BELL 206
FIELD MAINTENANCE TRAINING
Bell 206 – Field Maintenance Training
INDEX:
DESCRIPTION
PAGE
AIRCRAFT GENERAL
03
PNEUMATIC/ENVIRONMENTAL SYSTEMS
71
ELECTRICAL SYSTEMS
83
INSTRUMENT/AVIONICS SYSTEM
130
FUEL SYSTEM
159
HYDRAULIC SYSTEM
173
MAIN ROTOR SYSTEM
183
MAIN ROTOR DRIVE SYSTEM
196
TAIL ROTOR SYSTEM
214
TAIL ROTOR DRIVE SYSTEM
224
FLIGHT CONTROL SYSTEM
241
POWERPLANT
280
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AIRCRAFT
GENERAL
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INTRODUCTION
AIRFRAME
The 206B is a single engine, five place, utility type helicopter. Average empty weight is 1,604 pounds,
and the maximum internal gross weight is 3,200 pounds.
The standard configuration provides seating for one pilot and four passengers. The passengers seats
and doors can be removed to utilize the aircraft in a utility configuration. Removal of the seats in the
cabin area provides approximately 40 cubic feet of cargo area, with a maximum floor loading of 75
pounds per square foot. An optional cargo hook can be attached to the bottom of the aircraft to
accommodate external cargo operations. Also, the aircraft has a 16 cubic foot baggage compartment
with a maximum floor loading of 86 pounds per square foot.
The aircraft is powered by Allison turboshaft engine Model 250C-20J, rated at 420 SHP. The
transmission is limited to 317 SHP for takeoff, and 270 SHP continuous operation.
The main rotor, which employs preconing and underslinging to ensure smooth operation, is a semirigid,
seesaw type with two rotor blades. Each rotor blade is attached to a common hub by means of a grip,
pitch change bearings, and a tension-torsion strap assembly to carry the blade centrifugal force.
The tail rotor is a semirigid, delta hinged, two blade design.
Several different skid type landing gear configurations are available to better equip the aircraft for a
particular type operation.
AIRFRAME SECTIONS
The 206B is divided into the three major sections: forward, intermediate, and tailboom.
The forward section of the aircraft begins at the nose of the aircraft and extends aft to the rear landing
gear attachment point. Included in the forward section of the aircraft is the main rotor, transmission,
cabin, pilot and passenger doors, fuel cells, and the landing gear. The forward section is composed
primarily of aluminum honeycomb structure. This structure was utilized due to both excellent weight to
strength ratio and the ability to dampen noise and vibrations.
The intermediate section includes the engine, the equipment/electrical shelf, the baggage compartment,
and oil cooler assembly. It consists of semi-monocoque construction, and the aluminum skins are
various thickness to accommodate the different dynamic loads.
The tailboom section, which is a full monocoque structure, includes the tail rotor drive shafts, horizontal
stabilizer, vertical fin, tail rotor gearbox and the tail rotor.
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AIRCRAFT GENERAL
6-1. DIMENSIONS AND CHARTS.
Principal dimensions and station locations of the helicopter are included in this chapter. Metric
dimensions and station locations are included for reference.
6-2. PRINCIPAL DIMENSIONS.
Figure 6-1 illustrates the principal dimensions of the helicopter and includes a three view general
arrangement drawing.
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6-3. STATION DIAGRAM
Figure 6-2 illustrates the principal fuselage stations (STA), butt lines (BL), waterlines (WL), centerlines
(CL), and boom stations (BS) of the helicopter.
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LIFTING AND JACKING
7-1. LIFTING AND JACKING
Procedures for lifting a complete helicopter, lifting a helicopter minus a transmission, lifting a tailboom
only, and jacking a helicopter are provided in this chapter.
7-2. LIFTING COMPLETE HELICOPTER
WARNING
PERSONNEL SHALL NOT ENTER INTO OR CLIMB ONTO HELICOPTER WHILE IT IS BEING
RAISED OR WHILE SUPPORTED ON JACKS. AREA SHALL BE ROPED OFF AND WARNING SIGNS
DISPLAYED, 'THIS HELICOPTER IS ON JACKS.'
1. Secure main rotor blades with main rotor tie-down (3, Figure 7-1) and tail rotor with tail rotor strap (4).
2. Attach suitable hoisting cable and clevis (1) capable of supporting helicopter to eye of main rotor mast
nut (2). Connect suitable hoist and take up slack.
3. Position a person at tail skid to steady helicopter when hoisting. If lifting of helicopter will be beyond
arms reach from ground, secure a safety rope to tail skid.
4. Hoist helicopter slowly with constant lifting force.
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7-3. LIFTING HELICOPTER WITHOUT TRANSMISSION
1. Remove all loose equipment from cabin and passenger compartment.
2. Remove all loose or broken hardware (i.e., cowling, fairing, tailboom, etc.).
3. If landing gear crosstubes are intact, provide a four-point sling from crosstubes on each side of
fuselage to a common clevis. Use a standard four-point sling to provide lifting capability of a minimum of
5000 pounds (2268 kg) total. Protect with padding where sling may contact fuselage.
4. If lifting of helicopter cannot be accomplished using a hoisting clevis or a four-point sling, a large cargo
net may be used. Place cargo net under fuselage and attach four corners of net to hoisting clevis.
5. Secure safety rope to toe of one landing gear skid or tie-down fitting.
6. Hoist helicopter slowly with constant lifting force.
7-4. LIFTING TAILBOOM ONLY
Hoist tailboom as separate component with sling positioned at center of gravity of tailboom. Use sling
manufactured from web material. Position a person at each end of tailboom to steady tailboom in sling.
7-5. JACKING HELICOPTER
WARNING
PERSONNEL SHALL NOT ENTER INTO OR CLIMB ONTO HELICOPTER WHILE IT IS BEING
RAISED OR WHILE SUPPORTED ON JACKS. AREA SHALL BE ROPED OFF AND WARNING SIGNS
DISPLAYED, 'THIS HELICOPTER IS ON JACKS'.
1. When conditions require that helicopter be placed on jacks in an unsheltered area, the following
precautions shall be observed.
a. Secure main rotor blades with main rotor tie-down (3, Figure 7-1) and tail rotor with tail rotor
strap (4).
b. Place three 8587 or equivalent hydraulic jacks (6) in their respective positions under aft and
forward jack pads (5 and 7).
c. Operate jack handles slowly and carefully, simultaneously at three positions, being careful to
keep helicopter level, as it is raised to desired height.
2. When lowering helicopter, slowly and carefully lower the three jacks simultaneously.
3. Remove jacks and other equipment from area.
4. Return and secure loose equipment taken from cabin and passenger compartment.
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WEIGHT AND BALANCE
8-1. PURPOSE
This section gives the procedures that are necessary to find the actual weight empty and the center of
gravity (CG) of a helicopter configuration, and to find what changes, if any, are necessary to keep the
helicopter within the gross weight flight limits during operation.
