuminate.

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 Outflow Valves
The two outflow valves are used to adjust the amount of air exiting the cabin as directed by the pressurization
controller to maintain the desired cabin pressurization.

While on the ground, the left main landing gear squat switch activates the dump solenoid and applies full
vacuum to the outflow valve. This causes the outflow valve to fully open and prevent cabin pressurization. While
in-flight, the dump solenoid closes and the preset solenoid opens and allows the pressurization controller to
control the position of the outflow valve by adjusting the amount of vacuum applied. As vacuum is decreased,
the outflow valve closes and prevents air from exiting the cabin, thus pressurizing the cabin.

To prevent the cabin from exceeding the maximum pressure differential (5.0 ± 0.1 PSI), the outflow valve
compares ambient (static) pressure, from a set of static ports on the sides of the fuselage aft of the cargo door,
to cabin air pressure. In the event that the pressure differential should exceed 5.1 PSI, a poppet valve will open
and decrease the vacuum pressure on the outflow valve which will decrease the cabin pressure differential
below 5.1 PSI.

Pressurization Control

Cabin Pressure Control Switch

The cabin pressure control switch, located on the lower center pedestal, has 3 positions:

DUMP – The dump solenoid is OPEN and the preset solenoid is CLOSED, this prevents the cabin
from pressurizing (or dumps the cabin pressure).

PRESS – When on the ground, the dump solenoid is OPEN and the preset solenoid is CLOSED,
this prevents the cabin from pressurizing. Once in-flight, the dump solenoid CLOSES and the
preset solenoid OPENS, this allows the pressurization controller to control the cabin
pressurization.

TEST – Holding the switch to this position bypasses the landing gear squat switch, CLOSES the
dump solenoid and OPENS the preset solenoid, this allows the aircraft to pressurize while on the
ground for testing the pressurization.

Pressurization Controller
The pressurization controller is used by to maintain the cabin pressure at the selected altitude by
controlling the outflow valves. The pressurization controller has adjustments for cabin altitude and cabin
rate-of-change (Cabin Climb Rate). When the preset solenoid is open, the pressurization controller
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 compares the current cabin altitude to the selected cabin pressure and a reference volume tank to
control the outflow valves and achieve the desired cabin pressure and rate-of-change. Whenever the
selected cabin altitude is different from the actual cabin altitude, the pressurization controller will
change the cabin pressure at the rate-of-change selected to achieve balance. A pneumatic relay,
attached to the outflow valve control vacuum line of the pressurization controller is used to limit the
rate of change in cabin pressurization, as well as, to amplify the vacuum output so that both outflow
valves can be controlled in tandem.

Pressurization Gauges
Cabin Pressure Differential (Inner Scale)
The cabin pressure differential portion displays the
differential pressure between the cabin and
ambient pressure. Pressure differential is displayed
in PSI on the smaller inner scale.

Cabin Pressure Differential Cabin Altitude (Outer Scale)

0-5.1 Normal Operating The cabin altitude portion displays the current
Range cabin altitude. The cabin altitude is displayed in
>5.1 Prohibited Range 1000’s of feet.

Cabin Altitude
0-35,000’ Indication Only

Cabin Climb Rate

The cabin climb rate gauge displays the rate-of-
climb or rate-of-descent of the cabin altitude. The
rate of change is displayed in 100’s of feet.

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 Cabin Altitudes vs Differential PSI

Cabin Differential Warning

Should the pressurization controller and outflow poppet valve fail to control the cabin pressure, the red “CABIN
DIFF HI” annunciator will illuminate when a sensor in the center pedestal detects a cabin pressure differential of
> 5.25 ± 0.15 PSI. (See Malfunctions & Faults below)

Cabin Altitude Warning

Should the cabin altitude exceed 10,000, reach 10,000 +0 / -500 feet, the red “CABIN ALT HI” annunciator will
illuminate. (See Malfunctions & Faults below)

Optimal Pressurization Profile

Take-off & Final
Ground Cruise Step Descent Ground
Climb Descent
Pressure Field Field
Setting Cruise Alt Cruise Alt Cruise Alt Step Altitude Elevation Elevation
+ 1000’ + 1000’ + 1000’ + 1000’ + ___ + ___
(Note1) (Note1)
Cabin VSI +500
--- 0 -500 -500 ---
(FPM) (Nominal
Cabin Increasing Stable Decreasing
Stable
Differential 0 =< 5.1 ± 0.1 =< 5.1 ± 0.1 To 0 at 500’ 0
2 = 5.1 ± 0.1 3
(PSID) (Note ) (Note ) AGL
Note1: The correct setting is 500’ above the field elevation when the pressure at the field is standard (29.92
in HG). To compensate for non-standard pressure at the field, use settings displayed on page N-25 of the
Beechcraft 1900D checklists. The correct setting is important to avoid landing while the aircraft is
pressurized.

Note2: If the aircraft maintains a high rate-of-climb, the cabin pressure differential will reach maximum prior
to the cabin reaching the desired altitude. This will cause the poppet valve within the outflow valve to open
at 5.2 ± 0.1 psi and the cabin altitude will climb at the same rate at the aircraft until the altitude set on the
pressurization controller is set.

Note3: The cabin differential pressure will vary during depending on the rate of descent of the aircraft.
However, it will return to 5.1 ± 0.1 (or lower if cabin altitude is already at the adjusted field elevation) once
level at the step decent altitude.
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Malfunctions & Faults
“L ENVIRO FAIL” and/or “R ENVIRO FAIL”
Emergency Procedures – “Environmental Failure”

Disagreement between switch position and valve position or valve is not powered. Valve automatically closes in
the event that the environment air duct over pressure (Activates at > 44 ± 1 PSI) and/or over temperature
(Activates at > 500˚F ± 10˚F).

1) Bleed Air Valve (affected side) – Cycle the bleed air valve switch to “ENVIRO OFF”, then back to “OPEN”

If the respective “L ENVIRO FAIL” or “R ENVIRO FAIL” annunciator extinguishes, continue operation. If
annunciator remains illuminated:

2) Bleed Air Valve (affected side) – Select the bleed air valve switch to the “ENIVRO OFF” position.

“CABIN ALT HI”

Emergency Checklist – “Cabin Decompression”

Cabin altitude sensed at 10,000’ +0 / -500

1) Oxygen Drill (if required by SOPs)

2) Determine cause of pressure loss
a. Cabin Pressure Switch – check switch is in the PRESS position
b. Cabin Controller – check the controller is set correctly for required altitude
c. Bleed Air Valves – check each switch is in the OPEN position
3) If unable to correct problem and cabin altitude approaches 15,000’, or if decompression is rapid
a. Execute Emergency Descent Procedure – Drill & Checklist (Per SOPs)
b. Oxygen Duration – Confirm output is correct for existing cabin altitude

Note1: If operating with a single bleed air source inoperative, the cabin pressure differential will not be able to
reach/maintain the same pressure differential that would be available with both bleed air sources operating
normally.

Note2: If the cabin differential pressure does not reach the desired level (cabin altitude higher than selected),
increasing the cabin temperature will often correct the problem. This is due to the large number of potential
leaks present within the ACM unit, when a higher cabin temperature is selected, more bleed air directly enters
the cabin and bypasses the ACM.

“CABIN DIFF HI”

Emergency Checklist – “High Differential Pressure”

Cabin differential pressure sensed at > 5.25 ± 0.15 PSI

1) Bleed Air Valves – Select both bleed air switches to the “ENIVRO OFF” position.
2) Oxygen (Crew & Passengers) – May be required now that the cabin altitude will increase until it
reaches the aircraft’s altitude.
3) Descend - May be required now that the cabin altitude will increase until it reaches the aircraft’s
altitude.

Note: After turning off both bleed air valves, the cabin altitude will increase and cause the activate the
red “CABIN ALT HI” annunciator if the aircraft’s altitude is greater than 10,000 +0/-500 feet.

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Pressurization System Check
To cabin pressurization control switch is used in the test position the cabin pressurization on the ground. With all the
windows and doors closed, select both bleed air valve switches to the “OPEN” position, dial the cabin altitude selector
knob to 1000’ below the field pressure elevation and position the arrow on the rate control selector to the 12 o’clock
position. Next, hold the cabin pressurization control switch to the “TEST” position and wait for the cabin altitude begins
to descend (may take up to 45 seconds), then release the switch and ensure that the cabin altitude climbs back to field
pressure elevation to verify the aircraft has depressurized. The cabin pressurization control switch should now remain in
the “PRESS” position.

SOP – Pressurization Test

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Pressurization/Environmental Ducting

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ENVIRONMENTAL SYSTEM

General Description
To maintain the desired temperature within the cabin, the Beechcraft 1900D is equipped with a bleed air system for
heating, Air Cycle Machine for cooling and Vapour Cycle machine for additional cooling when required.

Components
Bleed Air
Bleed air from the P3 compressor stage of both engines is used to provide air for pressurization, heating and
cooling of the cabin. Once the P3 air exits the engine, it passes through a set of precooler valves (directs air into
or to bypass the heat exchanger) that regulates the bleed air temperature to 450˚F ± 25˚F, and through a
pressure regulator/shutoff valve that maintains a bleed air pressure of 37 ± 3 PSI before it is ducted into the Air
Machine Cycle or directly into the cabin.

To open the bleed air system on an engine, the respective bleed air valve switch, which is located on the left side
of the co-pilot’s subpanel, must be moved to the “OPEN” position. This powers the bleed air control circuitry
that controls the bleed air temperature and pressure.

To protect the pressurization and environmental systems, each bleed air line is equipped with an over-pressure
sensor (44 ± 1 PSI) and an over-temperature sensor (500˚F ± 10˚F) that will automatically close the precooler
valves and shutoff valve to isolate the malfunctioning bleed air source from the rest of the system. For
additional protection a secondary set of over-pressure and over-temperature sensors (same values) are installed
in the common bleed air line (after the left and right bleed air lines connect together). Should any sensor in the
left or right bleed air lines be triggered, the respective red “L ENVIRO FAIL” or “R ENVIRO FAIL” annunciators will
illuminate. Should either of the secondary sensors be triggered, both red “L ENVIRO FAIL” or “R ENVIRO FAIL”
annunciators will illuminate.

Air Cycle Machine (ACM)

The air cycle machine, located within the left wing root between the fuselage and engine nacelle, is used to
provide cooling for the cabin. The ACM cools the bleed air through the following process:

1) Hot bleed air enters the ACM and passes through the primary heat exchanger which removes some
of the excess heat.
2) The bleed air is ducted into ACM compressor where it is compressed further, resulting in an
increased pressure and temperature.
3) The compressed bleed air now passes over the secondary heat exchanger which removes the excess
heat created from its compression.

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 4) The compressed bleed air is now allowed to expand as it passes over the ACM turbine and as a
result of the expansion, the air cools to a temperature below its starting ambient temperature. The
expanding air is also used to drive the ACM turbine, which in turn drives the ACM compressor and
cooling fan.

Also, as the secondary heat exchanger cools the compressed bleed air, moisture condensation occurs
and is collected, then injected through a fogging nozzle into the expanding air to assist in cooling.

The ACM cooling fan draws cooling air form the ambient air intake, located under the ACM near the front of the
wing, and is ducted over the primary and secondary heat exchangers and then is discharged overboard.

ACM Vents

Vapour Cycle Machine (VCM)

The vapour cycle machine, located within the right wing root between the fuselage and engine nacelle, is used in
conjunction with the ACM when additional cooling is required. The VCM cools the cabin in the same manner
that an air conditioner cools a house and is powered by the right engine.

VCM Cycle
The VCM system utilizes a compressor to compress the refrigerant into a high pressure and high
temperature gas. The refrigerant then passes through a 40,000-BTU condenser to remove excess heat
and results in the refrigerant condensing into a high pressure, low temperature liquid. The refrigerant is
then allowed to expand which results in a low pressure, low temperature gas that passes through two
12,500-BTU evaporators within the cabin. Cabin air is circulated over the evaporators where heat is
transferred from the cabin air to the refrigerant, therefore cooling the cabin. The refrigerant then
returns to the compressor and the cycle is repeated.

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 To protect the VCM from damage, it is equipped with an refrigerant over-pressure (390 ± 10 PSI) and under-
pressure (2.5 ± 0.5 PSI) cut-off switches that will disable the VCM should an a pressure problem occur (should
either of these be activated, it can only be reset on the ground through switches located within the left side of
the avionics bay). Also, to prevent the condenser from icing up, the VCM compressor clutch is disengaged when
the air cooling inlet temperature is 40°C or below. In addition to protecting the VCM, to prevent overloading the
engine at low speeds the compressor clutch will automatically disengage should the engine speed fall below
62% N1.

During Flight the condenser is cooled by ram air, however, during ground operations a condenser blower is
activated by the right main landing gear squat switch, to draw air over the condenser.

Environmental Controls
The temperature control panel is located on the left side of the co-pilot’s subpanel and consists of the cabin
blowers switch and both the automatic and manual temperature controls.

Environmental Controls

Blowers Control
The two vent blowers are located directly forward of each of the evaporator coils and are used to
circulate air within the cabin and when the VCM is operating to force air through the evaporator coil.
The vent blowers are controlled by a 3 position switch:

HI – The vent blowers operate at high speed.

OFF – The vent blowers are off.

AUTO – The vent blowers operate at low speed whenever the temperature mode is selected to
any position other than “OFF”.

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 Mode Control

AUTO
When the mode controller is selected to “AUTO” the ACM bypass (therefore ACM) and VCM
operate automatically and are controlled by the cabin temperature control module. The “CABIN
TEMP” knob selects the desired temperature and the cabin temperature control module then
sends the appropriate “heat” or “cool” command to the ACM bypass valve. The bypass valve will
continue to change position in response to the temperature selected and current cabin
temperature.

Maximum heat provided to the cabin when the ACM bypass valve is fully open (bypasses ACM)
and maximum cooling when the ACM bypass valve is fully closed. Should the ACM cooling be
insufficient to reach the desired temperature selected, the VCM will engage to provide
additional cooling once a switch within the ACM bypass valve indicates that the valve is fully
closed. Once engaged, the VCM will stay engaged until a switch in the bypass valve indicates
that the valve is fully open.

Note: The “MAN TEMP” switch will only move the ACM bypass valve for as long as the switch is
held, however, the ACM bypass valve will return to the position commanded by the
temperature control module when the switch is release.

MAN
When the mode controller is selected to “MAN” the ACM bypass valve position is controlled by
the “MAN TEMP” switch. The switch is spring loaded to the center neutral position and must be
held in the “INCR” position to open the ACM bypass valve, or “DECR” to close the ACM bypass
valve. To move the ACM bypass valve from it’s fully open to fully closed position takes approx 30
seconds.

Note: The “CABIN TEMP” knob position is ignored by the temperature control module.

MAN COOL
When the mode controller is selected to “MAN COOL” the ACM bypass valve is automatically
moved to the fully closed position and the VCM is engage regardless of the ACM bypass valve
position.

Note: The “CABIN TEMP” knob position is ignored by the temperature control module and
“MAN TEMP” switch will cause the ACM bypass valve to move for as long as the switch is held,
however, the ACM bypass valve will return to the fully closed position when the switch is
released.

Temperature Gauges
To provide the flight crew with an indication of the cabin temperature a gauge is located on the left side of the
co-pilot’s subpanel. The temperature sensor, along with the cabin temperature control module are located
under the headliner, above the isle in row 6 and are covered by a perforated plastic cover.

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Air Distribution
Air Distribution Controls
The flight crew have four push-pull controls that manipulate butterfly valves to regulate the airflow to their
respective compartments. The PILOT AIR and DEFROST AIR controls are located just below and on each side of
the pilot’s yoke and the CABIN AIR and COPILOT AIR controls are located just below and on each side of the co-
pilot’s yoke.

PILOT AIR – Controls airflow from under the left side of the instrument panel near pilot‘s legs.

DEFROST AIR – Controls airflow from black eyeball vents on the top left and right side of the instrument
panel, as well as defrosting vents at the base of each front main window.

CABIN AIR – Controls the flow of air entering the cabin ducts.

COPILOT AIR – Controls airflow from under the right side of the instrument panel near co-pilot’s leg.

Air Distribution

The passengers each have an eyeball vent that allows for control of airflow rate and direction of recalculated/
VCM air to suit their individual requirements (plus one in the lavatory, if installed). The passengers have no
control over ACM air output as it enters the cabin through 6 vents in the bottom of the sidewall.

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System Operation
The environmental system uses any combination of bleed air, ACM air and VCM air to achieve the desired temperature.
Below is a general description of how each system provides different temperature air into the cabin when selected.

Heating
When a temperature is selected that requires an increase in cabin temperature (auto or manual modes), the
ACM bypass valve opens further to allow more hot air to bypass the ACM and directly enter into the cabin,
mixed with cabin air and distributed through 6 vents in the bottom of the sidewalls and into the cockpit through
a vent below each pilot seat, the two black eyeball vents on the top left and right of the instrument panel and
the defrosting vents at the base of each front main window.

Cooling (ACM Only)

The ACM provides cooling for most phases of flight and when a temperature is selected that requires a decrease
in cabin temperature (auto or manual modes), the ACM bypass valve closes further to direct more hot air into
the ACM so that it can be cooled. Once cooled, the ACM air enters the cabin, mixes with cabin air and
distributed through 6 vents in the bottom of the sidewalls and into the cockpit through a vent below each pilot
seat, the two black eyeball vents on the top left and right of the instrument panel and the defrosting vents at
the base of each front main window.

Cooling (ACM and VCM)

When a temperature is selected that is lower than the ACM can cool (ACM bypass valve fully closed) the VCM is
activated long as the inlet air temperature is greater than 40°C. The ACM will continue to cool the bleed air to its
maximum ability and enters the cabin, mixes with cabin air and distributed through 6 vents in the bottom of the
sidewalls and into the cockpit through a vent below each pilot seat, the two black eyeball vents on the top left
and right of the instrument panel and the defrosting vents at the base of each front main window. At the same
time, the VCM is operating and further cools the cabin through two evaporator coils. The VCM air mixes with
cabin air and is distributed through the silver eyeball vents at each passenger seat (plus one in the lavatory, if
installed) and above each pilot seat.