8-2. GENERAL
The CG is the balance point of a body and is used when calculating the weight and balance for the
helicopter. If a helicopter and pendulum are compared, the suspension point is where the main rotor hub
intersects the mast and the pendulum weight is the helicopter. The pendulum weight will stop with its CG
directly below the suspension point. For example, a helicopter will fly with its nose up if the CG is aft of
the hub/mast intersection. To fly the helicopter in a level manner, the pilot must move the cyclic control
stick forward. The more the pilot moves the cyclic control stick forward, the less power there will be for
forward speed and the control over the helicopter is decreased. Because this loss of control is unsafe, it
is important to keep the helicopter CG within the given gross weight flight limits. This is done in two
ways:
1. By changing the location of the helicopter weight empty CG through equipment relocation or by
adding or removing ballast, and
2. By deriving the combinations of useful load items which are permitted for each flight.
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LEVELING
8-4. LEVELING PROCEDURE
Helicopters 4 through 103 and 584 and subsequent can be leveled using the plumb bob method
(paragraph 8-5) or by alternate procedure using a spirit level and target leveling plate mounted on the
cabin floor (paragraph 8-6).
Helicopters 104 through 583 are leveled using a spirit level positioned on leveling pads located on aft
portion of cabin roof adjacent to pylon support (paragraph 8-7).
8-5. LEVELING PROCEDURE USING PLUMB BOB METHOD
NOTE
It will be necessary to loosen or remove upper upholstery at location of slotted plate (1, Figure 8-1) and a
portion of carpeting around leveling plate (3).
NOTE
A slotted plate (3) is located on the cabin roof approximately 14 inches (355.60 mm) inboard from edge
of passenger-cargo door at fuselage station 90.00, buttline -11.14 (Approximately 7.00 inches (177.80
mm) forward of aft seat structure). The slotted plate (1) is located directly above leveling plate (3). 1.
Hang a plumb bob (2) from slotted plate (1) in cabin roof. Plumb bob should be just above leveling
plate (3).
CAUTION
HELICOPTER MUST BE ON HARD LEVEL SURFACE PRIOR TO JACKING HELICOPTER.
2. Place two forward jacks (5) under, but clear of forward jack fittings (4). Place the aft jack (5) under aft
jack fitting (6).
3. Adjust the aft jack (5) until the helicopter is almost level.
4. Adjust forward jacks (5) until snug against forward jack fittings (4). Raise all three jacks evenly until
skids are clear of surface.
5. Level helicopter fore and aft and laterally by adjusting height of jacks (5) at forward and aft jack fittings
(4 and 6) while observing plumb bob (2). Helicopter is level when plumb bob is directly over intersection
of lines of leveling plate (3).
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8-6. ALTERNATE LEVELING METHOD (HELICOPTERS S/N 4 THROUGH 103 AND 584 AND
SUBSEQUENT)
1. Mount large leveling plate (8, Figure 8-1, detail A) on cabin floor approximately 14 inches (355.60 mm)
inboard from edge of passenger-cargo door at fuselage station 90.00, and buttline 11.14 (approximately
7.00 inches (177.80 mm) forward of aft seat structure).
NOTE
During this step ensure that forward jacks (5) are clear of forward jack fittings (4) to prevent side loads to
helicopter. 2. Place the two forward jacks (5) under, but clear of the forward jack fittings (4). Place the aft
jack (5) under the aft jack fitting (6).
3. Adjust the aft jack (5) until helicopter is almost level.
4. Adjust forward jacks (5) until snug against forward jack fittings (4). Raise all three jacks evenly until
skids are clear of surface.
5. Position spirit level (7) on leveling plate (8).
6. Level helicopter fore and aft and laterally by adjusting height of jacks (5) at forward and aft jack fittings
(4 and 6) while observing spirit level (7) indication. Helicopter is level when spirit level (bubble) is in
center of spirit level both longitudinally and laterally.
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8-7. LEVELING USING LEVELING PAD METHOD (HELICOPTERS 104 THROUGH 583)
NOTE
Helicopters S/N 104 through 583 incorporate four leveling pads located on aft right portion of cabin roof
adjacent to pylon support.
1. Open access door adjacent to firewall on right side of forward cowling.
2. Place spirit level across two lateral pads.
3. Level helicopter laterally by adjusting height of jacks (5, Figure 8-1) at forward jack fittings (4 and 6)
while observing spirit level indication. Helicopter is level when spirit level (bubble) is centered laterally.
4. Place spirit level device across fore and aft leveling pads.
NOTE
No correlation of leveling provision is required between lateral and fore and aft pads.
5. Level helicopter fore and aft and laterally by adjusting height of jacks (5) at forward and aft jack fittings
(4 and 6) while observing spirit level indication. Helicopter is level when spirit level (bubble) is centered
both longitudinally and laterally.
8-8. WEIGHING PROCEDURE
8-9. PREPARATION OF THE HELICOPTER FOR WEIGHING
Before weighing the helicopter, ensure configuration is as near the Weight Empty as possible. Perform
the following:
1. Remove, as much as possible, dirt, grease, moisture, and any equipment not required for weighing
from helicopter.
2. Ensure baggage compartment is empty.
3. Place all kits and required equipment for weighing in proper locations.
4. Ensure transmission, gearbox, and hydraulic reservoirs are serviced to proper levels (Chapter 12).
5. Ensure engine oil system is either fully drained or topped up to the full mark.
NOTE
The Weight Empty configuration is the weight of the basic helicopter plus the weight of the kits, special
equipment, fixed ballast, transmission and gearbox oil, hydraulic fluid, unusable fuel, and undrainable oil.
6. Drain fuel system (Chapter 12).
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8-10. WEIGHING
1. Do not weigh the helicopter outdoors or in an open building because wind, flapping rotors, and body
sway may seriously affect the accuracy of scale readings.
NOTE
If electronic platform scales are used, always align the jackpoint on center of the scale. Do not cross
scale coax wire on ground or put any weight on cable.
2. If electronic loadcells are used, ensure that loadcells and adapters are tightened to pads of the jacks
used to raise the helicopter. Place loadcells in position on jacks. Perform warmup recommended by the
scale manufacturer. Refer to instructions supplied by the manufacturer and adjust each loadcell to zero.
3. Ensure each of the scale calibrations have a zero reading before performing each weighing
procedure.
CAUTION
ENSURE LANDING GEAR SKIDS DO NOT TOUCH THE WEIGH SCALES OR FLOOR. IF SKIDS
TOUCH, THE SCALES WILL NOT BE BALANCED CORRECTLY.
4. Weigh helicopter on portable scales and place scales in position on level ground. Place a scale under
each jackpoint. Align jackpoint on center of scale. Use jacks to make helicopter level in longitudinal and
lateral directions (paragraph 8-4).
5. Balance each scale and make a note of the readings. If electronic scales are used, find the weight
on each cell from the digital counter. Refer to manufacturer instructions.
6. Remove the helicopter from jacks (Chapter 7). On each scale, weigh the weight tare. This includes the
applicable jack, blocks, and any other equipment in position between the helicopter and scale. Subtract
this weight tare from the first scale reading to get the net weights.
7. To ensure accurate readings are obtained, rotate loadcells/electric scales one position and reweigh
the helicopter. If the difference between the first total weight and second total weight is less than 10.0
pounds (4.54 Kg), and if forward weight and aft weight difference between the first and second
weighings (ie. [left+right -aft] - [left+right-aft]) is less than 5.0 pounds (2.268 Kg) the resultant readings
can be considered accurate.
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9-1. TOWING
The helicopter can be equipped for towing by attaching two ground handling wheels (4, Figure 9-1) to
landing gear skid tubes (3). A standard tow bar (1) may be attached to tow rings (2) provided on forward
end of each landing gear skid tube.