Unpressurized Ventilation
When operating unpressurized and with the cabin pressurization control switch in the “DUMP” position (or with the
aircraft on the ground as detected by the left main landing gear squat switch) a solenoid controlled door opens and
allows ambient air to be routed into the aircraft. The VENT control, located on the left side and below the co-pilot’s
subpanel, is used to regulate the amount of air entering the aircraft from the ram air scoop, located on the right lower
side of the nose of the aircraft, and into the ACM ducting. When the cabin pressurization switch is in the “PRESS”
position, the solenoid controlled door remains closed and prevents air from entering or exiting the aircraft through the
ram air scoop.

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Failures & Malfunctions
“L ENVIRO FAIL” and/or “R ENVIRO FAIL”
Emergency Procedures – “Environmental Failure”

Disagreement between switch position and valve position or valve is not powered. Valve automatically closes in
the event that the environment air duct over pressure (Activates at > 44 ± 1 PSI) and/or over temperature
(Activates at > 500˚F ± 10˚F).

3) Bleed Air Valve (affected side) – Cycle the bleed air valve switch to “ENVIRO OFF”, then back to “OPEN”

If the respective “L ENVIRO FAIL” or “R ENVIRO FAIL” annunciator extinguishes, continue operation. If
annunciator remains illuminated:

4) Bleed Air Valve (affected side) – Select the bleed air valve switch to the “ENIVRO OFF” position.

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Environmental System check
T-Test (Over-Temperature Test)
To check the control and over-temperature warning system of the environmental system, select both bleed air
valve switches to the OPEN position and verify that the respective white “L ENVIRO OFF” and “R ENVIRO OFF”
annunciators are extinguished to indicate that the bleed air valves are open. Move the environmental mode
selector to “T-TEST”, this simulates an over-temperature condition within the bleed air lines and verify that the
red “L ENVIRO FAIL” and “R ENVIRO FAIL” annunciators, as well as the white “L ENVIRO OFF” and “R ENVIRO
OFF” annunciators all illuminate. This indicates that the bleed air valve switches are in the OPEN position;
however, both bleed air valves are CLOSED. Return the environmental mode selector to “AUTO” and both bleed
air valve switches to the ENVIRO OFF position and the red “L ENVIRO FAIL” and “R ENVIRO FAIL” annunciators
should extinguish. Once again, select both bleed air valve switches to the OPEN position and verify that the
respective white “L ENVIRO OFF” and “R ENVIRO OFF” annunciators are extinguished and then return both bleed
air valve switches to the ENVIRO OFF position and verify that the respective white “L ENVIRO OFF” and “R
ENVIRO OFF” annunciators are extinguished.

SOP - T-Test

Maintenance - P-Test (Over-Pressure Test)

To check the over-pressure warning system of the environmental system, with both engines operating, select
both bleed air valve switches to the OPEN position and verify that the respective white “L ENVIRO OFF” and “R
ENVIRO OFF” annunciators are extinguished to indicate that the bleed air valves are open. Increase the engine
speed to 85% N1. The red “L ENVIRO FAIL” and “R ENVIRO FAIL” annunciators, as well as the white “L ENVIRO
OFF” and “R ENVIRO OFF” annunciators all illuminate.

Note: if the left and/or right annunciators do not illuminate within several seconds, increase the engine speed to
90% N1.

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OXYGEN SYSTEM

General
The Beechcraft 1900D is equipped with a supplemental oxygen system that can supply the crew and passengers with
oxygen during normal and/or for emergency situations. The manually activated system is designed to provide an
adequate flow of oxygen for 2 pilots and 19 passengers for approximately 60 minutes at an altitude of 25,000’.

System Components
The oxygen system consists of 2 oxygen cylinders with regulators, crew oxygen system and passenger oxygen system.
Each system component is described below.

Oxygen Cylinders and Pressure Regulators

The system stores all the available oxygen within 2 composite cylinders that are filled to nominal capacity of
77.9 ft3 (2205 litres) and pressure of 1850 PSI @ 15°C. The cylinders are located in the nose of the aircraft, below
the avionics bay and installed lengthwise. Attached to the aft end of each cylinder is a combination regulator,
shut-off valve and high pressure relief valve.

The regulated output from the regulators is up to 300 litres per minute at 70 ± 10 PSIG with cylinder pressures of
200 – 1850 PSIG. Control for the shutoff valve of each cylinder is located on the lower left side of the pilot’s
subpanel. Normal servicing for the oxygen cylinders is accomplished through the filling panel which is located on
the left side of the nose and below the avionics bay door. This single filling port will service both oxygen
cylinders and the gauge displays the average total pressure within both oxygen tanks. As well, the pressure
within each individual tank can be checked on pressure gauges in the cockpit, located on the co-pilot’s right
lower subpanel.

Cockpit Pressure Gauge

(One for each Cylinder)
Filling Panel
Oxygen Cylinders

In the event that the pressure within either oxygen cylinder exceeds 2775 PSIG @ 21.1°C (70°F), the pressure
will vent through its respective overboard discharge port, one located on each side of the lower portion of the
nose. Also, this will rupture the green indicating disc that covers the port to warn the flight crew of the
overpressure condition (the cylinder must be replaced following an overpressure condition).

Overpressure Relief Indicator

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 Crew Oxygen System
The crew oxygen system is controlled by the shutoff valve of each cylinder is located on the lower left side of the
pilot’s subpanel. Once this control knob has been pulled out, each of the crew oxygen masks is supplied the
regulated pressure and flow of oxygen and is ready for use. The crew oxygen system can be used without the
activation of the passenger oxygen system.

Crew Masks
There are 4 available crew oxygen mask types available for the aircraft. The two most common mask
types are the SCOTT-EROS and PURITAN BENNETT – SWEEP ON 2000. Both of these crew oxygen masks
are diluter demand type masks with an integrated vent valve, internal microphone, inflatable harness
and inline oxygen presence indicator. The integrated vent valve activates automatically when the crew
member is wearing the smoke protection goggles and supplies oxygen into the goggles to clear
contaminated air. Oxygen is also used to inflate the elastic type harness on the mask. To verify the
operation of the masks, an oxygen indicator is located on the hose, near the connection to the aircraft.
Both masks can be selected to operate in the following 3 modes.

1) 100% - On demand oxygen supply, 100% oxygen concentration. When the pilot inhales, the
controller releases 100% oxygen into the mouthpiece.

2) Normal – On demand oxygen supply automatically diluted. When the pilot inhales, the controller
releases diluted oxygen (mix of 100% oxygen and cabin air) into the mouthpiece.

3) Emergency – Constant Flow oxygen supply, 100% oxygen concentration. Oxygen is continuously
forced into the mouthpiece which maintains positive pressure within the mask. This prevents
contaminated air from entering the mouthpiece and protection goggles (when worn) through
the integrated vent valve.

Although the functions of both masks are identical, the appearance is different. Differences between the
two types of oxygen masks are listed below.

SCOTT-EROS PURITAN BENNETT – SWEEP ON 2000

Mode Selection Mode is selected by rotating the Mode is selected by rotating the knob
knob on the front of the mask to on the end of the mask to the required
the desired position. The setting. While wearing the mask,
positions are clearly labelled. centered is 100%, clockwise is NORM
and counter clockwise is EMERG.
Harness inflation Inflated by squeezing the red Inflated by squeezing the harness
“ears” together. Once the “ears” inflation button (red button on one side
are released the mask will deflate and black grip on the other). Once the

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 automatically. harness inflation button is released the
mask will deflate automatically.
Oxygen Presence Oxygen flow indicator. Turns Oxygen pressure indicator. Turns green
green to indicate the flow of when oxygen pressure is present in the
oxygen through the hose and red hose and red otherwise.
for no flow.

Note: Will indicate red when

oxygen is flowing into the mask
(during inhalation or when vent
valve open).

PURITAN BENNETT – SWEEP ON 2000

SCOT-EROS Mask

Passenger Oxygen System

The passenger oxygen system is controlled by the Cabin Oxygen Control knob, located is the center of, and
below pilot’s subpanel in the center. The Oxygen Control knob (see above) must first be pulled out (ON position)
to allow oxygen to reach the cabin oxygen control valve. The cabin oxygen masks are pressure regulated by an
altitude compensating regulator which automatically changes the pressure as required. The pressure schedule
ranges from approximately 21 psig (1.4 LPM) at 10,000 ft to 58 psig (2.9 LP) at 25,000 ft.

Passenger Oxygen Masks

There are 20 passenger oxygen masks (21 with optional lavatory installed) available in the passenger
cabin. Masks are located on the sides of the cabin above each seat, except for seats equipped with an
emergency exit (Seats 4B, 6A, 6B), where the masks are located immediately aft of the exit, 2 masks are
located in the ceiling (rows 6 and 8), and if a lavatory is installed, one in the ceiling within the lav. The
masks are re-breather type masks, which uses a soft plastic bag on the end of the mask to save some of
the person’s exhaled air to be used again. This reduces the amount of oxygen required by passenger
mask while maintaining a breathable mixture of air.

Deployment
When the cabin oxygen control knob is pulled out, a surge valve momentarily bypasses the altitude
compensating regulator and applies 70 psig to each of the passenger mask container assemblies. This

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 high pressure triggers the Door Opening Actuator which opens the oxygen mask container door. The
surge valve then closes and the altitude compensating regulator reduces the pressure within the lines to
the mask container assemblies to the required level. In order to allow oxygen to flow from any mask, the
lanyard pin must be pulled out of the oxygen inlet valve (this occurs automatically when the mask is
pulled out of the mask container). To stop the flow of oxygen to all of the cabin oxygen masks, the Cabin
Oxygen Control knob needs to put pushed in (OFF position). To stop the flow of oxygen to an individual
mask the lanyard pin must be inserted back into the oxygen inlet valve.

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Determining Available Oxygen
Step 1 - Calculating Oxygen Available
1) Read each cylinder pressure (PSI) and use the following formula to calculate the average:
(Cylinder 1 PSI) + (Cylinder 2) PSI
= Average PSI
2
2) Plot average PSI on the partially fill oxygen cylinder graph to determine % of oxygen available.

Step 2a – Calculating duration of oxygen available (crew and passengers using oxygen for same duration)
1) Select appropriate oxygen chart (crew on 100% or normal)
2) Using the chart, determine oxygen duration available (with full cylinders) by the total number of oxygen
masks in use (crew + passengers)
3) Compensate for partially full oxygen cylinders with the following formula:
Duration of oxygen Cylinder % (from Step 1) Actual duration available
from chart (minutes)
x 100
= (minutes)

Step 2b – Calculating duration of oxygen available (crew and passengers using oxygen for different durations)
1) Select appropriate oxygen chart (crew on 100% or normal)
2) Compensate for partially full oxygen cylinders with the following formula:
Duration of Crew Oxygen Cylinder % (from Step 1) Actual Crew Duration
from Chart (minutes)
x 100
= Available (minutes)

3) Determine duration of oxygen required (flight requirements)

4) Calculate the remaining oxygen available with the following formula:

Actual Crew Duration Duration Required Remaining Oxygen

Available (minutes)
- for Crew (minutes)
= Available (minutes)

5) Calculate the litre of oxygen available with the following formula:

Remaining Oxygen Crew Mask Oxygen Remaining Oxygen
Available (minutes)
x Flow Rate (LPM)
= Available (Litres)

6) Calculate the duration of oxygen available for passengers with the following formula:

Remaining Oxygen Available (Litres) Duration of Passenger Oxygen

=
Passengers Mask (LPM) x # of Passengers Available (minutes)

Effective Performance Time (EPT) or Time of Useful Conscientious (TUC)

The EPT/TUC is the time interval between exposure to an oxygen deficient environment and loss of useful mental
function. Typical EPT/TUC times are shown in the chart below; however, these times may be reduced do to other
factors.
Cabin Altitude Typical EPT / TUC
15,000 30 Minutes or More
18,000 20 - 30 Minutes
22,000 5 - 10 Minutes
25,000 3 - 5 Minutes
28,000 2.5 - 3 Minutes
30,000 1 - 3 Minutes
35,000 30 - 60 Seconds
40,000 15 - 20 Seconds
45,000 9 - 15 Seconds
50,000 6 - 9 Seconds

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Oxygen Use - Standard Operating Procedures (SOP’s)
Smoke or Fume Elimination
If smoke is prevalent in the cabin, cabin oxygen masks should not be deployed unless the cabin altitude exceeds
15,000 ft, and should only be used only until the cabin altitude is reduced to 15,000 ft or lower.

Note: As re-breather type masks recycles air that the passenger breathes, use of oxygen masks should be
carefully considered whenever environmental contamination exists. Also, in the event of a smoke or fumes, the
use of cabin oxygen may increase oxygen levels within the cabin and increase oxygen supply to a potential fire.

Cabin Decompression
The use of cabin oxygen masks should not be used unless the crew is unable to correct the problem and the
cabin altitude approaches 15,000 ft or if decompression is rapid.

System Checks
Oxygen System

The oxygen system is checked to ensure proper operation of the controls, crew masks and that sufficient oxygen
is present.

Ensure that the CABIN OXYGEN knob is pushed in fully, and then pull the OXYGEN CONTROL knob ON. Press then
release the harness activation button crew both oxygen masks and ensure that the harness inflates then
deflates. Select the NORM, 100% and EMERG modes on each harness and check for freedom of movement and
that air rushes out of the mask when the EMERG mode is selected. Reset the masks to the centered 100% mode.
Next, verify that both the pilot and co-pilot mic switches on the overhead panel are selected to normal. Lastly,
check the average quantity of oxygen displayed on the oxygen quantity gauges is sufficient for the intended
flight.

SOP – Oxygen System Check

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Oxygen Duration Charts
SCOTT-EROS & PURITAN BENNETT-SWEEP ON 2000 – Normal Mode

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PURITAN BENNETT-SWEEP ON 2000 (Inflatable Harness) - 100% Mode

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SCOTT-EROS (Inflatable Harness) - 100% Mode

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Determining Available Oxygen Worksheet
Step 1 - Calculating Oxygen Available
1) Read each cylinder pressure (PSI) and use the following formula to calculate the average:
(Cylinder 1 PSI) + (Cylinder 2) PSI
= Average PSI
2
2) Plot average PSI on the partially filled oxygen cylinder graph to determine % of oxygen available.

Step 2a – Calculating duration of oxygen available (crew and passengers using oxygen for same duration)
1) Select appropriate oxygen chart (crew on 100% or normal)
2) Using the chart, determine oxygen duration available (with full cylinders) by the total number of oxygen
masks in use (crew + passengers)
3) Compensate for partially full oxygen cylinders with the following formula:

Duration of oxygen Cylinder % (from Step 1) Actual duration available

from chart (minutes)
x 100
= (minutes)

Step 2b – Calculating duration of oxygen available (crew and passengers using oxygen for different durations)
1) Select appropriate oxygen chart (crew on 100% or normal)
2) Compensate for partially full oxygen cylinders with the following formula:

Duration of Crew Oxygen Cylinder % (from Step 1) Actual Crew Duration

from Chart (minutes)
x 100
= Available (minutes)

3) Determine duration of oxygen required (flight requirements)

4) Calculate the remaining oxygen available with the following formula:

Actual Crew Duration Duration Required Remaining Oxygen

Available (minutes)
- for Crew (minutes)
= Available (minutes)

5) Calculate the litre of oxygen available with the following formula:

Remaining Oxygen Crew Mask Oxygen Remaining Oxygen

Available (minutes)
x Flow Rate (LPM)
= Available (Litres)

6) Calculate the duration of oxygen available for passengers with the following formula:

Remaining Oxygen Available (Litres) Duration of Passenger Oxygen

=
Passengers Mask (LPM) x # of Passengers Available (minutes)

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PNEUMATICS

General Description
Pneumatic air (also known as service air) is supplied from both engines and used in several aircraft systems. These
systems include: surface deice system, brake deice and bleed air overheat warning systems, vacuum ejector, flight hour
meter and pressurizing of the hydraulic reservoir.

Pneumatic Air Source

The compressor section of both engines are used to produce air for the pneumatic systems, however, one engine is
capable of supplying sufficient bleed air to supply all the systems (with N1 speeds > 70-80%). P3 air is routed through the
Pneumatic Bleed Air Shutoff Valve, which is located in the leading edge of each wing between the fuselage and engine
nacelle, and is controlled by the Bleed Air Control switch, located on the left side of the co-pilot’s subpanel. This switch
has 3 positions: (This switch also controls the environmental system, see environmental system chapter for details)

Bleed Air Valves

OPEN – Pneumatic Bleed Air Shutoff Valve is OPEN (unpowered)

ENVIRO OFF – Pneumatic Bleed Air Shutoff Valve is OPEN (unpowered)

INST & ENVIRO OFF - Pneumatic Bleed Air Shutoff Valve is CLOSED (powered)

With the Pneumatic Bleed Air Shutoff Valve OPEN, P3 air passes through a one way check valve and into a pressure
regulator. This regulator reduces the P3 air to a nominal 18 psi (17.5 – 19.0 psi) and is then distributed to the various
pneumatic systems (See diagram below). The 18 psi regulator contains a pressure relief valve that will prevent the
pneumatic pressure from exceeding 21 ± 1 psi.

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Pneumatic Air Systems
Pressure Systems
The regulated pneumatic pressure can be monitored on the Pneumatic Pressure Gauge, located on the right side
of the co-pilot’s subpanel.

UE-21 and After

17.5-19.0 PSI Normal Operating

Range
----- 20 PSI Maximum Limit

Bleed Air and Brake Deice Warning Systems

Pneumatic air is used to pressurize 4 small EVA (ethylene vinyl acetate) tubes (1 for each bleed air line
and 1 for each brake deice line). This pressurized air is used by a sensor within each line to monitor the
EVA tubing for leaks. Should the pressure in that EVA tube drop below 1.5 ± 0.5 PSI it will cause its
respective annunciator will illuminate: “L BL AIR FAIL”, “R BL AIR FAIL”, “L BK DI OVHT”, and “R BK DI
OVHT”.

Note: Any time there is less than 1.5 ± 0.5 PSI in the EVA tube its respective annunciator will illuminate.
For example: If both pneumatic sources were turned off, all 4 of the annunciators would illuminate.

(See Fire Protection Chapter for Bleed Air/Brake Deice Warning System details)

Landing Gear Hydraulic Reservoir

Pneumatic air is used to pressurize the landing gear hydraulic reservoir to provide positive pressure on
the hydraulic fluid to prevent gas bubbles from forming within the flowing fluid which would result in
hydraulic pump cavitations as well as to prevent the hydraulic fluid from foaming on the surface. Pump
cavitations cause shockwaves to develop within the system and can significantly damage moving parts.

(See Landing Gear Chapter for Landing Gear Reservoir details)

Flight Hour Meter

Pneumatic air is used to trigger a pressure switch in series with the right landing gear squat switch to
activate the flight hour meter (records “air time”), located on right side of the co-pilot’s subpanel.

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 Surface Deice System (Inflation)
Pneumatic air is used to inflate the surface deice boots during the inflation cycle. The pressurized
pneumatic air routed into the surface deice boots through the surface deice distribution valve.

(See Ice Protection Chapter for Surface Deice System details)

Vacuum Ejector
Pneumatic air is routed through an ejector and then overboard to ambient air; this ejector utilizes a
venturi to create a vacuum within the vacuum system lines. A regulator limits the amount of vacuum
within the system to 4.3 - 5.9 in HG (except for the surface deice system). Each component that uses the
regulated vacuum system has its own air filter to prevent contaminants from entering into the system.

Pressurization Control
The pressurization control system uses vacuum pressure to operate the pressurization controller
and outflow valves. The pressurization control system has its own vacuum regulator which
further reduces the vacuum to 3.75 - 4.75 in HG.

(See Pressurization Chapter for Pressurization Control details)

Surface Deice System (Deflation)

Unregulated vacuum is used by the surface deice system during the deflation cycle and to
maintain the boots deflated when not in use (engines operating). Vacuum is applied through the
surface deice distribution valve to the deice boots.

(See Ice Protection Chapter for Surface Deice System details)

Malfunctions & Faults

“L BL AIR FAIL” and “R BL AIR FAIL” and “L BK DI OVHT” and “R BK DI OVHT”
No Checklist - Most likely indicates a failure of the pneumatic 18 psi regulator. However, checking the pneumatic
pressure gauge and vacuum gauge will confirm the problem.

If the pneumatic pressure gauge is normal and the vacuum gauge indicates abnormally low or zero:

Abnormal Procedures Checklist – Vacuum Source Failure

1) Failure of the vacuum source will be indicated by:

a. Abnormally low or zero indication on the gyro suction gage/vacuum gage.
b. An uncontrolled rate-of-descent until maximum cabin differential is reached.

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 2) Both Bleed Air Valves – Enviro Off
(as required to decrease cabin differential pressure to zero prior to landing)

Although not in the aircraft checklists, if the pneumatic pressure gauge and vacuum gauge indicate abnormally
low or zero, the above remains true as well as all of the pneumatic system will be inoperative.

Important Inoperative Systems:

- Surface Deice System

- Bleed Air/Brake Deice Warning Systems
- Pressurization Control (uncontrolled rate-of-descent to max differential)

System Checks
Pneumatic System
The pneumatic system is checked to insure proper supply and indications from the system.

With both engines operating and both BLEED AIR switches in the ENVIRO OFF position, check the gyro suction
gauge indicates in the green range (4.3 - 5.9 HG) and pneumatic pressure gauge indicates in the green range
(17.5 - 20 PSI). Select the left BLEED AIR VALVE to INST & ENVIRO OFF and check the gyro suction and pneumatic
pressure gauges remain in the green range. Select the right BLEED AIR VALVE to INST & ENVIRO OFF and check
that the gyro suction and pneumatic pressure gauges indicate 0 and the yellow “L BL AIR FAIL”, “R BL AIR FAIL”,
“L BK DI OVHT” and “R BK DI OVHT” annunciators illuminate. Select the left BLEED AIR VALVE to OPEN and verify
that the both gyro suction and pneumatic pressure gauges return to the green range and the respective
annunciators extinguish. Select the right BLEED AIR VALVE to OPEN.

SOP – Pneumatic System Check

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LANDING GEAR & BRAKES

General Description
The Beechcraft 1900D is equipped with a retractable, tricycle landing gear system that is electronically controlled and
hydraulically powered. The landing gear system also incorporates an indicating and warning system along with an
alternate gear extension procedure. Each main landing gear is equipped with conventional hydraulic brakes and optional
anti-skid system (if installed).

Landing Gear
The landing gear system consists of 1 nose gear assembly, 2 main gear assemblies, hydraulic power pack, landing gear
controls, alternate landing gear alternate extension system and landing gear position/warning system.

Nose Landing Gear

The nose landing gear assembly consists of an air-oil type shock strut, Torque link, drag brace, hydraulic
actuator, wheel, tire, steering linkage and integrated gear door with taxi light.

The air-oil type shock utilizes compressed air and red coloured MIL-5202 hydraulic fluid to dampen landing and
taxi forces. The torque link allows the lower gear assembly to move independently of the upper assembly while
maintaining the correct alignment. The drag brace secures the nose landing gear in the down and locked
position by extending slightly over-center and also collapses inwards to allow the hydraulic actuator to retract
the nose landing gear. The single nose landing gear tire is a 19.50 x 6.75 x 8, 10-ply, 190 mph rated, tubeless tire
and is mounted on a 6.50 x 8 magnesium wheel. The main landing gear shock strut and tire is properly inflated
when 5.25” - 5.75” of the shiny shock strut are visible and the tire pressure is 60 +5/-0 psi.

The nose landing gear bay door is connected the fuselage by a hinge at the top and a link assembly connects the
bottom of the door to the nose landing gear upper assembly. As the nose landing gear door extends or retracts
the nose landing gear bay door pivots with the gear to the required position. Also, within the nose landing gear
bay door is an integrated window and taxi light.

Steering (No Power Steering)

A direct mechanical linkage to the rudder pedals provides control of the nose landing gear for steering when the
aircraft is on the ground and allows the nose wheel to be turned 15° left or right. A spring mechanism in the
linkage dampens shock loads from reaching the rudder pedals and permits turning angles of 63° left or right
while being towed or during hard turns (brake and asymmetric power assited).

As the rudder pedals control the nose landing gear steering and rudder, once the left main landing gear squat
switch determines the aircraft is airborne, the steering disconnect actuator extends and moves the steering
linkage to the bottom of the steering cam which disconnects the nose wheel steering and allows the rudder
pedals to only control the rudder pedals. Once the aircraft lands the steering disconnect actuator retracts and
moves the steering linkage to the top of the steering cam to restore nose wheel steering. The nose gear also is
equipped with steering stop on the bottom-rear of the upper assembly that prevents damage to the steering
system.

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 Main Landing Gear
The two main landing gear assemblies consists of an air-oil type shock strut, Torque link, drag brace, hydraulic
actuator, wheels, tires, brakes and landing gear door actuator cams.

The air-oil type shock utilizes compressed air and red coloured MIL-5202 hydraulic fluid to dampen landing and
taxi forces. The torque link allows the lower gear assembly to move independently of the upper assembly while
maintaining the correct alignment. The drag brace secures the nose landing gear in the down and locked
position by extending slightly over-center and also collapses inwards to allow the hydraulic actuator to retract
the nose landing gear. The dual main landing gear tires are 22.00 x 6.75 x 10, 10-ply, 190 mph rated, tubeless
tires and are mounted two 6.50 x 10 magnesium wheel. The main landing gear shock strut and tire is properly
inflated when 5.12” - 5.6” of the shiny shock strut are visible and the tire pressure is 97 +5/-0 psi (or 93 +5/-0
when on jacks).

Each of the two main landing gear bay doors is connected at the top to the engine nacelle by a hinge and held to
the fully open position by a spring. Actuator pins on the main landing gear upper assembly are aligned with the
gear bay door actuating cam and close the doors as the respective landing gear is retracted into the gear bay.

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 Brake System
The brake system consists of a hydraulic reservoir, dual brakes, parking brake and disc brakes installed
on each main landing gear assembly.

The hydraulic reservoir is located at the rear side of the right avionics bay and contains a sight glass on
the side for checking the fluid level and filling cap on the top for servicing. The brake system utilizes
standard red Mil-5202 hydraulic fluid. The brake fluid first flows into the pilot’s set of toe brake master
cylinders, then into the co-pilot toe brake master cylinders. Either set of pilot cylinders can be used to
activate the brakes and the pressure applied between the cylinders is cumulative.

Application any pair of brakes (left or right) will apply force fluid into 5 pistons on the respective brake
assembly and creates friction between the stack of rotating discs and stationary brake pads within the
brake assembly.

A parking brake is installed between the brake pedals and brake assembly. The parking brake control is
located on the aft-right portion of the center pedestal and contains an integrated lock within the two
piece handle.

The proper procedure to apply the parking brake, apply the required brake pressure by using the tow
brakes then compress the lock on the parking brake handle by squeezing the lower and upper discs
together, raise it fully and release the lock. Brake pressure is now locked within the brake lines and
additional pressure can be added through a one way check valve within the parking brake assembly. To
release the brakes, apply pressure to the toe brakes to equalize pressure on both sides of the parking
brake, compress the lock on the parking brake handle, lower it fully and release the lock.

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 Landing Gear Controls
The landing gear controls are located on the right side of the pilot’s subpanel and consist of the landing gear
control handle and handle test switch, J-hook release, gear position indicators, landing gear relay circuit breaker
and gear warning horn silence button.

The landing gear control handle is used to select the landing gear UP or DOWN and contains 2 internal red light
bulbs that are used to indicate the gear is in transit (not up or down and locked). A momentary switch is located
immediately to the right of the landing gear control handle and activates the red lights within the handle for
testing. When the landing gear control handle is selected UP or DOWN its respective solenoid moves the
selector valve to the required position.

A landing gear safety switch or J-hook is a solenoid operated mechanical hook that prevents the landing gear
handle from being moved from the down position while the aircraft is on the ground and automatically unlocks
the landing gear control handle when the aircraft is airborne as indicated by the right squat switch. Should the J-
hook fail to unlock a manual switch labelled DN LCK REL, located immediately to the left of the landing gear
handle is used to move the J-hook to the unlock position.

The gear position indicator block contains 3 annunciator segments that have 2 light bulbs each. The horizontal
NOSE annunciator is a typical annunciator that is illuminated by either of the two bulbs. However, the two
annunciators for the main landing gear are mounted horizontally and function vertically between two
annunciators. The L of the center annunciator and H (left side) of the bottom annunciator indicate the position
of the LEFT main landing gear and the R of the center annunciator and H (right side) of the bottom annunciator
indicate the position of the RIGH landing gear. These annunciators contain a light divider between the L and R
(center annunciator) and also the H and H (bottom annunciator) and cannot be exchanged with any other
annunciators as the result may provide false indications depending of the fault.

Landing Gear Lights

The landing gear circuit breaker is used to protect and disable the control circuitry of the hydraulic power pack
(see Hydraulic Power Pack below for details).

Also located on the landing gear control panel is the WARN HORN SILENCE that is used to silence the landing
gear warning horn when activated by the power lever switches (low power configuration warning). However,
this switch will NOT silence the warning horn if activated by the flap system switches.

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 Hydraulic Power Pack
The hydraulic power pack consists of an electric motor, hydraulic pump, hydraulic reservoir, gear selector
solenoids and a pressure switch. The power pack is located in the left wing, between the fuselage and engine
nacelle and can be serviced through an access door on the top of the wing.

The electric motor is controlled by the landing gear control circuitry and is used to power the hydraulic pump
within the power pack. The hydraulic pump forces hydraulic oil from the reservoir and through the gear selector
solenoid and to the desired port of each landing gear accumulator.

The hydraulic reservoir contains a liquid level sensor that illuminates the yellow “HYD FLUID LOW” to alert the
crew that a low quantity of hydraulic fluid remains in the reservoir. Also, the hydraulic reservoir is designed with
a separate supply lines at different levels for the power pack hydraulic pump and alternate landing gear
extension pump so that in the event of a leak within the power pack hydraulic system, enough hydraulic fluid
remains for the alternate landing gear extension pump be able to fully extend the landing gear. Pneumatic air is
used to pressurize the landing gear hydraulic reservoir to provide positive pressure on the hydraulic fluid to
prevent gas bubbles from forming within the flowing fluid which would result in hydraulic pump cavitations as
well as to prevent the hydraulic fluid from foaming on the surface. Pump cavitations cause shockwaves to
develop within the system and can significantly damage moving parts.

A pressure switch is used in addition to the landing gear control handle to control the power pack during gear
retraction and maintaining the gear retracted, while microswitches within each hydraulic actuator control the
power pack during the gear extension. To assist the power pack in maintaining the landing gear in the retracted
position a hydraulic accumulator, located within the left landing gear bay on the outboard wall, is charged to
800 ± 50 PSI and maintains the hydraulic pressure within the system to reduce operating cycles of the power
pack. Any time the power pack is operating a timing circuit is activated that will trip the landing gear relay circuit
breaker, located on the right side of the pilot’s subpanel, anytime the power pack operates for greater than 16 ±
0.5 seconds.

(See Landing Gear Operation below for details)

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 Alternate Landing Gear Extension Handle
The alternate landing gear extension handle is located on the cockpit floor between the left pilot’s chair and the
center pedestal. This handle allows the flight crew to extend the landing gear manually when an emergency gear
extension is required.

During normal operations the alternate landing gear extension handle is stowed lowered fully and under the
securing clip. When its use is required, the alternate landing gear extension handle is used to bypass the
hydraulic power pack motor and gear selector solenoid and provide sufficient hydraulic fluid and pressure to the
extension ports on the landing gear actuators to fully extend the gear.

(See Landing Gear Operation below for details)

Note: when the alternate landing gear extension handle is lowered below the height of the securing clip an
internal pressure relief valve is actuated to relieve pressure from within the alternate landing gear extension
pump.

Landing Gear Position/Warning System

The landing gear position/warning system consists of multiple switches and secondary systems to provide
accurate gear position, troubleshooting and preventative warning information to the flight crew.

Inputs on the landing gear for the landing gear portion of the landing gear position/warning system are provided
through 3 switches on each landing gear assembly. One switch is internal to each hydraulic actuator and
activates when the actuator is in its locked in its fully extended position, one switch is mounted each drag brace
and activated when the drag brace is in its over-center down position, and the last one is mounted inside the
gear bay and actuates when the nose landing gear is fully retracted.

Other inputs for the landing gear position/warning system are switches on each power lever that actuate (when
the gear is not down and locked) when the respective engine speed decreases below 86-84% N1, two switches
within the flap system that activate (when the gear is not down and locked) when the flaps are either selected
to 35° or the actual flap position is greater than 17°. The WARN HORN SILENCE button ONLY will silence the gear
warning horn when activated by either power lever switch and the warning is reset when the power lever is
increased beyond the set point. If the flaps are selected, or are beyond 17° and the gear is not down and locked,
or the landing gear control handle is in the UP position with the aircraft on the ground the warning horn will
activate and cannot be silenced by using the WARN HORN SILENCE button.

Also, when on the ground the right squat switch removes power from the control circuitry and from the landing
gear handle J hook to prevent raising the landing gear control handle with the aircraft on the ground.

Additionally, the landing gear power pack contains a timing circuit that trips the landing gear relay circuit
breaker, located on the right side of the pilot’s subpanel, anytime the power pack operates for greater than 16 ±
0.5 seconds.

Landing Gear Operation

Retraction
Once the aircraft is airborne (or weight off wheels) the steering actuator extends and the nose gear
centers automatically. Once the landing gear control handle is moved to the UP position, the hydraulic
power pack is activated and the gear up solenoid moves the gear selector valve to allow hydraulic fluid
under pressure into the retract side of the system from the primary port on the reservoir, and fluid from
the extend side of the system to return to the reservoir. The landing gear control handle will illuminate
red to indicate the landing gear is in transit (not in the selected position).

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 Once the hydraulic pressure increases to 200-400 PSI the internal locking mechanism within each
landing gear actuator unlocks and the landing gear will start to retract. The main landing gear retracts
forward and the nose landing gear retracts rearward into their respective wheel wells. As hydraulic fluid
flows into the retract side of the system the landing gear will continue to retract, at the same time, fluid
from the extend side of the system returns through a separate port within the gear selector valve to the
reservoir. Each landing gear green position indicator lights will extinguish once its respective actuator
pins have unlocked.

Once the pressure within the retract side of the system reaches 2,775 PSI the power pack is disabled and
the pressure is held within the retract side of the system to maintain the gear in the retracted position.
The landing gear control handle red light will extinguish once each landing gear wheel well switch is
activated and indicates that the landing gear is fully retracted.

Note: The landing gear power pack contains a timing circuit that trips the landing gear relay
circuit breaker, located on the right side of the pilot’s subpanel, anytime the power pack
operates for greater than 16 ± 0.5 seconds.

As hydraulic pressure will slowly decrease within the retract side of the system, the power pack is
activated again once the pressure drops to 2,320 PSI to increases the pressure back to 2,775 PSI. To
assist the power pack in maintaining the landing gear in the retracted position a hydraulic accumulator,
located within the left landing gear bay on the outboard wall, is charged to 800 ± 50 PSI and maintains
the hydraulic pressure within the system to reduce operating cycles of the power pack.

Normal Extension
Once the landing gear control handle is moved to the DOWN position, the hydraulic power pack is
activated and the gear down solenoid moves the gear selector valve to allow hydraulic fluid under
pressure into the extend side of the system from the primary port on the reservoir, and fluid from the
retract side of the system to return to the reservoir. The landing gear control handle will illuminate red
to indicate the landing gear is in transit (not in the selected position).

As hydraulic fluid flows into the extend side of the system the landing gear will begin to extend, at the
same time, fluid from the retract side of the system returns through a separate port within the gear
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 selector valve to the reservoir. The main landing gear extends rearward and the nose landing gear
retracts forward from their respective wheel wells.