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9-2. TOWING PROCEDURES
1. Secure main rotor blades (Chapter 10).
2. Attach tow bar (1) to skid gear tow rings (2) and to towing vehicle. If helicopter is moved by hand, do
not push on any part of airframe that could result in damage to helicopter, i.e., vertical fin, elevator, etc.
CAUTION
DO NOT USE TAIL ROTOR TO HANDLE HELICOPTER.
3. Install and extend ground handling wheels (4) (Chapter 32).
CAUTION
WHEN TOWING HELICOPTER OBSERVE OBSTRUCTIONS TO PREVENT DAMAGE TO ROTOR
BLADES AND OTHER HELICOPTER PARTS.
4. Clear departure area of auxiliary support equipment, i.e., workstands, power units, etc.
CAUTION
TOWING HELICOPTER ON GROUND, UNPREPARED SURFACES, OR ACROSS HANGAR DOOR
TRACKS AT EXCESSIVE SPEED MAY CAUSE A PERMANENT DEFLECTION IN AFT CROSSTUBE.
NOTE
To guide helicopter and balance it on ground handling wheels during towing, position one person at tail
skid. On helicopters S/N 4 through 413, a receptacle grip is provided on the right side of vertical fin.
5. Tow or push slowly, balancing helicopter with tail skid (5).
WARNING
ENSURE THAT FEET ARE CLEAR OF LANDING GEAR SKID TUBE PRIOR TO LOWERING
HELICOPTER.
CAUTION
GRASP GROUND HANDLING WHEEL BAR FIRMLY PRIOR TO REMOVING PIN. DO NOT LEAVE
HELICOPTER UNATTENDED WITH GROUND HANDLING WHEELS IN EXTENDED POSITION.
6. Remove pin on ground handling wheels (4), slowly lower helicopter, and remove ground handling
wheels.
7. Remove tow bar (1).
8. Remove main rotor tie-down from main rotor (Chapter 10).
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PARKING AND MOORING
10-1. COVERS AND TIE-DOWNS
Protective covers and tie-downs are furnished as loose equipment and are used for the parking or
mooring of the helicopter. Additional equipment such as ropes, cables, clevises, ramp tie-downs or dead
man tie-downs will be required during mooring.
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10-2. COVERS — ENGINE INLET AND PITOT TUBE
Engine inlet plug assembly (4, Figure 10-1) and pitot tube cover (7) are red in color and flame resistant.
Each cover has a red streamer stenciled in white letters, REMOVE BEFORE FLIGHT. Install engine inlet
plug in each engine air inlet so that surface marked TOP is up. Cover pitot tube and tie cord to secure.
10-3. COVER — ENGINE EXHAUST
Engine exhaust covers (3, Figure 10-1) are red in color, flame resistant, and include a red streamer
stenciled in white letters, REMOVE BEFORE FLIGHT. Install engine exhaust cover on each exhaust.
10-4. TIE-DOWN — TAIL ROTOR
CAUTION
DO NOT TIE DOWN ROTOR TO THE EXTENT THAT TAIL ROTOR BLADE BECOMES FLEXED.
1. Secure tail rotor strap (2, Figure 10-1) to vertical fin.
2. Position tail rotor in horizontal position and secure tail rotor strap (2) to lower position of tail rotor
blade.
10-5. TIE-DOWN — MAIN ROTOR
CAUTION
DO NOT TIE DOWN ROTOR TO THE EXTENT THAT MAIN ROTOR BLADE BECOMES FLEXED.
1. Secure main rotor tie-down (1, Figure 10-1) to eyelet on rotor blade or place boot assembly (8) on
main rotor blade.
2. Pull main rotor tie-down (1) tight and secure to right side of tailboom or tie straps of boot assembly
around tailboom.
10-6. PARKING — NORMAL AND TURBULENT CONDITIONS (WINDS UP TO 50 KNOTS)
Park helicopter for normal and turbulent conditions with winds up to 50 knots in accordance with
procedures described in this paragraph. For conditions with winds above 50 knots, moor helicopter
(paragraph 10-7).
1. Position helicopter in desired parking area by hovering, taxiing, or towing (Chapter 9). Allow helicopter
to rest on landing gear skid tubes.
2. Secure main and tail rotor blades if helicopter is parked in an area subject to turbulence created by jet,
prop or rotor blast from other aircraft (paragraph 10-4 and paragraph 10-5).
CAUTION
MAXIMUM ALLOWABLE LOAD AT MAIN ROTOR BLADE TIPS IS 100 POUNDS (45 KG).
3. Install engine inlet plug assemblies (4, Figure 10-1), pitot tube cover, and engine exhaust covers
(paragraph 10-2 and paragraph 10-3).
4. Tighten friction locks on flight controls, check that all switches are in the OFF position, and disconnect
battery.
5. Close and secure all doors, windows, and access panels.
6. If helicopter is parked outside in a heavy dew environment, purge lubricate all exposed control
bearings every 7 days to ensure no voids exist that could trap moisture (Chapter 12).
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10-7. MOORING (WINDS ABOVE 50 KNOTS)
If helicopter must be parked in the open during period of high wind forecast, comply with the following
precautionary measures:
CAUTION
STRUCTURAL DAMAGE CAN OCCUR FROM FLYING OBJECTS DURING HIGH WIND
CONDITIONS. HELICOPTER SHOULD BE HANGARED OR EVACUATED TO A SAFE WEATHER
AREA WHEN WIND CONDITIONS ABOVE 75 KNOTS ARE EXPECTED.
1. If a paved ramp with suitable tie-down rings is available, position helicopter with nose headed in
direction from which highest forecast winds are expected. Secure helicopter to ramp tie-downs. Use
cable, rope, or manufactured tie-downs at helicopter jacking tie-down fittings. Use of a mooring clevis at
each of the three tie-down fittings will permit use of larger diameter rope.
2. If suitable tie-down rings are not available, park helicopter on an unpaved parking area with nose
headed in the direction from which highest forecast winds are expected and retract ground handling
wheels. Use dead man tie-downs.
3. Secure main and tail rotor blades (paragraph 10-4 and paragraph 10-5). If storage space and time are
available, remove main rotor blades (Chapter 62) and store in a secure building. Secure main rotor hub
to mast to prevent movement on flapping axis.
CAUTION
MAXIMUM ALLOWABLE LOAD AT MAIN ROTOR BLADE TIPS IS 100 POUNDS (45 KG).
4. Install engine inlet plug assemblies (4, Figure 10-1) and pitot tube cover (7). Secure red streamers
inside nearest access doors to prevent flapping.
5. Tighten friction locks on flight controls. Check that all switches are in OFF position. Disconnect
battery.
6. Close and secure all doors, windows, and access panels.
7. Fill fuel tank to maximum capacity with prescribed fuel (Chapter 12).
8. Secure all ground handling equipment and other objects which might be blown by high winds.
NOTE
After winds subside, inspect helicopter carefully for damage which may have been inflicted by flying
objects.
10-8. STORAGE
Preparation procedures to place the helicopter in storage and depreservation procedures to activate the
helicopter after storage are found in BHT-ALL-SPM. Storage of helicopter includes corrosion control,
which consists primarily of preventing moisture from contacting exposed material surfaces by the use of
preservatives. Refer to CSSD-PSE-87-001 Corrosion Control Guide.
Prior to returning helicopter to service, perform applicable Depreservation and Activation procedure
(BHT-ALL-SPM).
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11 - PLACARDS AND MARKINGS
Placards, markings, and stencils applicable to the helicopter are illustrated in Figure 11-1.
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SERVICING
12-1. SERVICING — GENERAL
This chapter contains the data required to service the helicopter.
Table 12-1 contains a list of applicable helicopters with their associated serial numbers and Flight
Manual document.
For data on the consumable materials, refer to Standard Practices Manual (BHT-ALL-SPM).
CAUTION
THE SERVICING INTERVALS PROVIDED IN THIS CHAPTER ARE THE MAXIMUM PERMITTED
INTERVALS UNDER NORMAL OPERATION OF THE HELICOPTER. DO NOT EXCEED THESE
INTERVALS. IT MAY BE NECESSARY TO DECREASE THESE INTERVALS IF THE HELICOPTER
OPERATES IN EXTREME ENVIRONMENTAL CONDITIONS.
For specific information on fuels, oils and fluids, refer as required to the applicable Flight Manual as
follows:
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12-2. ACCESS PANELS, DRAINS, AND VENTS
NOTE
Boost pumps and lower fuel quantity probe covers were introduced as basic equipment at S/N 3567.
Figure 12-1 shows the location of all the access panels, drains and vents required for servicing the
helicopter. For the location of sight and level gauges, refer to specific system Figure 12-3 or Figure 12-4
as required.
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FUEL SYSTEM — SERVICING
The 206A/B series helicopter fuel system incorporates a single bladder type fuel cell located below and
aft of passenger seat. Installed within the fuel cell are two electrically operated boost pumps, lower and
upper tank indicating units and sump drain valve. Some helicopters are also equipped with a low fuel
level switch mounted on the drain valve. Boost pumps are interconnected and supply fuel through a
single hose assembly to the fuel shut-off valve, and from there to the engine mounted fuel filter and
pump assembly.
Several helicopters incorporate an airframe-mounted fuel filter between the shut-off valve and the engine
fuel pump. The airframe fuel filter is installed on the aft face of the right hand side of the forward firewall.
The fuel filter has a manual drain valve, bypass capability and an impending bypass switch. The switch
is connected to the caution/warning and advisory panel in the flight compartment. Boost pumps
incorporate pressure switches in discharge ports and drain plugs in the pump drain port. The fuel cell is
filled from the right side. A ground jack is also installed on the right hand side of the helicopter near the
filler port. Access to boost pumps, lower tank unit and drain valve is from bottom of the fuselage. Access
to the upper indicating unit is gained from a cover plate located on the hat bin shelf. Access to fuel shutoff valve compartment and vent line is gained from an access door located on right side above filler cap.
Provisions are also made in the shut-off valve compartment for combustion heater fuel, fuel pressure
instrument line and fuel pump purge line (S/N 4 through 2123 only).
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OIL LEVEL — CHECK
The sight gauges in the transmission, the engine oil tank and the tail rotor gearbox permit personnel to
ensure the oil level in these components is within the specified limits. Refer to Figure 12-3 for the
location of the sight gauges. Stained or discolored sight gauges could give a false indication of oil/fluid
quantity. Clean or replace stained or discolored sight gauges.
ENGINE OIL SYSTEM
The engine oil system is serviced by removing the oil filler cap installed on the engine oil tank. The oil
tank is installed aft of the rear engine firewall on the roof of the helicopter and closed by the aft fairing
(Chapter 53).
TRANSMISSION OIL SYSTEM
Table 12-3 shows the recommended servicing intervals. The 206A/B series transmission has a vented
filler cap located on the top right side of the case. A sight glass is located on the lower right side of the
transmission case indicates the oil level. The oil filter assembly, mounted on the lower left side of the
transmission case, has a filter head that contains a bypass valve, a temperature-warning switch and a
temperature bulb. The transmission has two electric chip detectors in the main case and may have one
in the top case to monitor the mast bearing. A transmission oil cooler is mounted at the back of the top
case and is fed cooling air from the oil cooler fan shaft assembly. Lubrication to the freewheeling unit is
provided by the transmission oil system. The transmission oil system is serviced through the oil filler cap
installed on the transmission.
FREEWHEEL OIL SYSTEM
The freewheel oil system shares lubrication oil with the transmission oil system. The freewheel oil
system is serviced through the transmission oil filler cap.
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HYDRAULIC SYSTEM SERVICING
Table 12-5 shows the recommended servicing intervals. The hydraulic system is serviced through the
filler cap (3, Figure 12-4) on the hydraulic fluid reservoir and pump assembly. The hydraulic fluid
reservoir incorporates the hydraulic pump and is mounted on the forward left side of the transmission.
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GREASE LUBRICATION —SERVICING
If the helicopter is stored longer than 45 days without operating, purge lubricate the bearings before
returning it to service. Purge lubricate means that fresh grease is injected in the bearing to remove all of
the old grease. If flying in rain or if the helicopter is stored outside during a rain storm, lubricate the
exposed components before flight.
If it becomes necessary to change the brand of grease, remove the old grease by purging until only
new grease is present, except as otherwise specified.
NOTE
Prior to and following grease lubrication, visually examine grease fittings for presence of the spring
loaded steel ball. If the steel ball is not visible, does not spring back to the closed position, the grease
fitting indicates excessive leakage, or it is difficult to inject grease during lubrication, replace the grease
fitting (BHT-ALL-SPM, Chapter 8). If the grease fitting is removed due to the spring loaded steel ball not
being visible, insert a small steel probe into the back of the grease fitting to confirm the presence of the
steel ball. If it is identified that the steel ball is not present in the grease fitting, further investigation is
required to ensure that the steel ball has not migrated into the component and caused damage. Discard
grease fittings that have been removed.
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BATTERY — SERVICING
Access to the battery is through the battery door in the nose compartment. Service the battery in
accordance with Chapter 96, the Electrical Standard Practice Manual (BHT-ELEC-SPM) and the
instructions from the battery manufacturer.
12-38. FIRE EXTINGUISHER
The portable fire extinguisher is mounted on the front of the center post, between the seats of the pilot
and the copilot.
GROUND HANDLING WHEELS — SERVICING
The ground handling wheels are installed on the fittings on the landing gear skid tubes. Servicing the
ground handling wheels includes greasing the axle supports, the wheel bearings and inspecting the tires
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for
wear
and
adequate
inflation.
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EQUIPMENT AND FURNISHINGS
This chapter provides instructions for maintenance of equipment and furnishings provided in the crew
and passenger compartments. Included are such items as crew seats, passenger seats, seat belts,
interior trim, registration certificate case, first aid kit, and loose equipment. A cargo and maintenance
platform is available as optional equipment.
CREW SEATS
The two crew seats in helicopters S/N 4 through 3212 are constructed of aluminum honeycomb panels
and form an integral part of the airframe. Helicopters S/N 3213 and subsequent have stamped sheet
metal seat bottoms in place of aluminum honeycomb panels. The left seat converts from a passenger
seat to a copilot seat with the dual control kit installed. Each seat is equipped with cushions and a laptype safety belt.