Once each of the landing gear actuators has reached their fully extended position the internal locking
mechanism pin engages into the actuator piston to lock the respective landing gear in place and
activates the actuator pin switch. Once all 3 actuator pins switches are activated, the hydraulic power
pack is disabled.

Note: The landing gear power pack contains a timing circuit that trips the landing gear relay
circuit breaker, located on the right side of the pilot’s subpanel, anytime the power pack
operates for greater than 16 ± 0.5 seconds.

In addition, when the actuator is fully extended, the drag brace on each landing gear will be extend
slightly over-center, which also locks the gear down and will activate the respective drag brace switch.
Once all 3 drag brace switches are activated, the landing gear control handle red light will extinguish.

With the actuator pin switch and drag brace both activated on a single landing gear, the respective
green gear position indicator light will illuminate.

(See Landing Gear Position Troubleshooting below for details)

Alternate Landing Gear Extension Handle

Note: Use of the alternate landing gear extension handle should be in accordance with the aircraft’s
emergency checklist procedure.

When used, the alternate landing gear extension handle is pumped by the left seat pilot in order to
lower the landing gear without using the hydraulic power pack or gear selector valve. Hydraulic fluid is
drawn from secondary port on the reservoir, enters the manual pump and is routed to the extend side
of the system through the emergency extension lines. The pressure within the emergency extension
lines position shuttle valves within each actuator to allow fluid to enter. As the handle raised and
lowered, hydraulic fluid and pressure will lower the landing gear and the 3 green gear position indicator
annunciators will illuminate once the gear is down and locked.

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 Note: When the alternate landing gear extension handle is lowered below the height of the
securing clip an internal pressure relief valve is actuated to relieve pressure from within the
alternate landing gear extension pump.

Note: After an EMERGENCY landing gear extension has been made, do not move any landing
gear controls or reset and switches or circuit breakers until the malfunction has been
determined and corrected.

Landing Gear Troubleshooting

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Landing Gear Switch Items
Nose Landing Gear
Actuator
- Gear extended signal for landing gear NOSE position annunciator
- Gear extended signal for power pack relay
Drag Brace
- Gear extended signal for landing gear control handle in-transit light
- Antiskid system pressure switch control
Uplock
- Gear retracted signal for landing gear control handle in-transit light

Left Main Landing Gear

Actuator
- Gear extended signal for landing gear L H position annunciators
- Gear extended signal for power pack relay
- Gear selector valve power
Squat Switch
- Stall transducer heat mode control
- Nosewheel steering disconnect actuator control
- Pressurization system preset and dump solenoids
- Flight hour meter control
Drag Brace
- Brake deice timing circuitry
- Gear extended signal for landing gear control handle in-transit light
Uplock
- Gear retracted signal for landing gear control handle in-transit light

Right Main Landing Gear

Actuator
- Gear extended signal for landing gear R H position annunciators
Squat Switch
- Ground idle signal for low pitch stop solenoid system
- Gear extended signal for power pack relay
- Gear selector valve power
- Landing gear handle J-hook solenoid power
Drag Brace
- Gear extended signal for landing gear control handle in-transit light
- Air conditioner system condenser fan power
Uplock
- Gear retracted signal for landing gear control handle in-transit light

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System Checks
Landing Gear Control and Indication
The landing gear control and indication system is checked to ensure proper operation and indication of the
various components.

Press down on the landing gear down lock release lever and verify that the J-hook moves away from the top of
the landing gear handle arm. Check that the landing gear control handle is in the down position, then press the
HDL LT TEST button and verify that both red lights illuminate within the landing gear control handle. Verify that
the green “NOSE”, “L H” and “R H” annunciators are all illuminated. Check the WARN HORN SILENCE button for
freedom of movement and that the and LANDING GEAR RELAY circuit breaker is IN. Verify that the alternate
landing gear extension handle is against the cockpit floor and stowed under the handle clip.

SOP – Landing Gear Control Check

Limitations
Airspeeds

Maximum Landing Gear Operating Speed (VLO)

Extension 180
Retraction 180

Landing Gear & Tire Inspection

Shock Strut Inflation Tire Pressure

Nose Landing Gear 5.25” - 5.75” 60 +5/-0 psi
Main Landing gear 5.12” - 5.6” 97 +5/-0 psi (or 93 +5/-
0 when on jacks)

Landing Gear Hydraulic Accumulator

Accumulator Charge 800 ± 50 PSI

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FLIGHT CONTROLS

General Description
All the primary flight controls on the Beechcraft 1900D, with exception of the flaps, are cable operated convention
surfaces. The flap system is electrically powered, as well as the following supplemental systems: electric elevator trim,
yaw damp and autopilot (if installed).

Primary Controls
The primary flight controls are manual controlled through a cable-pulley-bellcrank system and can be controlled through
either of the dual controls by the pilot or co-pilot. The ailerons, elevators and rudder can be secured with control locks
to prevent damage during windy conditions. The controls locks are installed on the pilot’s yoke and consists of a pin that
in inserted through the floorboard to lock the rudder, another pin that is inserted into the right side of the pilot’s yoke
(with the yoke full forward and rotated approximately 15° counter-clockwise) to lock the ailerons and elevator and a U-
shaped clamp that fits around the engine power control levers to warn the flight crew not to start the engines with any
portion of the controls locks installed.

Ailerons
One aileron is located on the trailing edge of each wing immediately outboard form the outboard flap segments.
The ailerons are designed as frise-differential type ailerons to minimize adverse yaw during their use.

Frise Ailerons – A portion of the forward edge of the aileron will protrude into the airflow of the down-
going wing to increase drag on the respective wing.

Differential Ailerons – The ailerons deflection is larger up and smaller down. This increases drag when
deflected up and decrease drag when deflected down.

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 Aileron Trim
The aileron system has two trim tabs located on the left aileron, a fixed trim tab that can only be
adjusted by maintenance on the ground, and a pilot adjustable trim tab that is operated through a
drum-cable system attached to jackscrews and controlled by a trim knob on the left, lower side of the
center pedestal that rotates clockwise and counter clockwise. The aileron trim knob has an integral trim
position indicator that has a scale of 6 units left or right.

Aileron Trim
Elevator
The elevator consists of 2 separate elevators jointed by a torque tube and is a convention T-tail design with the
elevators located on the top of the vertical stabilizer and aft of the vertical stabilizer.

Elevator Trim (Manual)

The elevator trim system had one pilot adjustable trim tab on each elevator that is operated through a
drum-cable system attached to jackscrews and controlled by a trim knob on the left side of the center
pedestal that rotates forward and aft. The elevator trim knob has an integral trim position indicator that
has a scale of 3 units nose down and 10 units nose up with an additional takeoff range indentified by the
green portion between 0 units (neutral) and 6 units nose up.

Elevator Trim

Elevator Trim (Electric)

The optional electric elevator trim system is controlled by the pilot or co-pilot through a split switch
located on the outboard side of each yoke. Should the pilot and co-pilot activate the trim system at the
simultaneously, the pilot’s input will override the co-pilots input. The elevator trim system has a master
ON/OFF switch located on the left-rear portion of the lower pedestal and is powered through a 5 amp
“PITCH TRIM” circuit breaker on the co-pilot’s circuit breaker panel.

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 In the event of a trim runaway situation, the red disconnect switch on either yoke can be pressed to its
2nd level (fully depressed) to disable the electric trim system and the yellow ‘PITCH TRIM” annunciator
will illuminate. To reset the system, the master switch must be turned off then on again.

Rudder
The rudder is a single piece construction and mounted behind the vertical stabilizer and is the only primary flight
control surface that is equipped with a control assist system.

Rudder Trim
The rudder has a single pilot adjustable trim tab on the lower portion of the rudder. The trim tab is
operated through a drum-cable system attached to jackscrews and controlled by a trim knob on the
right, lower side of the center pedestal that rotates clockwise and counter clockwise. The rudder trim
knob has an integral trim position indicator that has a scale of 6 units left or right.

Rudder Trim

Flight Control Assist Systems

Yaw Damp (Non-Autopilot Aircraft)

The yaw damp system is installed to assist the pilot in maintaining directional control during
normal operations. The yaw damp system senses changes in heading through a yaw rate sensor
and dual yaw damp computers (Left and Right computers) located in the in the avionics bay and
controls an electric rudder servo to counteract this adverse yaw.

The yaw damp system is controlled through a YAW momentary switch located on the left-rear
portion of the lower pedestal. To disconnect the yaw damp system, the YAW switch can be
pressed or the red disconnect switch on either yoke can be pressed to its 1st level to disable the
yaw damp system.

Rudder Boost (Non-Autopilot)

The rudder boost system is installed to assist the pilot in maintain directional control during an
engine failure or power loss. The rudder boost system senses engine torque from sensors on the
torque transmitters of each engine. The system constantly compares the torque levels and
should the difference exceed 1000 ft-lbs, the electric rudder servo activates proportionally to
the torque difference and deflects the rudder to assist the pilot.

The rudder boost system is ON as long as the rudder boost control switch on the left-rear
portion of the center pedestal is in the RUDDER BOOST position. If this switch not in this
position, the yellow “RUD BOOST OFF” annunciator illuminates to alert the flight crew that the
rudder boost system is disabled. Also, the red disconnect switch of either yoke can be held to its
1st level to interrupt the rudder boost system.

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 Yaw Damp & Rudder Boost (Autopilot Aircraft)
For aircraft equipped with a Hawker-Beechcraft installed autopilot, the autopilot incorporates
both the yaw damp and rudder boost capabilities (see “Miscellaneous Systems and/or
supplements” chapter for autopilot supplement).

Flaps
The flap system is a slotted, fowler type flap type and consists of four flap segments, one outboard and inboard
on each wing, and is driven by an electric motor attached to a gearbox.

Slotted – A slot (gap) between the wing and flap segment allows high pressure air from under the wing
to flow over the top of the flap segment and maintain the boundary layer over the flap to increase its
efficiency.

Fowler – The flap segment moves rearward as it moves downward. This effectively increases surface
area of the wing (by increasing the effective chord and camber of the wing) and provides better low
speed handling characteristics.

The motor incorporates a magnetic clutch (two sets of motor windings) which helps prevent the flaps from
traveling beyond the selected position. Through a gearbox the motor drives 4 flexible driveshafts, each
connected to a jackscrew for each flap segment.

The flap position is controlled by a sliding lever, located on the right side of the center pedestal, below the
condition levers. The flap control lever can select the flaps to UP, 17° and 35° and the current flap position is
displayed on the flap position indicator, located on the left side of the center pedestal, above the power levers.

Flap Selector / indicator

Two safety systems are incorporated into the flap system, the first is an aural warning that sounds whenever the
flaps are selected to 35° and the landing gear is NOT down and locked. The warning horn cannot be silenced by
pressing the landing gear warning silence button and will only stop once the gear is down and locked or the flaps
are retracted to 17° or less.

The second safety system is a split flap protection system. This system compares the position of the inboard flap
to the outboard flap on each wing (left wing segments are not compared to the right wing segments). Should the
position of a pair of flap segments exceed 3° - 6°, power to the flap motor is disabled and the flaps will remain at
their current position until the aircraft is serviced. So that the flaps can be retract, the flap system deferred and
the aircraft be returned to service (operating flapless) an override switch is located at the top of the avionics bay
and accessed through the left avionics bay door. With the flap selection lever set to the flaps UP position,
holding the override switch either left or right (depends on which pair of flaps activated the split flap protection)
will bypass the split flap protection and allow the flaps to be retracted.

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Stall Warning System
The stall warning system consists of a lift transducer, lift computer and aural warning system. The lift transducer is
located on the left wing between the engine nacelle and wingtip and provides information to the lift computer. The lift
transducer is also heated for flight into icing (see “Ice Protection” chapter for details). The lift computer processes the
information from the lift transducer and compensates for changes in aircraft attitude, airspeed, flap position and gear
position to activate the aural warning prior to alert the crew of an impending stall condition. The stall warning system is
disabled by the left main landing gear squat switch when the aircraft is on the ground.

A stall warning test switch is located on the left-bottom side of the co-pilot’s subpanel and bypasses the squat switch in
order to test the aural warning portion of the system only.

Limitations
Flap Limitations

Speed (At 17,120 lbs) KIAS Remarks

Maximum Flap Do not extend flaps or operate with
Extension/Extended Speed 188 flaps in the prescribed position above
(VFE) 154 these speeds.
Flaps 17
Flaps 35

Gust Locks

The gust locks are used in order to prevent damage to the primary controls by preventing any movement of the
controls, cables and respective control surfaces. The control locks must be removed before towing the aircraft
to avoid serious damage to the steering linkage.

System Checks
Electric Elevator Trim Check
The electric elevator trim system check ensures that each split switch functions correctly, the pilot’s elevator
trim switch overrides the co-pilot’s switch and that both red disconnect switch disables the electric elevator trim
system.

SOP - Electric Elevator Trim

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 Flap System Check
The flap system check verifies the correct function of the flap control and indicating systems in both extension
and retraction.

SOP – Flap System Check

Flight Controls Check

The flight controls check is conducted prior to every flight to verify free and correct operation. To properly check
the full range flight controls the following steps should be followed.

(FWD + Roll Left, FWD + Roll Right, Roll right + RWD, RWD + roll left, Roll Left + FWD, level +FWD, Level + AFT)

SOP - Flight Controls Check

Stall Warning System Test

The stall warning system (aural warning portion) is tested by holding the combination stall warning/overspeed
warning test switch to STALL WARNING TEST and verify the appropriate stall warning sound is heard.

SOP – Stall Warning System Test

Yaw Damp Check (includes Rudder Boost warning system)

The yaw damp system and rudder boost warning system are checked together as they both use the same switch,
located on the left-rear portion of the lower pedestal, to test the system. The yaw damp system is checked the
same way on both rudder boost computers and ensure that the system operates correctly and can be disabled
by pressing the 1st level of either red disconnect button and that the YAW CONTROL TEST switch disables the
yaw damp system when selected.

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 SOP - Yaw Damp Check

Rudder Boost Check

The rudder boost system check ensures that the system correctly identifies and corrects for a loss of torque
(simulates an engine failure) of either engine. The rudder boost system check also incorporates the overspeed
propeller test to simplify the engine run-up procedures. When a single engine torque is increased sufficiently
the rudder boost system applies rudder pressure towards the engine with the higher torque (simulated good
engine) and the red disconnect switch is held to the 1st level while the rudder boost is active to verify that the
system is rudder boost is interrupted.

SOP – Rudder Boost Check

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ICE AND RAIN PROTECTION

General Description
The Beechcraft 1900D is equipped for flying in a variety of inclement conditions and certified for flight into known ice
conditions. Rain protection is provided by the windshield wipers and weather radar (see Avionics Chapter for details).
The ice protection system can be divided into 2 categories; anti-ice system (used prior to entering and during icing
conditions) and de-icing system (used after the accumulation of ice).

Rain Protection
Each main windshield is equipped with a separate windshield wiper and is controlled by a single switch, located on the
upper-right portion of the overhead panel. The use of the windshield wipers should not be used on dry glass to prevent
damage to the thin anti-static film. The switch that controls the windshield wipers has the following positions/modes:

PARK – Must be held in this position. The wiper blades will return to the parked position on the inboard side of
each windshield. However, with the windshield wipers in the parked position, if held to this position for several
seconds, the windshield wipers will cycle once and return the park position. (Momentary Switch position)

OFF – Windshield wipers are off

SLOW – The windshield wipers will operate continuously at slow speed

FAST – The windshield wipers will operate continuously at fast speed

Ice Protection
Anti-Ice Systems

Engine Intake Heat

To prevent ice accumulation on the engine air intake, a scoop in the left exhaust stack deflects some of
the hot exhaust gases through the hollow engine air inlet lip and is exhausted out of the right exhaust
stack. The engine intake lip heat operates whenever the respective engine is operating. No control or
warning systems are required for this system.

Engine Anti-ice
Each engine nacelle is equipped with an inertial bypass separator that prevents moisture particles from
entering the engine plenum during freezing conditions. The inertial separators are used during ground
operations to prevent FOD (foreign objects and debris) from entering the engine and causing damage.

The engine anti-ice inertial separators are controlled by 4 switches located on the right side of the pilot’s
left subpanel. The top 2 switches select the engine anti-ice ON/OFF for each engine and the bottom 2

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 switches select which actuator are used to move the ice vanes and bypass doors on its respective
engine.

Engine Anti-Ice

Note: There is only one actuator within each engine anti-ice system; however, each actuator contains a
main winding and standby (backup) winding that can be used to move the actuator.

When the ENGINE ANTI-ICE switch is turned ON, the respective ice vane extends downward into the
engine air intake airflow creating a venturi effect, while at the same time; the bypass door retracts
upwards and opens the bypass exit duct. As the mixture of air and contaminants (ice, water droplets
and/or FOD) enter the engine intake, it is accelerated by the venturi effect and then the air turns and
enters the engine plenum. Due to the greater inertia of the heavier contaminants, they cannot make the
turn and continue out the bypass exit duct. Once the engine anti-ice vane and bypass door have fully
opened, its respective green “L ENG ANTI-ICE” or “R ENG ANTI-ICE” annunciator will illuminate.

Note: When the engine anti-ice is turned ON, the respective engine torque and fuel flow will
decrease and ITT will increase due to the reduction of ram air and the loss of air through the
bypass exit duct.

When the ENGINE ANTI-ICE switch is turned OFF, the respective ice vane retracts upwards out of the
engine air intake airflow and the bypass door extends downwards and closes the bypass exit duct.
Anything that enters the engine intake is ducted directly into the engine plenum. Once the engine anti-
ice vane and bypass door have moved out of the fully open position, its respective green “L ENG ANTI-
ICE” or “R ENG ANTI-ICE” annunciator will extinguish.

Note: When the engine anti-ice is turned OFF, the respective engine torque and fuel flow will
increase and ITT will decrease due to the increase in ram air.

If either engine anti-ice vane and bypass door does not reach the selected position within 30 seconds its
respective yellow “L ENG ICE FAIL” or “R ENG ICE FAIL” annunciator will illuminate.