PASSENGER SEATS
The aft compartment provides seating for three passengers or, with seats removed, space for cargo is
provided. The seat support is constructed of aluminum honeycomb panels and covers the forward
portion of the fuel cell. The seat deck is composed of three panel assemblies and the center panel is
removable to gain access to the forward part of the fuel cell. The aft seats are provided with cushions
and lap-type safety belts.
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CARGO AND MAINTENANCE PLATFORMS
Dual purpose cargo and maintenance platforms are available as optional equipment. When used as an
interior cargo platform, either or both rear seats are removed and platform sections are installed in the
rear compartment. The platforms consist of bonded aluminum and plywood panels and provide airframe
attachment points and cargo tie-down rings. When not in use as a cargo or work platform, the sections
are stowed in the baggage compartment. The work platform may be installed on either side of the
helicopter.
25-31. BAGGAGE COMPARTMENT
The baggage compartment is located on the left side of helicopter. A hinged access door is provided and
the compartment provides 16 cubic feet (0.45 m3) of space. The baggage compartment is constructed of
aluminum alloy and honeycomb paneling and provides access to the heater and electrical compartment
door. Procedure for lock replacement is the same as for cabin door locks except that a spacer is used in
the baggage door.
25-32. FIRST AID KIT
The first aid kit is provided as loose equipment and can be mounted to the console pedestal or in map
and data case.
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DOORS AND WINDOWS
DOORS
The helicopter is equipped with four entrance/exit doors for crew and passengers, as well as a variety of
hinged doors and panels, which provide access for inspection and servicing. Crew and passenger doors
are located on both sides of the fuselage, and a baggage compartment door is on the left side of the
helicopter (Figure 52-1).
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CREW DOORS.
A crew door is installed on each side of the forward fuselage to provide access to the crew area. Each
door is equipped with a latch assembly, which may be operated from either side of the door, and a lock,
installed in the exterior door handle. Each door incorporates a sliding window and a stationary window.
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PASSENGER DOORS.
A passenger door is installed on each side of the fuselage to provide access to the cabin area. Each
door is equipped with a latch assembly, which may be operated from either side of the door, and a lock
installed in the exterior door handle. Each door incorporates a sliding window and a stationary window.
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BAGGAGE COMPARTMENT DOOR.
A hinged baggage compartment access door is located on the left side of the helicopter. The 16 cubic
feet (0.45 m3) baggage compartment constructed of aluminum alloy and honeycomb paneling also
provides access to heater and electrical compartment access panel. Procedure for lock replacement is
the same as for cabin door locks, except that a spacer is used in the baggage door.
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BATTERY ACCESS DOOR.
The battery access door (2, figure 52-9) is located on the nose of the
helicopter and provides access to the battery, battery relay, and the
hourmeter and circuit breaker. The battery access door is hinged aft. Two
camloc fasteners secure the forward edge of the door to the fuselage
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MISCELLANEOUS ACCESS DOORS.
Doors and panels are provided at various locations in the cowling and fairings for servicing and
inspection of interior areas. Oil reservoir access door (1, figure 52-1), and oil cooler access door (2) are
located on aft fairing. Engine side cowling (3) has side panels which are hinged for easy access, and air
induction cowling doors (5) are located on both sides for inspection of transmission area. All
miscellaneous access doors open on piano hinges and are secured with flush-type latches and/or winghead stud fasteners.
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WINDOWS
Windshield (1, Figure 52-10), lower windows (5), and passenger and crew windows (3 and 4) are bluetinted acrylic plastic. Roof windows (2) are tinted polycarbonate plastic or plex 55. There are two
methods of securing windows to helicopter. One method is to secure retainer or edging to structure with
rivets
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WINDSHIELD
The windshields are fabricated of blue-tinted acrylic plastic, MIL-P5425, supported by formed aluminum alloy sections. Water-tight
sealant is applied to the faying surfaces of the windshield.
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LOWER WINDOWS
Lower windows are located in the lower cabin nose section. Sealant is
applied to mating areas of window panels and nose structure, providing a
water-tight seal.
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ROOF WINDOWS
Two roof windows (skylights) (2, Figure 52-14) constructed of tinted polycarbonate plastic are provided in
the roof of the forward compartment. Windows are supported by formed aluminum alloy sections and
secured with aluminum alloy retainer strips. Sealant is applied to the faying surfaces of the retainer strips
and support to provide a water-tight seal.
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CABIN DOOR -WINDOWS.
The cabin door windows are fabricated of tinted acrylic plastic and secured to the door assembly with
rivets. Sliding windows constructed of acrylic plastic are provided for ventilation. Each window is installed
with water-tight sealing compound of polysulfide rubber applied to the Paying surfaces of the window
and door.
CREW DOOR WINDOWS.
Crew door windows are riveted and sealed to door frame.
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PASSENGER DOOR WINDOWS.
Passenger door windows are tinted plastic. The sliding window is adjustable and moves in a track. The
sliding window handle also functions as a retainer, keeping the window from sliding out of the track.
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LANDING LIGHT WINDOW.
The landing light window (1, figure 52-19) is fabricated of polycarbonate plastic and is located in the
forward end of the console access door. The window protects and provides access to the landing light.
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FUSELAGE
FUSELAGE
The fuselage consists of three main sections: the forward section (1, Figure 53-1) which extends from
cabin nose to bulkhead aft of passenger compartment, intermediate section (2) which extends from the
bulkhead aft of passenger compartment to tailboom attach and tailboom section (3).
FUSELAGE — FORWARD SECTION
The forward section utilizes aluminum honeycomb structure and provides the major load-bearing
elements of the forward cabin. The forward section provides for pilot and passenger seating, fuel cell
enclosure and pylon support.
FUSELAGE — INTERMEDIATE SECTION
The intermediate section utilizes an aluminum semimonocoque construction and provides a deck for
engine installation, a baggage compartment and a compartment under the engine deck for heater and
electrical equipment.
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TAILBOOM
The tailboom is a full monocoque structure except for the forward 10 inch (254.0 mm), where the loads
are redistributed by means of four intercostals load-carrying members. The tailboom supports the tail
rotor driveshaft, tail rotor, gearbox, vertical fin and horizontal stabilizer. Tail rotor driveshaft bearing
supports are mounted to the top of the tailboom. The supports located underneath the bearing support
and inside the tailboom support tail rotor control guide tubes. Covers are provided to protect and provide
a fairing for the tail rotor driveshaft and gearbox.
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VERTICAL FIN
The vertical fin (1, Figure 53-3) is an aerodynamic surface that gives stability to the helicopter while it is
in flight. The vertical fin (1) is installed with the leading edge positioned outboard (with reference to the
aircraft longitudinal axis). This helps to unload the tail rotor while the helicopter is in forward flight. The
vertical fin (1) is made up of an aluminum honeycomb core with aluminum outer skins. The leading and
trailing edge caps (2 and 3) are made of formed aluminum alloy. The anticollision light (4) is installed on
the upper fairing (6) of the vertical fin. The rubber bumper and the tail skid (8) are bonded into the base
of fin. The tail skid (8) absorbs shock in the event of a tail low landing (Chapter 32).
As various combinations of vertical fins (1) and attachment supports (11 and 12) are authorized for
installation, please refer to paragraph 53-12 and paragraph 53-13 to identify the specific configuration
used on your helicopter. Identification of the vertical fin (1) and attachment support combination (11 and
12) will ensure that the proper inspection and installation procedures are used.
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53-12. VERTICAL FIN — IDENTIFICATION
Vertical fins (1, Figure 53-3) have been produced or field modified with various configurations of external
skins and doublers. These vertical fins (1) can also be attached to the helicopter with two types of
attachment supports (11 and 12) (paragraph 53-13). As the various combinations of vertical fins (1) and
attachment supports (11 and 12) require differing inspection and installation procedures. The following
information will ensure that you can identify the vertical fin (1) used on your helicopter. To ease
identification of the 206A/B vertical fins (1), they have been categorized into three types.
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TAIL ROTOR GEARBOX FAIRING
The tail rotor gearbox fairing encloses the tail rotor gearbox and is attached to tailboom and vertical fin.
The fairing incorporates a white position light and two screens/doors. Screens/doors are used for tail
rotor gearbox oil level inspection and cooling of tail rotor gearbox.
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HORIZONTAL STABILIZER
The horizontal stabilizer is constructed of aluminum and is attached to a spar with clamps. The inboard
rib of the horizontal stabilizer contains a fitting which secures stabilizer to tailboom with bolts.
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LANDING GEAR
LOW AND HIGH SKID LANDING GEAR SYSTEMS.
The low (standard) and high skid landing gear consists of two tubular aluminum alloy main skid tubes
and two curved tubular aluminum crosstubes. Airflow type fairings are provided on the forward and aft
crosstubes. The landing gear is attached to the fuselage structure with four strap assemblies.
Provisions are made on skid tubes for installing ground handling wheels and tow rings for towing. Each
skid tube is equipped with replaceable skid shoes. The skid shoes absorb the wear caused by normal
ground contact of the landing gear.
WARNING
NO COMPONENTS SHALL BE ATTACHED TO LANDING GEAR ASSEMBLY EXCEPT AS
DESIGNATED BY MANUFACTURER. TO DO OTHERWISE MAY RESULT IN PREMATURE FAILURE
OF THE CROSSTUBE.
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LOW SKID FAIRINGS
The airfoil type, skid landing gear fairings enclose the forward and aft crosstubes and are constructed
from white thermoplastic.
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HIGH SKID LANDING GEAR.
WARNING
NO COMPONENTS SHALL BE ATTACHED TO LANDING GEAR ASSEMBLY EXCEPT AS
DESIGNATED BY MANUFACTURER. TO DO OTHERWISE MAY LEAD TO PREMATURE FAILURE
OF CROSSTUBE.
The high skid landing gear provides approximately 13 inches (330 mm) additional ground clearance. The
high skid landing gear assembly consists of two tubular aluminum alloy main skid tubes and two cured
tubularaluminum alloy crosstubes (figure 32-12). The landing gear is attached to fuselage structure with
four straps. Provisions are made on skid tubes for installing ground handling wheels and tow rings are
provided for towing. Each skid tube is provided with replaceable skid shoes constructed from normalized
4130 steel alloy conforming to MIL-S-18729. Four fuselage mounted cabin steps are provided to
facilitate entrance and exit.
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FIXED STEP ASSEMBLY.
NOTE
Four externally mounted fixed steps are mounted on the fuselage structure to provide safe entrance
and/or exit of helicopter when the high skid landing gear is installed.
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GROUND HANDLING WHEELS.
Hand operated ground handling wheel assemblies (14, figure 32-21) are mounted on each skid tube (13)
near helicopter center of gravity to facilitate helicopter handling or movement. Wheel assemblies (4) are
retracted and extended manually and are removable. Two 6 ply, 3.50 x 6, nylon tires and tubes are used
on the 2 wheel assemblies (4).
WARNING
MAINTAIN WIDE STANCE BALANCE, HOLDING LIFT TUBE (1) FIRMLY WHILE RAISING OR
LOWERING GROUND HANDLING WHEELS.
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TAIL SKID.
A tubular steel tail skid and bumper are installed on the lower portion of the vertical fin and act as a
protective device for the tail rotor and tailboom in the event of a tail-low attitude in landing.
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PNEUMATIC
ENVIRONMENTAL
SYSTEMS
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AIR DISTRIBUTION (VENTILATION)
VENTILATION SYSTEM
Air for cabin ventilation is obtained by opening sliding windows in each of the entrance doors (Figure 211). Helicopters S/N 4 through 253 provide additional air for cabin ventilation by a ram air scoop mounted
under the forward transmission fairing. Four adjustable valves, located in forward and aft cabin roof,
provide air distribution for passenger and crew areas. Helicopters S/N 254 and subsequent provide ram
air ventilation for the crew area only, through ram air grilles located in nose of helicopter (Figure 21-1).
This additional air is obtained by pulling the VENT control knob under instrument panel. Positioning
DEFOG BLOWER switch or circuit breaker on overhead panelto ON will circulate air onto windshield to
defog.
RAM AIR SYSTEM (HELICOPTERS S/N 4 THROUGH 253)
In flight, ram air is forced into the ram air scoop (Figure 21-1) and is distributed through ducts, located
in the cabin roof, to the cabin through four adjustable valves in crew and passenger compartments.
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RAM AIR SYSTEM (HELICOPTERS S/N 254 THROUGH 2488)
The ram air system is part of the vent and defog system (Figure 21-1). With the VENT control knob (2,
Figure 21-3 for S/N 254 through 448 and Figure 21-4 for S/N 449 through 2488) pulled out, ambient ram
air will be forced through ram air grilles, located on helicopter nose, into air plenum assembly then into
crew compartment through defog system.
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RAM AIR SYSTEM (HELICOPTERS S/N 2489 AND SUBSEQUENT)
With the VENT control knob (10, Figure 21-5) pulled out, ambient ram air will be forced through ram air
grille on the nose of helicopter, and directed through the plenum and flapper valve into the crew
compartment.
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VENT CONTROL CABLE
Vent control cables (7, Figure 21-3 for S/N 254 through 448 and Figure 21-4 for S/N 449 through 2488,
or 12, Figure 21-5 for S/N 2489 and subsequent) are mounted on brackets on lower instrument panel.
The VENT control knob will lock in all positions when pulled and will unlock by pressing release button in
control knob.
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INLET PLENUM AND OUTLET PLENUM ASSEMBLY (HELICOPTERS S/N 449 THROUGH 2488)
A one-piece combination inlet plenum and valve assembly (1, Figure 21-4), and outlet plenum (19),
constructed of molded polycarbonate, provides direction and control of ram air ventilation. The inlet
plenum (1) contains a flapper valve that is connected to a control cable (7) and VENT control knob (2)
and is used by the pilot to control ventilation air flow. Drain tubes (23 and 24) are connected to the inlet
plenum (1) and outlet plenum assembly (19) to drain moisture from ram air ventilation system.
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INLET PLENUM AND TURN VANE ASSEMBLY (HELICOPTERS S/N 2489AND SUBSEQUENT)
An inlet plenum (26, Figure 21-5) and turn vane (18), constructed of molded polycarbonate, direct and
control ram air ventilation. The inlet plenum contains a flapper valve (45, Detail A) that is connected to a
control cable (12) and VENT control knob (10), to provide control of ram air flow by pilot. A drain tube
(25) connected to the inlet plenum drains moisture from ram air ventilation system.