Note: If the engine anti-ice vane and bypass door reach their selected position after the
illumination of the “L ENG ICE FAIL” or “R ENG ICE FAIL” have illuminated, the respective
annunciator will extinguish.

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 Windshield Heat
The pilot and co-pilot main windshields are heated by internal heating elements to prevent ice
accumulation. Each window is controlled by separate, (3 position) switches located on the pilot’s right
subpanel.

NORMAL – The windshield heat is ON and the entire windshield is being heated. (Gated
Position)

OFF – Windshield heat is OFF. (Gated Position)

HI – The windshield heat is ON and only the inboard 2/3 of the windshield is being heated.
(Gated Position)

When the windshield heat is in the NORMAL or HI position, the temperature controller regulates the
windshield temperature to 90-100°F. When operating in the NORMAL mode, the amount of current
supplied to heat the windshield may be insufficient at higher airspeeds to prevent ice accumulation. The
HI mode is more effective at preventing ice accumulation as the same current is now concentrated over
a smaller area and will allow the windshield heat to reach and maintain the desired temperature faster.

The maximum airspeed for effective windshield anti-icing is 223 KTS. For prolonged flight in sustained
icing conditions, the airspeed may need to be reduced in order to prevent ice accumulations.

Note: Use of either windshield heat will cause erratic compass indications.

Windshield Anti-Ice

Pitot-Static Heat
The left and right pitot-static masts, located on the upper sides of the fuselage below and in front of the
cockpit side windows, are heated to prevent ice accumulation. Each pitot-static mast heat is controlled
by a circuit breaker type switch located on the pilot’s right subpanel.

Whenever there is insufficient current being supplied to either pitot-static mast, or the switch is in the
OFF position, the respective yellow “L PITOT HEAT” and/or “R PITOT HEAT” annunciators will illuminate.

Pitot Heat

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 Stall Warning Heat
The stall warning transducer, located on the left wing between the engine nacelle and wingtip, is heated
to prevent ice accumulation. The metal plate and tab are heated by separate elements within the stall
warning transducer and are both controlled by the STALL WARN switch located on the pilot’s right
subpanel.

When the STALL WARN switch is ON, the stall warning heat output changes, depending if the aircraft is
on the ground, or airborne as determined by the left main landing gear squat switch. While operating on
the ground the voltage being supplied to the stall warning heat is reduced to prevent overheating of the
stall warning transducer and surface deicing boot. Once in flight, full voltage is supplied to the stall
warning heat to provide sufficient anti-icing at low temperature and high airspeeds.

Whenever there is insufficient current being supplied to the stall warning heat for operating in ground
mode, or the switch is in the OFF position, the respective yellow “STALL HEAT” annunciator will
illuminate.

Note: There is no warning system for the flight mode and the stall warning plate and tab must
be visually monitored by the flight crew to ensure proper operation and no ice is accumulating.

Stall warning heat

Alternate Static Port Heat

The left and right alternate static ports located on the lower portion of the fuselage, below the cockpit
windows, are heated to prevent ice accumulation. Both alternate static heats are controlled by a 5 AMP
circuit breaker type switch located on the pilot’s right subpanel.

Note: There is no warning system to indicate a failure of the alternate static port heat and its operation
must be verified during the first flight system checks.

Alternate Static Port Heat

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 Fuel Vent Heat
The left and right ram air fuel tank vents, located on the bottom of the wing and inboard of the wing tip,
are heated to prevent ice accumulation. Each ram air fuel tank vent heat is controlled by a toggle switch
located on the pilot’s right subpanel

Fuel Vent Heat

Fuel Heater
The oil-to-fuel heat exchanger uses the heat from the engine’s oil system to pre-heat the fuel and
prevent icing in the other fuel system components. A temperature sensing bypass valve constantly
adjusts the flow of oil through the heater core to maintain a fuel temperature of 21-32°C (70-90°F).

Deice Systems

Surface Deice
The airframe is equipped with a surface deice system to remove ice accumulation from the leading edge
of critical flight surfaces. Surface deice boots are installed on the following: outboard wing, inboard
wing, horizontal stabilizer, stabilons and tailets.

The surface deice system is controlled by an electronic distributor valve and powered by the
pneumatic/vacuum system. The distributor valve is controlled by the SURFACE DEICE switch, located on
the right pilot’s subpanel. The switch has 3 positions:

SINGLE –Activates the single automatic surface deice cycle (Momentary switch position)

OFF – Surface deice system is not active, boots are held against wing. (Spring loaded position)

MANUAL – Must be held in this position to activate all the deice boots simultaneously
(Momentary switch position)

While one/both engines are operating, and its respective bleed air switch is in either the ENVIRO OFF or
OPEN positions, 18 PSI pneumatic air and vacuum are supplied to the surface deice distributor valve.
When the surface deice system has not been activated, the distributor valve routes the vacuum supply
to all the deice boots to hold them tightly against the leading edge of the surface.

When the SINGLE mode of the surface deice system is selected, 18 PSI pneumatic air is routed into the
outboard wing surface deice boots for 6 seconds, while vacuum is still applied rest of the deice boots.
Once the pressure within the outboard wing deice boots reaches 12 PSI the green “OUTBD WG DEICE”
annunciator will illuminate. After 6 seconds has elapsed, 18 PSI pneumatic pressure now is routed into
the inboard wing deice boots, horizontal stabilizer, stabilons and tailet for a further 6 seconds, while
vacuum is now applied to the outboard wing deice boots. Once the pressure within the inboard wing
deice boots reaches 12 PSI the green “INBD WG DEICE” annunciator will illuminate and once the
pressure within the horizontal stabilizer, stabilons and tailet boots reaches 12 PSI the green “TAIL DEICE”

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 annunciator will illuminate. After 6 more seconds has elapsed, vacuum is once again applied to all the
surface deice boots.

When the MANUAL mode of the surface deice system is selected, 18 PSI pneumatic air is routed into all
the de-icing boots for as long as the switch is held in the MANUAL position. Once the pressure within
the respective de-icing boots reaches 12 PSI, the green “OUTBD WG DEICE”, “INBD WG DEICE” and “TAIL
DEICE” annunciators will illuminate.

Note: The illumination of the green OUTBD WG DEICE”, “INBD WG DEICE” and “TAIL DEICE”
annunciators indicate that pressure within the respective de-icing boots is sufficient for de-icing.
However, this does not mean that the de-icing boots are inflating and a must be verified visually.

Note: The surface deice system must not be activated below 40°F. Exceeding this limit can result in
permanent damage to the de-icing boots.

Propeller Deice
The propeller deice system consists of an electrically heating deice boot on each propeller blade, a brush
block and slip ring assembly on each propeller, system timer and AUTO and MANUAL switches.

The propeller deice system is controlled by 2 switches, located on the pilot’s right subpanel. The left
switch has 2 positions (AUTO/OFF) and turns the automatic propeller deice system ON heats all the
blade elements on one propeller for 90 seconds, then heats all the blade elements on the other
propeller for 90 seconds. This complete cycle repeats every 3 minutes. The right switch has 2 positions
(MANUAL/OFF) and when held in the MANUAL position, all the propeller blade elements on both
propellers are heated for as long as the switch is held in the position (Momentary switch position).

Prop De-Ice

When the propeller deice system is active, the system timer controls the heat cycles and applies power
to the appropriate propeller blade elements through a rotating slip ring and brush block assembly.

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 The propeller deice system is monitored by the PROP AMP gauge, located on the lower right side of the
overhead panel. When one or both propeller deice systems are active, the gauge displays the amperage
(AMP) supplied to each propeller. When active the respective prop amp gauge should indicate 26-32
amps when all the elements are operating correctly. As the amperage is distributed between 4 blade
elements on each propeller, should an element fail, the resulting amperage be lower than required.

To protect the propeller deice system from damage, when an engine’s “LOW OIL PRESS” annunciator is
triggered by the oil pressure sensor, the signal also disables the propeller deice system on the respective
propeller when the AUTO mode is selected. The manual mode will override this restriction and allow all
the propeller blade elements to be heated, even if the engine is not operating. The propeller system
operation is limited to one cycle per propeller at ambient temperatures above 10°C.

Brake Deice
The brake deice system consists of bleed air lines, solenoid operated shutoff valve, brake deice
manifold, control switch and a fire protection system.

Brake De-Ice

The brake deice system is controlled by the ON/OFF BRAKE DEICE switch, located on the pilot’s right
subpanel. When the switch is in the ON position, the solenoid operated shutoff valve opens and allows
raw P3 compressor bleed air to flow through a line on the left side of the wheel well and down the back
of the gear leg and into each of the brake deice manifolds. The manifold distributes the hot air onto the
brakes of each brake assembly. When the switch is placed in the OFF position, the solenoid operated
shutoff valve closes and stops the flow of raw P3 bleed air.

When the brake deice system is turned ON and the engines are operating the respective green “L BK
DEICE ON” and “R BK DEICE ON” annunciators will illuminate to indicate that the shutoff valve is open.

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 As the brake deice system utilizes P3 bleed air, when the system is turned ON, the engine ITTs will
increase by approximately 20C°. Once the brake deice system is turned OFF, the respective green “L BK
DEICE ON” and “R BK DEICE ON” annunciators will extinguish to indicate the shutoff valve is fully closed.

Note: The brake deice system should be used only as long as needed as prolonged use will increase the
brake temperature sufficiently to degrade the aircrafts braking ability.

(See Fire Protection for brake deice fire protection system)

System Checks
Windshield Anti-ice
The windshield anti-ice is checked to ensure power is being supplied to each windshield in the different modes.

Select the left windshield anti-ice switch to HI and check for an increase on the loadmeters, then select the right
windshield anti-ice switch to HI and check for a further increase on the loadmeters. Select both windshield anti-
ice switches to OFF, then Select the left windshield anti-ice switch to NORMAL and check for an increase on the
loadmeters, then select the right windshield anti-ice switch to NORMAL and check for a further increase on the
loadmeters. Select both windshield anti-ice switches to OFF.

SOP – Windshield Anti-ice Check

Prop Deice
The propeller deice is checked to ensure that all the heating elements on both propellers are functioning in the
different modes.

Select the auto prop deice switch to AUTO and check that either the left or right prop amp gauge indicates 26-32
Amps. Cycle the prop deice switch OFF then back to AUTO and check that the opposite prop amp gauge
indicates 26-32 Amps. Select the auto prop deice switch to OFF.

Hold the manual prop deice switch to MANUAL and check that both the left and right prop amp gauges indicate
26-32 Amps each. Release the manual prop deice switch.

SOP – Prop Deice Check

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 Surface Deice
The surface deice system is checked to ensure proper operation of the deicing boots and indication system.

Hold the surface deice switch to MANUAL and observe an initial decrease then increase back to normal on the
pneumatic pressure gauge, then verify that the green “OUTBD WG DEICE”, “INBD WG DEICE” and “TAIL DEICE”
annunciators illuminate. Release the surface deice switch and verify that the green “OUTBD WG DEICE”, “INBD
WG DEICE” and “TAIL DEICE” annunciators extinguish.

Next momentarily select the surface deice switch to SINGLE and observe an initial decrease then increase back
to normal on the pneumatic pressure gauge and that the green “OUTBD WG DEICE” annunciator illuminates
within 6 seconds (from selecting SINGLE) and then extinguishes followed by the green “INBD WG DEICE” and
“TAIL DEICE” annunciators illuminating within 6 seconds (from when the “OUTBD WG DEICE” annunciator
extinguishes).

Note: Whenever the “OUTBD WG DEICE” or “INBD WG DEICE” annunciators illuminate, visually verify
that the surface deice boots have inflated.

SOP – Surface Deice Check

Misc Deice
The misc deice systems are checked to ensure proper operation and where applicable, operation of the warning
systems.

Turn off the right generator in order to more accurately detect changes to the aircraft’s electrical load and verify
that the yellow “R DC GEN” annunciator illuminates.

Select the left and right fuel vent switches individually ON then OFF and check for a slight increase then decrease
on the left loadmeter.

Turn the stall warn heat switch to ON, check that the yellow “STALL HEAT” annunciator illuminates and that the
left loadmeter increases slightly. Turn OFF the stall warn heat switch and check that the respective annunciator
extinguishes and that the left loadmeter decreases slightly.

Turn the left and then right pitot head switches to ON, check that the yellow “L PITOT HEAT” then “R PITOT
HEAT” annunciators illuminate and that the left loadmeter increases slightly. Turn OFF the left then right pitot
heat switches, and check that the respective annunciators extinguish and that the left loadmeter decreases
slightly.

Turn the alternate static heat switch ON then OFF and check for increase then decrease on the left loadmeter.

Turn the right generator ON and verify that the yellow “R DC GEN” annunciator extinguishes.

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 SOP – Misc Deice Check

Brake Deice
The brake deice system is checked to ensure proper operation of control and indication components of the
system.

Turn the brake deice switch to ON, check the left and right engine ITT gauges for a slight increase and verify that
the green “L BK DEICE” and “R BK DEICE” annunciators illuminate. Turn the brake deice switch to OFF, check the
left and right ITT gauges for a slight decrease and verify that the respective annunciators have extinguished.

SOP – Brake Deice Check

Limitations
Icing Limitations

1) Maximum Airspeed for Effective Windshield Anti-icing 223 knots

2) Minimum Temperature for Operation of Deicing Boots -40°C
3) Minimum Airspeed for Sustained Level Flight in Icing Conditions 160 Knots
4) Sustained flight in icing conditions with flaps extended is prohibited except for approach and landing.
5) Engine Anti-ice shall be on for operations in ambient temperatures of +5C or below or when flight free of
visible moisture cannot be assured.
6) Operation of propeller deice is limited to one cycle per propeller at ambient temperatures above 10°C.

Propeller Deice System

To avoid exceeding the temperature limit of the adhesives which attach the deice boots and the erosion
shield to the propeller, observe the following:

1) Ensure system is OFF when propellers are static.

2) System operation is limited to one cycle per propeller at ambient temperatures above 10°C.

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AVIONICS

General Description
The Beechcraft 1900D is equipped with complex flight instruments, Collins Pro Line II avionics and navigation package,
GPS and integrated weather radar system to enable the aircraft to be operated in most flight conditions while reducing
the workload on the crew. Most of the avionics and navigation system have control units mounted on the instrument
panel with the respective processing unit/computer mounted within line replaceable units (LRUs) within the nose of the
aircraft in the avionics bay.

Avionics Master Switch

The design of the avionics master switch is such that if the switch should fail, the avionics bus circuit breaker can be
manually pulled out and power will be restored to the avionics.

Avionics Master Switch and C/B

Avionics Master Switch Operation

When the avionics master switch is turned ON, power is removed from each of the 3 avionics bus relays and
allows power to flow into the respective avionic buses. Regardless of the position of the avionics master switch,
if the avionics circuit breaker is pulled out, power will be removed from the 3 avionics bus and power all the
avionics.

Note: It is important to always turn ON the battery master switch prior to turning ON the external power
switch as power is used to remove power from the avionics buses. The avionics buses must be
unpowered prior to introducing an external power source to prevent damage to the individual
components.

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Primary Flight Instruments
Airspeed Indicator
Each of the two airspeed indicator has an indication range from 40 to 300 knots and contains two pointers. The
white pointer displays the aircrafts indicated airspeed as determined by the pitot-static system and the stripped
red/white pointer (barber pole) displays the current maximum allowable airspeed (MAAS). The MAAS, on a
standard day, is 248 knots from seal level to 13,200’and is a structural limitation (VNE), the MAAS then decreases
from 248 knots to 195 knots at 25,000’ and is a limitation for airflow over the wing going past the speed of
sound (VMO).

Should the aircrafts indicated airspeed (white pointer) increase and equal to or greater than the MAAS (stripped
red/white pointer), an overspeed aural warning sounds through the audio system. The airspeed indicator
operates on pitot-static air pressure; however, the overspeed warning system requires 28 VDC to function.

Altimeter
Each of the two encoding altimeters provides the flight crew with a bar-corrected altitude and an altitude signal
to the transponder. The altimeter displays the aircraft altitude as determined by the static portion of the pitot-
static system and has a correction knob to adjust for non-standard pressure. The pilot’s altimeter is connected
to a KEA-346 encoding altimeter and connected to the KAS-297A altitude alerter which produces a chime
through the audio system when the aircraft passes 1000’ from selected altitude and when the aircraft deviates
400’ from the selected altitude. If the aircraft is equipped with an autopilot, an ALI-80A encoding altimeter is
installed and connected directly to the EFIS-84 system. The EFIS-84 system produces a chime through the audio
system when the aircraft passes 1000’ from the selected altitude and when the aircraft deviates 200’ from the
selected altitude. The ALI80A contains a test button that is used when testing the autopilot system. Regardless
of which encoding altimeter is installed, static air pressure and 28 VDC are required for its operation.

Instantaneous Vertical Speed Indicator

Each of the two instantaneous vertical speed indicators indicates the vertical velocity of the aircraft in feet per
minute (FPM). The instantaneous vertical speed indicators operate on static air pressure.

EFIS-84 System
The Collins 4-tube version of the Electronics Flight Instrument System (EFIS) consists of four, panel mounted
Electronic Flight Displays (EFD-84), two center console mounted Display Select Panels (DSP), two remote Display
Processor Units (DPU), a reversionary control panel, power control panel and integrated with a weather radar
system (WXR-350). The connections between these various components provide fail-safe operation as shown
below.

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 The power supply for the EFIS system is controlled by 4 switches that are located on the top-left side of the
lower pedestal. The pilot and co-pilot EFIS screens are controlled by the respective PILOT EADI EHSI and COPILOT
EADI EHSI switches. The pilot and co-pilot DPU systems are controlled by the respective PILOT DSPL PRCSR and
COPILOT DSPL PRCSR.

Note: To protect against switch failure, the pilot’s EFIS screen and DPU power switches are overridden
when airborne and turning on these switches will have no effect on the respective systems until the
aircraft lands.

EFIS-84 Panel

Electronic Attitude Direction Indicator (EADI-84)

The EFIS-84 EADI is a multi-colour Cathode Ray Tube (CRT) display that provides aircraft attitude, flight
control system commands and track/deviation information from the various navaids. The top of the
display is used to annunciate the flight control system modes in operation, and include: autopilot
engagement annunciation, altitude sensor annunciation, marker beacon annunciation, radio altitude,
decision set height and annunciation and comparator warnings. An inclinometer is also mounted on the
faceplate at the bottom of the EADI. The EADI screen is powered by 28 VDC, equipped with an EFIS Aux
Power backup system, and the DPU (see below) is powered by 115 VAC.

Note: All of the EFIS screens (EADI and EHSI) are identical and can be swapped after changing
the inclinometer is required.

The brightness of the pair of EADI and EHSI screens is controlled by a dual rotary knob that is located to
the left of the EHSI. The outer knob controls the EADI brightness level and the inner knob controls the
EHSI brightness level.

Note: Operating the EFIS screens at maximum brightness for extended periods of time may
result in a condition known as imprinting. Imprinting causes the images to remain on the CRT
display even when the unit is OFF, or seeing and image other than desired. This is most
commonly visible when an EHSI and EADI screen are swapped.

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 The EADI detects excessive aircraft attitude and removes all data displayed except for the aircraft
symbol and attitude. An excessive attitude is defined as: >30° nose up, >20° nose down or 65° of bank. If
the excessive aircraft attitude is due to excessive pitch, a series of red chevrons appear on the display
and point towards zero pitch.

Electronic Horizontal Situation Indicator (EHSI-84)

The EFIS-84 EHSI is a multi-colour Cathode Ray Tube (CRT) display that provides true or magnetic
heading, DG mode annunciation, selected heading, selected navaid course and deviation, navigation
sensor annunciation, digital selected course/desired track readout, heading sensor annunciation, ILS and
heading comparator warnings, to/from information, back course localizer annunciation, distance to
station/waypoint, glideslope, groundspeed, time-to-go, elapsed time, course information and sensor
annunciation from a seconds navigation sensor, weather radar target alert, waypoint alert when LNAV is
the navigation sensor, and two bearing pointers that can be driven by the various navigation sensors.
The EHSI screen is powered by 28 VDC, equipped with an EFIS Aux Power backup system, and the DPU
(see below) is powered by 115 VAC.

Note: All of the EFIS screens (EADI and EHSI) are identical and can be swapped after changing
the inclinometer is required.

The brightness of the pair of EADI and EHSI screens is controlled by a dual rotary knob that is located to
the left of the EHSI. The outer knob controls the EADI brightness level and the inner knob controls the
EHSI brightness level.

Note: Operating the EFIS screens at maximum brightness for extended periods of time may
result in a condition known as imprinting. Imprinting causes the images to remain on the CRT
display even when the unit is OFF, or seeing and image other than desired. This is most
commonly visible when an EHSI and EADI screen are swapped.

EFIS Aux Power

The aircraft is equipped with an electrical backup system for the pilot’s EFIS screens (EADI and EHSI
only), located in the lower-left side of the instrument panel, and is capable of providing power to the
pilot’s EFIS screens for approximately 3-5 minutes.

Note: There is not EFIS Aux Power system for the co-pilot’s EFIS screens.

The purpose of this system is to prevent the pilot’s EFIS screens from failing during engine airstarts as
part of the FAA single pilot certification requirements. However, the system will as provide a time
limited backup during certain system failures.

Note: The crew must be aware that if power is not restored within approximately 3-5 minutes
the EFIS Aux Power battery will become depleted and the unit will fail.

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 With the aircraft’s electrical systems ON and the EFIS Aux Power control switch in the ARM position, the
green “AUX ARM” annunciator illuminates. Should an electrical failure occur or the system voltage drops
below 18 VDC, the EFIS Aux Power battery will power the pilot’s EFIS screens and the yellow “AUX ON”
annunciator illuminates and the aural warning is heard (beeping tone). If desired, the crew can
acknowledge and silence the aural warning by momentarily pressing the HORN SILENCE button.

Note: The crew must be aware that if power is not restored within approximately 30 minutes
the EFIS Aux Power battery will become depleted and the EFIS screens will fail.

Display Processing Unit (DPU-84)

The Display Processing Unit (DPU) provides the video display for the EFIS screens. The DPU also provides
sensor input processing, switching, the necessary deflection and power for the information presented
on the EFIS screens. The DPU units are mounted in the avionics bay within the nose of the aircraft.

Display Select Panel (DSP-84)

The Display Select Panel provides the respective EHSI with format and input selection. The DSP controls
navigation sensor selection, bearing pointer selection, weather radar selection, navigation data selection
(groundspeed and time-to-go), and the selection of course information from a section navigation sensor.

DSP-84

The functions of the DSP buttons and knobs are described below.

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 EHSI Format Switch (Center Rotary knob)

MAP – The MAP mode uses the expanded compass format but places VOR and/or
waypoint symbols for both the active and preselected course in proper position in
respect to the airplane symbol and selected range.

ARC – The ARC mode selects an expanded compass segment at the top of the display
with the airplane symbol centered at the bottom. Information similar to that presented
the full compass rose is displayed in an enlarged, easily readable format.

HSI – The HSI mode displays the full compass rose (similar to a conventional HSI).

ARC/WX – The ARC/WX mode is the same at the regular ARC mode with the addition of
weather radar information being added to the display.

MAP/WX – The MAP/WX mode is the same at the regular ARC mode with the addition
of weather radar information being added to the display.

In the center of the EHSI Format Switch is the RNG knob and is used to select the range on the
EHSI display when in MAP or ARC modes (does not include MAP/WX and ARC/WX modes)

NAV DATA Switch (Left Rotary Knob)

Rotating the NAV DATA switch selects the information displayed on the top right of the
EHSI. The selections are:

GSP – Groundspeed from the active navigation sensor.

TTG – Time-To-Go to the active navigation sensor VOR/Waypoint.

ET – Elapsed Time countdown/up as set by the Timer Set knob (see below)

In the center of the NAV DATA switch is the HDG knob that is used to position the
heading cursor (bug) as desired on the EHSI. Within the HDG knob is the PUSH HDG
SYNC button that when pressed, causes the heading cursor (bug) to rotate and match
the aircrafts current heading.

COURSE Switch (Right Rotary Knob)

Rotating the CRS SEL switch selects the active and preset navigation sensors for course
control and selection, as well as replacing the active navigation sensor with the preset
navigation sensor. The positions are:

PRE – In this position the selection of navigation sensors and course control is
entered into the preselect position. The preselected course is shown in cyan as a
double dashed line.

ACT – In this position the selection of navigation sensors and course control is
entered into the active position. The active course is shown in white as a solid
single line.

XFR – Holding the COURSE switch in the XFR position replaces the ACT
navigational sensor with the PRE navigational sensor and return to the ACT
position when release. When transferred, the preselected navigation sensor
becomes the active navigation sensor and changes to a white solid line. The
preselect position will be blank as no navigation sensor is now preselected.

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 In the center of the COURSE switch in the CRS knob that is used to control the selected
course of the selected navigation sensor. Within the CRS knob is the PUSH CRS DIRECT
button that when pressed, causes the selected course (as determined by the COURSE
switch position) to rotate directly towards the station until the lateral deviation is zero
(only for VOR navigation sensors).

TIMER SET knob (Smaller Left Rotary Knob)

Rotating the TIMER SET knob is used to preset the countdown timer that is displayed on
the EHSI. The TIMER SET knob is also the Start/Stop (S/S) button that is used to control
the timer. After presetting a time, pressing the S/S button will start the countdown.
Once the timer has reached 00:00, the digits flash for 5 seconds then change to a line of
dashes. With the timer at zero, press the S/S button to start the count-up timer. Press
the S/S button again to stop and hold the timer at the current value. Press the S/S
button a third time to reset the timer to zero.

DH SET knob (Smaller Right Rotary Knob)

Rotating the DH SET knob is used to set the decision height (DH) that is displayed on the
EADI. The DH can be set between 0 and 999’. The DH SET knob is also the test button
(TST) that is used to start the radio altimeter test.

Bearing Pointer Buttons (Single and Double Bar Buttons)

The single line and double line bearing pointer select buttons are used to select the
navigation sensor and respective cyan single line or magenta double line bearing
pointers on the EHSI. Repeatedly pressing either button will cycle through the available
navigation sensors and once released on a selected, will display the respective bearing
point on the EHSI. Pressing and holding a Bearing Pointer Button will remove the
respective bearing pointer from the EHSI.

The cyan single line bearing pointer can select the NAV1, GPS and ADF1 navigational
sensors, and the magenta double line bearing pointer can select the NAV2, GPS and
ADF2 navigational sensors.

CRS SEL Button

The CRS SEL button is used to connect the desired navigation sensor to the EHSI ACT and
PRE positions. When the COURSE knob is in the ACT position, pressing the CRS SEL
button will cycle through the available navigations sensors and when the desired
navigation sensor is displayed, holding the CRS SEL button will select that navigation
sensor. When the COURSE knob is in the PRE position, pressing the CRS SEL button will
cycle through the available navigation sensors, holding the CRS SEL button will remove
the selected PRE navigation sensor.

Reversionary Panel
The EFIS-84 reversionary panel contains multiple switches to control the various reversionary modes for
the pilot and co-pilot EFIS systems, as well as the Yaw Damper computer switch or Autopilot computer
switch and Air Data Computer test switch (ADC) (autopilot equipped aircraft).

Reversionary Switches
PLT CMPST – Selects the pilot composite reversionary mode ON/OFF. When turned ON,
both of the pilot’s EFIS screen display a composite version of the EADI and EHSI. The

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 resulting display incorporates mostly the normal EADI with heading, courses and bearing
displayed in a compass arc at the bottom of the display.

DSP – Selects which Display Select Panel controls the pilot’s and co-pilot’s EFIS systems.
In the NORM position, the pilot’s DSP controls the pilot’s EFIS system and the co-pilot’s
DSP controls the co-pilot’s EFIS system. When selected to PLT, the pilot’s DSP controls
both EFIS systems and when selected to COPLT, the co-pilot’s DSP controls both EFIS
systems. To remind the flight crew that this EFIS reversionary mode is active, when PLT
is selected, an amber “DSP 1” annunciation is displayed in the lower-right corner of the
co-pilot’s EFIS screens and the data is coloured yellow. When COPLT is selected, “DSP 2”
is displayed in the lower-right corner of the pilot’s EFIS screens and the data is coloured
yellow.

DR XFR – Selects which Display Processing Unit controls the pilot’s and co-pilots EFIS
displays. In the NORM position, the pilot’s DPU produces the video display for the pilot’s
EFIS system and the co-pilot’s DPU produces the video display for the co-pilot’s EFIS
system. When selected to PLT, the pilot’s DPU produces the video display for both EFIS
system and when selected to COPLT, the co-pilot’s DPU produces the video display for
both EFIS system. As the DPU produces the video display that is requested by the DSP,
when the switch is in the PLT position, the pilot’s DSP controls all the EFIS systems and
when in the COPLT position, the co-pilot’s DSP controls all the EFIS systems. To remind
the flight crew that this EFIS revisionary mode is active, when PLT is selected; an amber
“ATT 1” annunciation is displayed in the lower right corner of both EADI screens and an
amber “MAG 1” annunciation is displayed at the top of both EHSI screens. W hen COPLT
is selected; amber “ATT 1” and “ATT 2” are displayed. Also, the external amber “DRIVE
XFR” annunciator illuminates on the EFIS system that is receiving cross-sided data and
the respective Flight Control Panel go blank (Flight Director is removed from respective
EADI).

COPLT CMPST – Selects the co-pilot composite reversionary mode ON/OFF. When
turned ON, both of the co-pilot’s EFIS screen display a composite version of the EADI
and EHSI. The resulting display incorporates mostly the normal EADI with heading,
courses and bearing displayed in a compass arc at the bottom of the display.

ATT – Selects which Attitude and Heading Reference System (AHRS) supplies attitude
data to which EFIS system. In the NORM position, the pilot’s AHRS supplies attitude data
to the pilot’s EFIS system and the co-pilot’s AHRS supplies attitude data to the co-pilot’s
EFIS system. In the ALL ON NO.1 position, the pilot’s AHRS (AHRS #1) supplies attitude
data to both EFIS systems and in the ALL ON NO.2 position, the co-pilot’ AHRS (AHRS #2)
supplies attitude data to both EFIS systems. To remind the flight crew that this EFIS
revisionary mode is active, when PLT is selected, an amber “ATT 1” annunciation is
displayed in the lower right corner of both EADI screens and when COPLT is selected, an
amber “ATT 2” is displayed.

HDG – Selects which Attitude and Heading Reference System (AHRS) supplies heading
data to which EFIS system. In the NORM position, the pilot’s AHRS supplies heading data
to the pilot’s EFIS system and the co-pilot’s AHRS supplies heading data to the co-pilot’s
EFIS system. In the ALL ON NO.1 position, the pilot’s AHRS (AHRS #1) supplies heading
data to both EFIS systems and in the ALL ON NO.2 position, the co-pilot’ AHRS (AHRS #2)
supplies heading data to both EFIS systems. Also, to remind the flight crew that this EFIS
revisionary mode is active, when PLT is selected, an amber “HDG 1” annunciation is
displayed at the top of both EHSI screens and when COPLT is selected, an amber “HDG
2” is displayed.

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 Other Switches
YD – Selects the left (YD/L) or right (YD/R) Yaw Damp computer for controlling the Yaw
Damp and Rudder Boost system (Non-Autopilot equipped aircraft only)

AP – Selects the Left (AP/L) or right AP/R) autopilot computer for controlling the
autopilot, Yaw Damp and Rudder Boost systems (Autopilot equipped aircraft only)

Other Switches

Magnetic Compass Unit (MCU-65)

The Magnetic Compass Unit is used by the EFIS-84 System to detect the aircraft’s orientation to the
Earth’s magnetic field and control the compass system (displayed heading). The single flux valve is
installed in within the left horizontal stabilizer to minimize electrical interference created by the various
aircraft systems and provides information to both compass systems.

The AHRS Control Panel, located on the outboard side of each EHSI on the instrument panel, is used to
control the compass system. When the green “NORM” annunciator is illuminated, the flux valve controls
the compass system and applies the required corrections. When the annunciator pushbutton is pressed,
the amber “FREE” annunciator is now illuminated and the compass system must be manually corrected
through the use of the non-stabilized magnetic compass and the toggle switch on the AHRS control
panel. When the toggle switch is held to “SLEW CW” the compass system changes the displayed heading
clockwise, and when held to “SLEW CCW” the displayed heading changes counter-clockwise until the
switch is released.

Note: When operating in “FREE” mode, the respective compass system acts like a regular
Directional Gyro (DG) and must be reset regularly to maintain accuracy.

Flight Guidance System (FGS-65) and Autopilot/Flight Director Computer (APC-65H)

Both the Flight Guidance System and Autopilot/Flight Director Computer function similarly. Both units
consist of the same flight director computer (autopilot system has several addition modes) that
processing inputs from many aircraft systems and determines the aircraft’s 3-dimensional orientation
and rate of change. The Flight Director (FD) can then determine the required attitude change to produce
the result selected on the Flight Control Panel.

The main difference between the systems if that the Autopilot/Flight Director Computer outputs
commands to the control surfaces to control the aircraft’s attitude, where the Flight Guidance system
only provides guidance on the pilot’s EADIs and the aircraft’s attitude must be manually changed.

Flight Control Panel (FCP-65)

The Flight Control Panel is used to select the various modes of the Flight Guidance System (Flight
Director) so that the Flight Guidance System can control the V-Bars to achieve the desired result.

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 The aircraft is equipped with 2 FCP, located above the pilots and co-pilot’s EADIs and controls
the respective Flight Directors.

To select a mode on the Flight Control Panel, press the desired mode (must select a lateral
mode prior to selecting a vertical mode) and verify the annunciator on the Flight Control Panel
illuminates and the EADI annunciation is the same. On the EADI, vertical modes are displayed on
the top left and lateral modes on the top right of the screen. When the EADI annunciation is
green, that mode is active, when white, that mode is armed. Once the mode is confirmed, the
pilot now must match the yellow aircraft attitude symbol within the magenta Flight Director V-
Bars to achieve the desired result.

The various modes of the Flight Control Panel are:

HDG – The lateral HDG mode causes the flight director to command turns to match the
aircrafts heading to the heading bug selected by the DSP and a green “HDG”
annunciation is displayed on the Flight Control Panel and EADI.

NAV – The lateral NAV mode cause the flight director to enter the HDG mode and NAV
ARM mode. When selected , a green “HDG” and green “NAV” and white “ARM”
annunciations will be displayed on the FCP and a green “HDG” and white “navigation
sensor” (NAV1, NAV2, LOC1, LOC2, GPS) annunciation on the respective EADI. The
aircraft must be flown in the HDG mode on a reasonable intercept angle towards the
active course until the flight director captures the course. Once captured, the green
“HDG”, green “NAV” and white “ARM” annunciators on the FCP will be replaced with a
single green “NAV" annunciation and the EADI will show a green “navigation sensor”
(NAV1, NAV2, LOC1, LOC2, GPS) annunciation. The flight director will now command
turns to maintain the aircraft on the active course.

APPR – The lateral APPR mode operates similarly to the NAV mode with the addition of
vertical navigation. When selected, the Flight Control Panel and EADI display a green
HDG and white APPR ARM annunciations. Once the localizer is captured, the green
“HDG”, green “APPR” and white “ARM” annunciations replaced by a green “APPR” and
white “GS” on the FCP and a green “navigation sensor” (NAV1, NAV2, GPS, LOC1, LOC2)
and white “GS” on the respective EADI. The Flight Director will command turns to
maintain the aircraft on the active course. Once the glideslope is captured, the white
“GS” annunciation turns green (on the FCP and EADI) and the Flight Director will
command pitch changes to maintain the aircraft on the slope.

B/C – The lateral B/C mode operates similarly to the APPR mode with the ability to track
a localizer course in the opposite direction (back course approaches or tracking away
from a localiser) and disregard all glideslope indications. When selected, a green “HDG”,
green “APPR”, green “B/C and white “ARM” annunciations are displayed on the FCP, and
a green “HDG”, green “B/C” and white “navigation sensor” (NAV1, NAV2, GPS, LOC1,
LOC2) annunciations on the respective EADI. Once the back course is captured, the
green HDG annunciation is replaced by a green “navigational sensor’ (Nav1, NAV2, GPS,
LOC1, LOC2) annunciation, and the green “B/C” remains. The flight director will now
command turns to maintain the aircraft on the active course.