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DEFOG SYSTEM
One defog nozzle is installed on each side of the console for distribution of windshield defogging air. Air
is supplied by an electrically driven blower for ventilation and defogging, primarily during ground
operation of the helicopter. On helicopters S/N 4 through 253 and S/N 2489 and subsequent (Figure 217), the blower is mounted to the forward end of the defog nozzle. On helicopters S/N 254 through 2488
(Figure 21-3 and Figure 21-4), the blower is mounted between the ram air outlet plenum and the defog
nozzle. In all configurations, the bloweris controlled by a DEFOG BLOWER circuit breaker-type switch
on the overhead console. On helicopters S/N 254 through 2488, it is recommended that both VENT
control knobs be pulled out and lockedin the full open position. On helicopters S/N 4 through 253, and
2489 and subsequent, the defog blowers may be operated with the VENT control knobs in the CLOSED
position.
BLOWER
The defog blower is an electrically operated, axial flow, multivane-type unit.
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DEFOG NOZZLE
Defog nozzles are constructed of molded polycarbonate. Outlet slots are located along upper edge of
nozzle to direct air onto windshields. Three spacers are located just below outlet slots for mounting defog
nozzles to structure support angles.
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ENVIRONMENTAL CONTROL
SYSTEM (ECS).
The environmental control is capable of heating, cooling, and dehumidifying air to the cabin. The kit
consists of an environmental control unit (ECU), control panel, circuit breaker, switches, solenoid valve,
relay, and associated ducting and hardware. The ECU is located in the equipment compartment directly
above the baggage compartment. Electrical controls for the ECU are located in the overhead console.
The ECU receives bleed air directly from the engine.
OPERATION – ENVIRONMENTAL CONTROL SYSTEM.
1. After closing the ECS circuit breaker and positioning the COOL/HEAT switch to MAX HEAT,
the ECS is activated. An audible "thump" sound indicates solenoid opened heater.
2. The temperature control plus ECS control switch located in the overhead console, determines
the temperature delivered to the cabin from the heat exchanger.
3. The duct overheat switch and/or the airframe thermostat switch keep the ECS operating within
the temperature limits determined by the position of the COOL/WARM knob. When temperature
limit is exceeded, one or both switches close. This closes the bleed air valve and de-energizes
the blower relay.
4. The ECS overload sensor protects the blower motor from being overloaded. A motor overload
condition causes the sensor to close. This de-energizes the blower relay.
BLEED AIR HEATER.
The bleed air heater delivers heated forced air from the engine and outside to the forward and
passenger compartment and to the windshield. The kit consists of electrically actuated shutoff and airmixing valves, circuit breaker, switches, relay, overheat sensing switch caution light, and associated
ducting and hardware. The heater and overheat sensing switch are located in the equipment
compartment directly above the baggage compartment. Electrical controls for the heater are located in
the overhead console.
OPERATION - BLEED AIR HEATER.
NOTE
Engine must be operating for heater operation.
1. After selecting HEAT on HEAT/VENT switch (overhead console), closing HTR CONT circuit
breakers activates the bleed air heater. An audible "thump" sound indicates solenoid opened
damper.
2. The temperature control, located below and to the right of the overhead console, determines
the setting of the overheat sensing switch. The overheat sensing switch is located in the heater
duct in the equipment compartment.
3. The overheat sensing switch (S69) opens the HTR CONT circuit breaker within six seconds
when temperature selected by the temperature control is exceeded. This de-energizes the heater
valve solenoid, which shuts off bleed air.
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ELECTRICAL
SYSTEMS
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ELECTRICAL SYSTEMS.
Models 206A, 206B and 206B JetRanger III helicopters are equipped with a 28 volt direct current (Vdc)
electrical system. Power for this system is obtained from a nickel-cadmium, vented, 24 volt, 13 amperehour or 17 ampere-hour (S/N 4299 and subsequent) battery and a 30 volt, 150 ampere (derated to 105
amperes) combination starter-generator. Major components of do power system include battery, startergenerator, voltage regulator, relays, and circuit breakers. All circuits in electrical system are single wire
with a common ground return. Negative terminals of starter-generator and battery are grounded to
helicopter structure.
Controls for electrical systems are located on overhead console and instrument panel. For location of
control relays, power relays, voltage regulators, and other electrical components, refer to figure 96-1.
Refer to Chapter 98 for electrical systems wiring diagrams. External power may be supplied to the
helicopter by means of a receptacle located at the lower front section of the fuselage.
OPERATIONAL CHECKS.
When performing operational checks, external power should be utilized whenever possible. Perform
operational checks to ensure circuits are free of malfunctions after equipment has been replaced or
airframe wiring repaired or replaced.
NOTE
For checks, adjustments, or repairs not covered in this manual, consult the handbook published by the
applicable
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ELECTRICAL COMPONENTS.
Included in this category are relays, solenoids, variable resistors, switches, circuit breakers, plugs, leads,
connectors, wiring, receptacles, shunts, capacitors, diodes, transistors, resistors, inductors, transducers,
synchros, and panel lights.
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CONTROL PANELS - ELECTRICAL.
Controls are mounted in the overhead console instrument panel and instrument panel pedestal. See
figure 96-2 forconsole/panel illustration and Chapter 95 for instrument panel.
CIRCUIT BREAKERS.
Circuit breakers are mounted in overhead console. Circuits can be opened and closed with these pushpull circuit breakers (figure 96-2).
.
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POWER SYSTEMS.
DC power systems include battery, external power, generator, and starter-igniter systems.
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BATTERY SYSTEM.
Battery system includes battery, battery relay, and battery switch. Helicopters S/N 716 and subsequent
are equipped with BATT TEMP and BATT HOT sensors and related wiring.
BATTERY.
Battery (BT1) is located in nose section of helicopter. For helicopters S/N 4 through 4298, battery is a
vented, 24-volt, 13-ampere-hour, nickel-cadmium battery. For helicopters S/N 4299 and subsequent,
battery is a vented, 24-volt, 17 ampere-hour, nickel-cadmium battery
WARNING
DO NOT USE ACID. INJURY MAY RESULT AND EQUIPMENT DAMAGE CAN OCCUR. NEVER
ALLOW ANYTHING ASSOCIATED WITH (INCLUDING ACID FUMES) OR CONTAMINATED BY LEADACID TO COME IN CONTACT WITH BATTERY.
NOTE
Nickel-cadmium batteries are different from lead-acid batteries. Terminal voltage remains constant over
90 percent of total discharge time; a terminal voltage test is not conclusive. A hydrometer test is not
effective because electrolyte specific gravity remains constant if battery is either in a charged or
discharged condition.
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CHARGING BATTERY IN HELICOPTER (EXTERNAL POWER).
1. Ensure electrical power is OFF.
2. Plug external power into helicopter.
3. Check and set voltage on external power unit at 27.5 to 28.0 volts
4. Turn external power unit ON and note reading of external power unit ammeter.
5. Set battery switch to BAT and note rise in ammeter reading on external power unit (approximately 100
amps).
6. Continue charging until ammeter has dropped to about the same reading recorded before the
helicopter battery switch was actuated. This time will be approximately 10 minutes.
7. When current has dropped to previously recorded reading, the battery is charged and ready for
service. Set battery switch to OFF; turn external power OFF, and disconnect power unit.