Note: The front course must still be selected on the EHSI as the flight simply
inverts the course signal and uses the reciprocal course for tracking.

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 ALT – This vertical mode causes the Flight Director to command pitch changes to
maintain a certain barometric altitude. When selected, the Flight Director will maintain
the current altitude (or preselected altitude when automatically capturing in autopilot
equipped aircraft) a green ALT annunciation is displayed on the Flight Control Panel and
EHSI. Pressing the yoke mounted Control Wheel Steering (CWS) button will set the
current altitude as the new altitude to be maintained.

VS – This vertical mode causes the Flight Director to command pitch changes to
maintain a desired vertical speed. When selected, a green VS annunciation is displayed
on the Flight Control Panel and EADI. Pressing the yoke mounted CWS button will set
the current vertical speed as the new vertical speed to be maintained.

IAS – This vertical mode causes the Flight Director to command pitch changes to
maintain a desired Indicated Airspeed (IAS). When selected, a green IAS annunciation is
displayed on the Flight Control Panel and EADI. Pressing the yoke mounted CWS button
will set the current IAS as the new IAS to maintain.

The additional modes of the Flight Control Panel for autopilot equipped aircraft are:

CLIMB – This vertical mode causes the Flight Director to command a climb at a
predetermined climb rate and a green CLIMB annunciation is displayed on the Flight
Control Panel and EADI.

DSC – This vertical mode causes the Flight Director to command a descent at a
predetermined descent rate and a green DSC annunciation is displayed on the Flight
Control Panel and EADI.

ALT SEL – This vertical mode causes the Flight Director to automatically maintain the
selected altitude once it is reached. To select this mode, a lateral mode, plus another
vertical mode (IAS, VS) must be selected. When selected, a white ALT SEL annunciation
is displayed on the Flight Control Panel and EADI. The Flight Director calculates when to
capture the selected altitude based on the aircrafts vertical speed. At the calculated
altitude, a green ALT annunciation replaces the white ALT SEL annunciation and the
Flight Director will maintain the selected barometric altitude.

Autopilot System Control (APS-65) – If Installed

When installed, the Autopilot System Control and Autopilot/Flight Director Computer are used to
control the Autopilot System.

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 The buttons and modes of the Autopilot System Control panel are described below.

YAW ENG – The Yaw Damp Engage button is used to engage or disengage the Yaw Damp
system. Once engaged, a green YD annunciation is displayed on both EADIs and on the Flight
Control Panel on the same side as the autopilot computer in use.

AP ENG – The Autopilot Engage button is used to engage or disengage the Autopilot system
once engaged, a green AP/L or AP/R annunciation will be displayed on both EADIs depending if
the left or right autopilot computer is in use.

L / R Knob – The Turn Knob is used to manually control the bank angle. The knob in the center
indicates wings level and any movement left or right will change the bank angle in proportion to
the turn knobs displacement from center. The turn knob use is limited to 30° of bank.

SR – The Smooth Ride button is used to engage or disengage the Smooth Ride autopilot mode.
The SR mode is indented for use in turbulence and changes circuitry within the autopilot system
to provide smoother autopilot movements. When engaged, a green SR annunciation is displayed
on the Flight Control Panel and EADI.

½ Rate – The ½ Rate button limits the maximum bank angle that the Flight Director can
command to ½ of standard. When engaged, a green 1/2 annunciation is displayed on the Flight
Control Panel and EADI.

UP / DOWN Switch – The UP/DOWN vertical control switch provides manual control of the
vertical flight modes. When no vertical mode is selected (pitch hold), the UP/DOWN vertical
control switch directly controls the elevator servo and changes the pitch ½° for every
momentary click. When a vertical mode is selected (IAS, VS) the UP/DOWN vertical control
switch can be used.

Radio Altimeter (ALT-55B)

The single radio altimeter system provides radar altitude data to both EFIS-84 system through the use of
a set of transmitting and receiving antennas that are located on the lower aft portion of the fuselage.

Backup Flight Instruments

Standby Attitude indicator (AI-804AF)
The aircraft is equipped with an electrically powered standby attitude indicator, located in the top-center of the
instrument panel, and is capable of providing attitude reference data for a minimum of 30 minutes in the event
of an aircraft electrical failure. The control panel is located on the right side of the Standby Attitude Indicator.

With the aircraft’s electrical systems ON and the Standby Attitude Indicator control switch in the ARM position,
the green “AUX ARM” annunciator illuminates. The Standby Attitude Indicator also contains a manual gyro
caging system that is manually controlled. Should an electrical failure occur, the Standby Attitude Indicator
battery will power the unit, the yellow “AUX ON” annunciator illuminates and the aural warning is heard
(continuous tone). If desired, the crew can acknowledge and silence the aural warning by momentarily pressing
the HORN SILENCE button.

Note: The crew must be aware that if power is not restored within approximately 30 minutes the
Standby Attitude Indicator will battery will be depleted and the unit will fail.

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 Standby Magnetic Compass
The standby magnetic compass, located in the center and below the overhead panel, is a conventional liquid
filled, non-stabilized compass. The compass is susceptible to errors created by: magnetic fields generated within
the aircraft structure and basic electrical systems (deviation), magnetic field produced by other various electrical
systems (windshield heat, exterior lights), acceleration/deceleration errors and turning errors. A compass
correction card is mounted on the overhead panel and provides corrections to errors caused by deviation only.

Environmental Inputs
Pitot-Static System
The pitot-static air pressure system is used to provide the various flight instruments with a source of ram air and
static air pressure. The 2 pitot-static masts are located on the uppers sides of the fuselage on the nose, below
and ahead of the side cockpit windows. The pitot portion of the left and right pitot-static masts is used to
provide an independent source of ram air pressure for the respective left and right side instruments. Each pitot-
static mast contains 2 independent static ports, one static port for the left side flight instruments and another
for the right side flight instruments to minimize static pressure errors created by changes in the aircraft’s
attitude.

In the event that either primary static ports should become blocked (which would cause erratic indications to
the flight instruments that utilize that static air port) a set of alternate static ports are located on the lower sides
of the fuselage and below the pitot-static mast. The left and right static air source selector valve is located on
the lower outboard sidewalls and has 2 selectable positions:

NORMAL – The respective side flight instruments utilize the static ports within the pitot-static masts. A
red metal guard prevents the lever from moving from this position inadvertently.

ALTERNATE AND DRAIN – The respective side flight instruments utilize the alternate static ports.

Note: When using the alternate static ports, the airspeed and altitude indications change due to
the different location and type of static air being supplied. The flight crew must reference the
airspeed correction and altimeter correction graphs located in the aircrafts Pilot Operating
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 Manual/Aircraft Flight Manual for the correct indications during use of the alternate static air
source.

The following tables list the various flight instruments and the systems there are connected to.

Left Pitot Pressure Left Static Pressure Right Static Pressure Right Pitot Pressure
- Pilot’s Airspeed - Pilot’s Airspeed Indicator - Co-pilot’s Airspeed - Co-pilot’s Airspeed
Indicator - Pilot’s Altimeter Indicator Indicator
- Air Data Sensor #1 - Pilot’s Vertical Speed - Co-pilot’s Altimeter - Flight Data Recorder
- Air Data Computer Indicator - Co-pilot’s Vertical Speed - Air Data Sensor #2
- Air Data Sensor #1 Indicator
- Air Data Computer -Pneumatic Pressure
Indicator
- Flight Data Recorder
- Air Data Sensor #1

Alternate Static Pressure

- Cabin Differential Pressure Switch
- Cabin Differential Pressure Gauge

* All the Left or Right Static Pressure items listed above,

except for the flight data recorder on the right static system,
can be substituted with Alternate Air Pressure on the side
select.

Outside Air Temperature

The single outside air temperature probe is located on the pilot’s upper left sidewall and consists of an indicator
dial and stem that protrudes through the aircraft’s skin and is contained within a sunshield/support to more
accurately display the actual outside air temperature. The indicator displays the outside air temperature
between -60°C to 60°C. While in flight, the indicated outside air temperature may be up to 5 degrees celcius
higher than actual due to the ram rise and friction on the exposed probe.

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Communications
VHF Communication Receiver Control (CTL-22)
The VHF communication receivers receive the communication signals from a two antennas. The antenna for the
#1 communications receiver is mounted on the forward top portion of the fuselage and the antenna for the #2
communications receiver is mounted on the lower aft portion of the fuselage. The frequency range of the VHF-
22A is 118.00 to 135.95 MHZ.

Audio Control System

The aircraft is equipped with one audio control system for each pilot, the two systems are mounted on one
panel and located at the top-center of the instrument panel.

Audio Control System

The switches and associated functions are listed below for the pilot’s system (co-pilot systems works identically):

Audio Input Control

COMM 1 & 2 – Connects the selected communications receiver to the pilot’s headset.

NAV 1 & 2 – Connects the selected navigation receiver to the pilot’s headset for station
identification and/or listening to broadcasts made over the selected VOR frequency.

MKR BCN 1& 2 – Connects the selected marker beacon signal to the pilot’s headset.

DME 1 & 2 – Connects the selected navigation receiver to the pilot’s headset for station
identification.

ADF – Connects the selected ADF receiver to the pilot’s headset for station identification and/or
listening to broadcasts made over the selected NDB frequency.

Voice-Both-Range – When selected to VOICE, the Morse code portion of the signal is removed
so that only the voice broadcast is heard. When selected to RANGE, the voice portion of the
signal is removed so that only the Morse code is heard. When selected to BOTH, both the Morse
code and voice broadcasts are heard.

Microphone Output Control

COMM 1 – Inputs from the pilot’s headset or hand microphone will be broadcast on the
frequency selected on communications receiver #1 whenever the respective Push-To-Talk (PTT)
switch is depressed.

COMM 2 – Inputs from the pilot’s headset or hand microphone will be broadcast on the
frequency selected on communications receiver #2 whenever the respective Push-To-Talk (PTT)
switch is depressed.

PA – Inputs from the pilot’s headset or hand microphone will be heard over the aircraft’s cabin
speakers whenever the respective Push-To-Talk (PTT) switch is depressed.

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 Volume Controls
SPKR – Regulates the speaker volume output to the pilot’s overhead speaker.

PH – Regulates the headset volume output to the pilot’s headset speakers.

AUDIO SPKR – Turns the speaker above the pilot’s seat ON, any incoming/outgoing radio
broadcasts will duplicated over the speaker. Crew interphone communications are cannot be
heard over the speaker.

Other Controls (only one set for both pilot’s)

DME 1 or DME 1 VOL – Regulates the volume of respective DME identification signal.

MKR BCN 1 & 2 VOL – Regulates the volume of both Marker Beacon signals

MKR BCN HI/LOW – Selects the marker beacon receiver sensitivity level. The marker beacon
signal will be heard at a greater distance away from the station when selected to HI and must be
very close to the station when selected to LOW.

HOT INTPH – Actives the hot interphone for communications between the pilot and co-pilot.
When selected ON, the interphone (pilot to pilot) communications are automatically activated
when either pilot’s headset microphone reaches the required noise level. When selected OFF,
the interphone communications must be manually activated by depressing the yoke mounted
interphone switch.

ENCD ALTM 1-ALTM 2 – Selects whether pilot’s (ALTM 1) or co-pilots (ALTM 2) air data encoding
altimeter signal is sent to the transponder for broadcasting.

XPNDR NORMAL-OVERRIDE – (Feature is disabled on all company aircraft) – In the NORMAL

position, the transponder automatically changes to the standby mode when the aircraft is on
the ground (determined by either a landing gear squat switch input or low groundspeed
depending on the transponder system type and installation). When selected to OVERRIDE, the
transponder remains in mode “C” for airports that utilize the transponder signal for its ground
surveillance radar.

GND COMM PWR – This Push ON/Push OFF switch provides power to communications receiver
#1 and the audio system for establishing communications while the aircraft electrical systems
are OFF (Battery Master switch must be off).

ANN PUSH BRT – When pulled out and rotated, the switch regulates the brightness of some
lighting systems and signals the communications receivers, navigation receivers, ADF receiver
and transponder enter dim mode. In dim mode these units use a photocell to control their
individual brightness levels in relation to ambient light conditions. (See Lighting Chapter for
details)

Navigation
Navigation Receiver Control (VIR-32)
The aircraft is equipped with two navigation receivers, located in the center of the instrument panel and provide
the EFIS-84 system and RMI units with: VOR (magnetic bearing to the station, distance to station (DME), to-from,
course deviation and audio identification signal information), localizer and glideslope deviations, as well as
marker beacon signal and audio signal. The navigation receivers receive the VOR and localizer signal from a pair
of antennas mounted of either side of the vertical stabilizer, DME transmitting and receiving through a pair of
antennas mounted on the lower forward and lower aft portion of the fuselage, and the glideslope signal from an

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 antenna within the nose cone. The selector switch on the right side of the display is pressed up to transfer the
preset and active frequencies and pressed down to cycle the preselect frequency through the 4 user stored
frequencies. The frequency range of the VIR-32 is 108.00 to 117.95 MHZ (VOR and localizer signals), FOR GS and
FOR DME.

Automatic Direction Finder Control (CTL-62)

The aircraft is equipped with one automatic direction finder (autopilot equipped aircraft have two ADF units
installed), located on the lower portion of the center instrument panel and provide the EFIS-84 system and RMI
units with NDB bearing information. The selector switch on the right side of the display is pressed up to transfer
the preset and active frequencies and pressed down to cycle the preselect frequency through the 4 user stored
frequencies.

Radio Magnetic Indicator (RMI-36)

The aircraft is equipped with two radio magnetic indicators, one on each of the pilot’s instrument panels. The
RMI contains a single-needle and a double-needle that displays the gyro-stabilized magnetic heading
information to the respective station on the compass card as selected by the switches at the bottom of the unit.
The single-needle indicator can display information from VOR #1 or ADF #1 and the double-needle indicator can
display information from VOR #2 or ADF #2. The RMI units require 28 VDC and 26 VAC to operate and will
display a red flag in the top right corner of the unit to indicate a loss of power, invalid compass condition and/or
servo error.

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 Note: If only 1 ADF unit is installed then selecting ADF on the single-needle or double-needle will display
information from ADF #1 only.

The compass card gyro-stabilized magnetic heading information is provided by the opposite magnetic compass
unit (MCU-65). This means that the compass card on RMI #1 is driven by MCU #2 and RMI #2 is driven by MCU
#1. This arrangement allows both flight crew members to retain gyro-stabilized magnetic heading information in
the event of either an RMI or EFIS-84 failure.

Transponder (CTL-92)
The aircraft is equipped with one transponder control (CTR-92) and two transponders (TDR-94). The transponder
control unit is located on the lower-center instrument panel and broadcasts the encoding altimeter signal (from
either air data encoding altimeter as selected on the audio panel) through two antennas mounted on the. The
toggle switch on the right side of the display is used to select which transponder is active. The transponder has
the following modes of operation:

OFF – Transponder control unit and transponders are OFF

STBY – Transponder system is ON; however, the transponders transmitting function is OFF

ON – Transponder system is ON (Mode A) and will transmit the 4 digit identification code when an
RADAR interrogation is received.

ALT – Transponder system is ON (Mode C) and will transmit the 4 digit identification code and pressure
altitude when an RADAR interrogation is received.

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 Global Positioning System (KLN 90B)
The aircraft is equipped with one Bendix-King (KLN-90B) Global Positioning System and is located near the
center of the instrument panel, with external annunciators and switches located inboard of each pilot’s Flight
Control Panel.

The panel mounted portion of the KLN-90B is a sensor/navigation computer and contains a removable database
cartridge, CRT display, controls and is connected to the EFIS system and altitude encoders. Also, the single GPS
antenna is mounted on the forward upper portion of the fuselage.

The KLN-90B is IFR approved in the Beechcraft 1900D for enroute, terminal and non-precision approach
operations.

Note: The KLN-90B is capable of performing overlay approaches (VOR, ADF) and stand-alone LNAV GPS
approaches. Unfortunately, due to the age of the system, the KLN-90B cannot perform LNAV+V,
LNAV/VNAV or LPV approaches and must be considered during flight planning.

The KLN-90B is a split screen design where the left dual control knob and CRSR button control the left side of the
display and the right triple control knob control the right side of the display. The 5 pushbuttons below the
screen perform the following functions:

MSG – When the KLN-90B determines that a situation requires the pilot’s attention, a MSG annunciation
appears at the bottom center of the screen. To review the message, press the MSG button and the MSG
page is displayed in the entire screen. If multiple messages appear, they are listed newest at the top to
the oldest at the bottom. Press the MSG button again to cycle to the next page of messages, or if there
are no more messages, it will exit the MSG page.

ALT – Pressing the ALT button will display the ALT page across the screen. This page contains the
barometric entry field and altitude alerter function on the left side of the screen and the Vertical
Navigation (VNAV) page on the right side. The cursor defaults active on the BARO field so that the
current altimeter setting can be changes using the left inner and outer knobs. The right inner and outer
knobs can be used to setup the VNAV profile which is used to plan for descent profiles and produces a
message alert on the MSG page near the required top of descent.

– The Direct To ( ) button is used to initiate direct to operations. When pressed, the Direct
To page will be displayed on the Left side of the screen with a flashing curser over the waypoint
identifier. The waypoint identifier that will automatically appear in the Direct To can be chosen by the
following most common methods:

1) If the cursor is not active and the Direct To button is pressed, the active waypoint is the
waypoint identifier that will appear in the Direct To page.
2) If commencing a GPS approach, once the aircraft has passed the Missed Approach Point
the waypoint identifier that will appear in the Direct To page will be the Missed
Approach Waypoint.
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 3) If the Flight Plan 0 page is displayed on the left side of the screen and the cursor is over
one of the waypoint identifiers when the Direct To button is pressed, this waypoint will
appear in the Direct To page automatically.
4) If the Super NAV 5 page is displayed across the screen, and the right inner knob is in the
“out” position, then the waypoint highlighted in the lower right corner of the display will
become the waypoint identifier that will appear in the Direct To page.

CLR – The CLR button is used to delete waypoint identifiers and flight plans. When the cursor over a
waypoint identifier and the CLR button is pressed, the delete DEL option appears on the screen. When
the ENT button is pressed, the item will be deleted. If CLR is pressed when on a flight plan page FPL with
the cursor not active, the option to delete the entire flight plan appears.

ENT – The ENT button is used to select commands and confirm entries within the KLN-90B system.

In the event of an emergency, the KLN-90B can be used to quickly display the nearest airport that meets the
minimum criteria specific in the SET pages. To access this function, press the MSG button then the ENT button
and use the right inner knob in the “out” position to scroll through the nearest airports.

The various pages types and number of pages are listed below:

Left Screen Selections

Page Annunciation Page Name Number of Pages
TRI Trip Planning 0-6
MOD Mode 1-2
FPL Flight Plan 0-25
NAV Navigation 1-5
CAL Calculator 1-7
STA Status 1-5
SET Setup 0-9
OTH Other 1-4

Right Screen Selections

Page Annunciation Page Name Number of Pages
CTR Center Waypoint 1-2
REF Reference Waypoint None
ACT Active Waypoint Based on FPL 0
D/T Distance/Time 1-4
NAV Navigation 1-5
APT Airport Waypoint 1-5
VOR VOR Waypoint None
NDB NDB Waypoint None
INT Intersection Waypoint None
SUP Supplemental Waypoint None

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Weather Avoidance
Weather Radar Control Panel
The Weather Radar Control Panel provides the sensor data processing, mode control, range selection and other
system controls for the WXR-350 Weather Radar System. The weather radar system provides four color weather
display (green, yellow, red and magenta) to indicate the severity of reflected weather.

W/X RADAR control panel

We have different units in different aircraft (some have WX/tur) and changes the selections below

The functions of the Weather Radar Control Panel buttons and knobs are described below.

MODE switch (Left Rotary Knob)

OFF – The system is not powered

SBY – The system is in Standby mode (warm up, minimum 30 seconds prior to use)

TEST – The system is in TEST mode (Screens ability to display colours and radome sweep)

RANGE switch (Right Rotary Knob) – adjusts range

Changes the range selection on the EHSI ARC or MAP weather modes to: 10, 25, 50, 100, 200 or
300 nautical miles

GAIN switch (Center Left Rotary Knob) – Adjusts power output level from the radome and sensitivity of
the receiver, the higher the gain the higher the output power and sensitivity generating stronger returns

TILT switch (Center Right Rotary Knob) – Adjusts the tilt angle of the radome up or down relative to the
aircraft longitudinal axis.

TGT pushbutton – Warns the pilot of potentially hazardous targets in the path of the aircraft which are
outside of the current selected range and will display TGT on the outer ring in the direction of the hazard

HLD pushbutton – Maintains the picture from the last sweep, it is used by the pilot to see movement of
a cell or system more clearly. After a few minutes, removing the HOLD feature will allow a more
pronounced view of cell movement as a new return is painted.

STB pushbutton – Stabilization. This is normally “on” and allows the radar sweep to maintain reference
to the horizon thereby eliminating unwanted ground returns during banked attitudes or
acceleration/deceleration. A gyroscopic input maintains the level reference.

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Misc Systems
ST3400 Multifunction Display
The Sandel ST3400, located in the center of the instrument panel, is a full colour multi-function display with self-
contained Terrain Awareness Warning System (TAWS) and is connected to a Traffic Collision and Avoidance
System (BF Goodrich TCAS 791). The ST3400 is connected to an external control panel, located above the
ST3400, and remotely mounted TAWS override and test buttons.

Ground Proximity Warning System / Terrain Awareness Warning System

The terrain protection is enabled during all airborne phases of flight: departure, Enroute, terminal and
approach. In addition to the traditional/standard Ground Proximity Warning System (GPWS), the ST3400
contains new enhanced terrain alerts and various advisories that are automatically displayed on the unit
and on external caution and warning annunciators and aural warnings. The difference between the
GPWS modes and TAWS modes is that the TAWS modes rely on the internal database to function,
whereas the GPWS modes uses aircraft sensor inputs only. During enroute flight, the colour coding of
terrain is illustrated below.

The following tables describe the various GPWS and TAWS modes and trigger conditions.

GPWS Alerts
Mode 1 –ERD Uses a combination of barometric and
Excessive Rate of Descent radar altitudes to alert of an excessive
barometric rate of descent when close to
the ground. If radar altitude is not
available, the calculated height above
terrain from the terrain database is
substituted.

Mode 2 - ECRT Uses the radar altimeter to alert of a high

Excessive Closure Rate to rate of change when close to the ground.
Terrain

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 Mode 3 - ALAT Uses a combination of barometric and
Altitude Loss After Takeoff radar altitudes to detect altitude loss after
or Missed Approach takeoff or a missed approach as calculated
from above the takeoff altitude. The
takeoff alert is disabled after climbing 770’
above takeoff elevation; the missed
approach alert is disabled upon reaching
1000’ on the radar altimeter.

Mode 4 - FITNL Uses the radar altimeter to detect too low

Flight Into Terrain Not in of an altitude without the gear and/or flaps
Landing Configuration in the landing configuration.

Mode 5 – EDGSD When in the landing configuration, uses

Excessive Downward the radar altimeter to alert of an excessive
Glideslope Deviation downward deviation from an ILS glideslope
when below 950’.

Mode 6 – VC Uses the radar altimeter to trigger a voice

Altitude Callouts callout when the aircraft descends through
500’ with the gear down.

Mode 6 Enhanced – VC Uses the radar altimeter to trigger

Additional Callouts additional voice callouts at 400, 300, 200,
100, 50, 40, 30, 20 and 10 feet (may be
customized). Also, a voice callout of
“Minimums, Minimums” is triggered by the
DH setting on the pilot’s EFIS system.

TAWS Alerts
FLTA Looks ahead of the aircraft, lateral and
Forward Looking Terrain vertical flight paths are compared to the
Avoidance terrain database to provide an alert if a
potential terrain threat exists.

PDA The current vertical and lateral position of

Premature Decent Alert the aircraft is compared to the proximity
of the nearest airport to determine is the
aircraft is abnormally below a reasonable
altitude.

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 Traffic Collision and Avoidance System (TCAS I)
The ST3400 is used as the display for the BF Goodrich TCAS 791 system. The TCAS 791 system
consists of the ST3400 display, TCAS computer, transmitting antenna (mounted on the forward
upper portion of the fuselage), receiving antenna (mounted on aft lower portion of the fuselage)
and is connected to various other aircraft systems.

The TCAS I system provides visual and aural traffic advisories only; any corrective action must be
determined and initiated by the flight crew. Below is the traffic symbology used on the ST3400.

Traffic Collision and Avoidance System (TCAS II)

Some of the aircraft are equipped with Bendix-King CAS-67A TCAS II system. The CAS-67A system
consists of a digital Vertical Speed Indicator (VSI)/TCAS display, TCAS computer, transponder/TCAS
control unit, transmitting antenna (mounted on the forward upper portion of the fuselage), receiving
antenna (mounted on aft lower portion of the fuselage) and is connected to various other aircraft
systems.

The TCAS II system provides visual advisories and visual/aural resolution advisories. The VSI/TCAS
display will display traffic advisories, and if required, visual and aural resolution advisories that provide
the required vertical flight path action required to clear the conflict.

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 Static Discharge Wicks
During flight, a static electrical charge may build up on the surface of the aircraft and can cause interference
with the avionics and operation of other electrical equipment. To keep this static charge to a minimum, static
discharge wicks are installed on the trailing edges of the flight surfaces and continuously dissipate the static
charge back into the atmosphere.

There is a total of 24 Static wicks and are installed at the following locations: 2 on each winglet,3 on each
aileron, 3 on the elevator, 4 on the rudder, 1 on each tail-let, 1 on each horizontal stabilizer tip, 1 on the top of
the horizontal stabilizer and 1 on each ventral strake.

System Tests
Overspeed Test
The overspeed warning system (aural warning) is tested by holding the combination stall warning/overspeed
warning test switch to OVERSPEED WARNING TEST position and verify the appropriate stall warning sound is
heard.

SOP – Overspeed Warning System Test

Standby Attitude Indicator

The Standby Attitude Indicator system is checked to ensure proper operation of the attitude indicator,
independent battery and indication/warning system.

On the Standby Attitude Indicator control panel, hold the switch to the TEST position (for a maximum of 5
seconds) and verify the green “AUX TEST” annunciator illuminates. Select the ON position and verify the yellow
“AUX ON” annunciator illuminates and the aural warning is heard (beeping tone), then press the small red HORN
SILIENCE button and the aural warning should silence. Uncage the standby attitude indicator gyro and verify that
the red flag moves out of view, then select the switch to the OFF position and verify that the red flag move into
view again. Cage the standby attitude indicator gyro and verify that the yellow “AUX ON” annunciator is
extinguished.

SOP – Standby Attitude Indicator Check

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EFIS Aux Power
The EFIS Aux Power system is checked to ensure proper operation of the independent battery and
indication/warning system.

On the EFIS Aux Power control panel, hold the switch to the TEST position (for a maximum of 5 seconds) and
verify the green “AUX TEST” annunciator illuminates. Select the ON position and verify the yellow “AUX ON”
annunciator illuminates and the aural warning is heard (continuous tone), then press the small red HORN
SILIENCE button and the aural warning should silence. Select the switch to the OFF position and verify that the
yellow “AUX ON” annunciator extinguishes.

SOP – EFIS Aux Power Check

Ground Proximity Warning System

The Ground Proximity Warning System is tested to ensure proper operation and indication. The GPWS system in
internal to the ST3400, located in the lower-center of the instrument panel, and has a remote GPWS control
panel located on the top-left of the instrument panel, as well as a flap override switch below the engine
instruments.

With the flaps at 0°, press the GPWS Test Switch and observe the following: The GPWS “FLAP OVRD”, “INOP”
and “GS CANCLD” annunciators illuminate during the test and the “BELOW GS” annunciator illuminates the aural
“Glideslope” warning is heard, and then the aural “Pull-UP” warning is heard. The test is completed when the
GPWS annunciators extinguish.

SOP – Ground Proximity Warning System Test

Altitude Alerter
The Altitude Alerter is checked to ensure that the system operates correctly and provides an aural warning.

Set the pilot’s altimeter to the nearest 100’ and set the same altitude in the Altitude Alerter. Change the pilot’s
altimeter setting to increase the altitude 300’ and verify that the aural warning chime activates and the yellow
“ALT” annunciator is blinking. Reset the pilot’s altimeter setting to the altitude on the Altitude Alerter and verify
the aural warning chime is silent and yellow “ALT” annunciator is extinguished. Change the pilot’s altimeter
setting to decrease the altitude 300’ and verify that the aural warning chime activates and the yellow “ALT”
annunciator is blinking. Set the current altimeter setting on the pilot’s altimeter and verify that the aural
warning chime is silent and yellow “ALT” annunciator is extinguished.

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 SOP – Altitude Alerter Check

EFIS Test
Both EFIS systems are tested to ensure proper function and annunciation. The EFIS test buttons are located to
the right of the EFIS power switches, at the top of the lower pedestal.

To test the pilot’s EFIS system, press and hold the PLT button and verify that initially the pilot’s EADI indicates
10° pitch up and right roll, pilot’s EHSI indicates a 20° heading increase and the EADI displays a red TEST
annunciator in the upper left corner. After 2-3 seconds, both the pilot’s and co-pilots EADIs display amber “PIT”
and “ROL” annunciators, both EHSIs displays an amber “HDG” annunciator and the yellow “PUSH TO COMPARE”
annunciator is illuminated. After 5 seconds the pilot’s EADIs attitude is removed and displays red “ATT”,
“XDATA”, “RA” and “DSP” annunciators blink to the 10 second mark and then remain on, as well as the pilot’s
EHSI displays red “HDG”, “XDATA” and “DSP” annunciators that blink to the 10 second mark and then remain on.
Now release the PLT test button, press the yellow “PUSH TO COMPARE” annunciator button and the all the EFIS
screens should return to normal, with the yellow “PUSH TO COMPARE” annunciator extinguished.

The co-pilot’s EFIS systems is tested the same way, except that initially the co-pilot’s EADI indicate 10° pitch
down and roll left and the co-pilot’s EHSI indicates a 20° heading decrease.

SOP – EFIS Test

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 Radios Test

Each of the communications receivers, navigation receivers and ADF receivers are tested to ensure the internal
fault monitoring systems do not detect any faults.

Press the TEST button on each unit and observe the two-digit code displayed on the unit. If the unit displays a
code other than 00 (no trouble found) or 05 (invalid NAV frequency = station out of range) then it should be
reported to maintenance for repair.

SOP – Radios Test

TAWS / TCAS
The TAWS / TCAS system is tested to insure that the TCAS internal system test passes and the altitude readout is
correct. The TCAS I control panel is located directly above the ST3400 unit and if installed, the TCAS II unit in
integrated in the transponder control unit.

TCAS I
Press the TCAS button and verify that the green “ON” annunciator illuminates. Set the pilot’s altimeter
to standard (29.92) and press the TCAS TEST/ALT button. Verify that the altitude displayed on the TCAS
screen is within 100’ of the pilot’s altimeter and that the test screen is shown. After a short time, the
TCAS system will announce “TCAS Test Passed” through the audio system. Reset the Pilot’s altimeter to
the current altimeter setting.

TCAS II
Rotate the control knob on the transponder control unit to the TST position for a moment, and then
back to SBY. The VSI/TCAS displays will show the test screen and after a short time, the TCAS system will
announce “TCAS Test OK” through the audio system.

SOP- TAWS / TCAS Test

Transponder
The transponder is tested to ensure the internal fault monitoring systems do not detect any faults.

Press the TEST button on the unit and verify that the display indicates the current pressure altitude (3 digits) and
then returns normal display that shows the 4 digit transponder code. If the unit displays an error code instead of
the pressure altitude , reference error code list at the end of this chapter for any faults or abnormalities
detected.

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 SOP- Transponder Test

EFIS reversionary
The EFIS reversionary system is tested to ensure proper operation or all the reversionary modes.

Select the PLT CMPST switch to CMPST and verify that both pilot EFIS screens display the composite screen
mode. Select the PLT CMPST switch to OFF.

Select the DSP switch to PLT and verify that the yellow “DSP 1” annunciator is displayed on the lower right
corner of the co-pilot’s EADI and the radar altimeter portion is now yellow and all data on the EHSI is now yellow
(when cross-sided data is selected (or GPS selected). Select the DSP switch to COPLT and verify that the yellow
“DSP 2” annunciator is displayed on the lower right corner of the pilot’s EADI and the radar altimeter portion is
now yellow and all data on the EHSI is now yellow (when cross-sided data is selected (or GPS selected). Select
the DSP switch to NORM.

Select the DR XFR switch to PLT and verify that the yellow “ATT 1” is displayed on both EADIs and “MAG 1” on
both EHSIs as well as the amber “DRIVE XFR” annunciator on the co-pilots side is illuminated. Also, the co-pilot’s
flight director is removed and the FCP goes blank. Select the DR XFR switch to COPLT and verify that the yellow
“ATT 2” is displayed on both EADIs and “MAG 2” on both EHSIs as well as the amber “DRIVE XFR” annunciator on
the pilot’s side is illuminated. Also, the plot’s flight director is removed and the FCP goes blank Select the DR XFR
switch to NORM.

Select the COPLT CMPST switch to CMPST and verify that both co-pilot EFIS screens display the composite screen
mode. Select the COPLT CMPST switch to OFF.

Select the ATT switch to ALL ON NO.1 and verify that the yellow “ATT 1” annunciator is displayed on the lower
right corner of both EADIs. Select the ATT switch to ALL ON NO.2 and verify that the yellow “ATT 2” annunciator
is displayed on the lower right corner of both EADIs. Select the ATT switch to NORM

Select the HDG switch to ALL ON NO.1 and verify that the yellow “HDG 1” annunciator is displayed on the top
center of both EHSIs. Select the HDG switch to ALL ON NO.2 and verify that the yellow “HDG 2” annunciator is
displayed on the top center of both EHSIs. Select the HDG switch to NORM.

SOP – EFIS Reversionary Test

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RADAR
The RADAR system is tested to ensure that the EFIS system is receiving data from the RADAR system and the
internal fault monitoring system does not detect any faults.

Select WX ARC or WX MAP on the pilot and co-pilot DSPs, select the range on the RADAR control panel to 25NM
and set the RADAR to the TEST position. Verify the test rainbow is displayed correctly on both EHSI screens, then
set the RADAR mode to STBY.

SOP – Radar Test

Briefer
The briefer is tested to ensure the system is operating and check the volume.

Select and play any briefing and verify that the playback is audible and volume is set so that the passengers will
be able to hear the recording with engine running.

SOP - Briefer

Cockpit Voice Recorder

The Cockpit Voice Recorder (CVR) is tested to ensure the recording and playback functions are operating
correctly.

Insert a headset speaker plug into the CVR plug and place the headset on your head. Hold the green TEST button
(minimum of 5 seconds), listen for a tone verify that the CVR meter increases into the green range. Speak
normally and listen for playback from the CVR system in the headset. Hold the red ERASE button (2 seconds
minimum) to erase the previous recording.

SOP – Cockpit Voice Recorder

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Flight Director
The Flight Control panel self test feature is used to ensure proper system operation and that the internal fault
monitoring system does not detect any faults.

Press the red TEST button and the FCP annunciators will test quickly (blink once) and then the diagnostics test is
started. When complete, a yellow “TEST” annunciation is displayed on the top-center of the respective EADI. If
the GA annunciator is the only annunciator remaining illuminated after the test, no faults were found. However,
if there are other annunciators that remain illuminated, a fault has been detected and all the illuminated
annunciators (except for GA) should be reported to maintenance for repair. To exit the test mode, press the red
TEST button again.

SOP – Flight Director Check

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