CHARGING BATTERY OUT OF HELICOPTER (EXTERNAL POWER).
1. Use same procedure outlined in paragraph 96-27 except battery is not installed in helicopter.
2. With this method the proper adapter must be used to interconnect battery and external power unit.
CAUTION
DO NOT USE CABLE CLIPS. DAMAGE TO BATTERY TERMINALS WILL RESULT.
3. Ensure connections and cables are capable of carrying at least 100 amps.
4. Charge battery as outlined in paragraph 96-27, recording first the current required by external power
unit, then closing generator power switch and continuing charge until current has dropped to value first
recorded on external power unit.
5. When current has dropped, battery is charged and ready for installation in helicopter.
CHARGING BATTERY - SLOW CHARGE.
Set battery charger on 24-volts slow rate. If battery is known to require a complete recharge, due to
accidental discharge or because battery has been stored for exceptionally long periods of time
(particularly at high temperatures), charging time should be approximately 10 hours (BHT-ELEC-SPM).
NOTE
The Marathon battery is sufficiently vented when installed in the helicopter to expel accumulated gases
during charging and discharging.
NOTE
A new battery is discharged and may require approximately 10 hours charging (BHT-ELEC-SPM).
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DUAL BATTERY.
The auxiliary battery offers increased electrical power for cold weather starts, short trips, and frequent
starts. The kit consists of a 13 ampere-hour battery, relay, heat sensors, and a three-position switch. The
BAT SEL switch (three-position battery select switch) is located in the bottom edge-lit panel which is
mounted in the instrument panel pedestal. The switch allows selective usage of the main battery (FWD
BAT) and/or the auxiliary battery (AFT BAT). The auxiliary battery is located in the left aft compartment.
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EXTERNAL POWER SYSTEM.
CAUTION.
MAXIMUM CURRENT OF EXTERNAL. POWER SOURCE SHOULD NOT EXCEED 1000 AMPS..
External power system includes external power receptacle, external power relay, and related wiring.
EXTERNAL POWER RECEPTACLE.
External power receptacle (J16) is located on front center of the nose section. It is a polarized receptacle
and used to connect external power to the helicopter.
EXTERNAL POWER RELAY.
External power relay (K1) located in nose section forward of pedestal, electrically controls external power
to main bus bar. Small positive pin of external power receptacle energizes circuit to relay coil, causing
contacts to close.
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GENERATOR SYSTEM.
For helicopters S/N 4 through 4310 the generator system consists of the generator portion of startergenerator, voltage regulator, reverse current relay, generator reset switch, generator shunt, generator
reset relay, field control relay, and over-voltage sensing relay.
The generator furnishes regulated power for all do electrical circuits of helicopter. Generator output is
transferred to main bus when generated voltage exceeds bus voltage by 0.30 to 0.42 volts.
Reversecurrent relay connects generator power to bus when voltage regulator senses adequate
generator voltage. Over-voltage sensing relay and reverse current relay provide protection against overvoltage and reverse-current conditions. Voltage regulator compensates for voltage fluctuations caused
by varying load conditions.
For helicopters S/N 4311 and subsequent the generator system consists of the generator portion of the
starter-generator, solid state voltage regulator, line control relay, field ignition relay, generator reset
switch, and generator shunt.
The generator furnishes regulated power for all do electrical circuits of the helicopter. Generator output is
transferred to main bus when a minimum of 24 vdc is achieved.
The solid state voltage regulator monitors for: a minimum of 24 vdc output from the generator; overvoltage of 31 ± Vdc; low voltage of 18 ± 1.8 Vdc, and reverse-current protection range between 16 and
25 amps. The voltage regulator compensates for voltage fluctuations caused by varying load conditions.
GENERATOR.
Starter-generator (G1) is located on underside of the engine to right of helicopter centerline. This unit is
used to start engine, charge battery, and supply power for operation of do equipment.
VOLTAGE REGULATOR.
A carbon pile type voltage regulator (VR1) is located below the instrument panel on helicopters S/N 4
through153 and on the equipment shelf above the baggage compartment on helicopters S/N 154 and
subsequent. The VR1 functions as a variable resistor in generator shunt field circuit to maintain a
generator output voltage constant at adjusted value, regardless of the varying load.
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REVERSE CURRENT RELAY
(Helicopters S/N 4 through 4310). Reverse current relay (K5) is located on equipment shelf above
baggage compartment. Reverse current relay prevents generator from being connected to line until
operating voltage is attained, prevents reverse current flow, and keeps generators on line unless voltage
drops to a point where continued operation would be detrimental to electrical equipment.
GENERATOR RESET SWITCH
(Helicopters SIN 584 and subsequent). Generator reset switch (S90) is located in overhead console
(figure 96-2). This switch is a double pole, double throw, spring-loaded switch with only momentary
contact in RESET position. It completes generator field circuit in ON position and supplies voltage to
reset generator field reset relay in RESET position.
GENERATOR SHUNT
Generator shunt (R3) is located just inboard of starter relay on the equipment shelf above baggage
compartment. It provides a voltage drop proportional to the generator load current for indication on
loadmeter.
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GENERATOR FIELD CONTROL RELAY (Helicopters S/N 4 through 4310).
Generator field control relay (K4) is located aft of generator shunt on equipment shelf above baggage
compartment. This unit is an electrically operated switch, which opens the starter-generator shunt field
when generator is used as a starter and it also completes the igniter circuit. Resistor installed between
terminals Al and X2 provides approximately one volt (positive) to terminal A of starter-generator during
engine starts.
GENERATOR FIELD RESET RELAY
(Helicopters SIN 4 through 4310). Generator field reset relay (K41) is located on the equipment shelf
above the baggage compartment. This unit is a double action type relay, which opens generator shunt
field and disconnects generator from line when over-voltage condition exists. It can be electrically reset
by the generator reset switch (S90).
OVERVOLTAGE SENSING RELAY
(Helicopters SIN 4 through 4310). The overvoltage sensing relay (K42) is located on equipment shelf
above the baggage compartment. This relay is energized when line voltage reaches 31 (±1) volts. In turn
it energizes the generator field reset relay to trip position, removing generator from line.
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STARTER-IGNITER SYSTEM.
The starter-igniter system includes starter portion of starter-generator, starter relay, generator field
control relay, igniter, and starter switch.
STARTER.
The starter-generator (G1) located on underside of engine, is energized by starter relay to start engine.
Refer to Chapter 71 for removal, inspection, repair or replacement, and installation procedures.
STARTER RELAY.
The starter relay (K3) is located on electrical equipment shelf above baggage compartment and supplies
direct current to starter when starter switch is depressed.
IGNITER.
Igniter (Z1) is furnished with the turbine engine and is located below power turbine tachometer generator
on lower left section of engine. This unit consists of a low tension capacitor discharge ignition exciter,
which provides continuous ignition arc during engine start cycle.
STARTER SWITCH.
The starter switch (S6), located in collective stick switchbox, is a double-pole, single-throw, pushbutton
type switch. When switch is depressed to START, starter relay and generator field control relay energize.
This applies power to starter and igniter and completes the generator field shunt weakening circuit.
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INDICATING SYSTEMS
INDICATING SYSTEMS
The following procedures cover only instrument system components, which are engine, transmission, or
airframe mounted. Refer to Chapter 95 for individual indicator and indicator circuitry maintenance,
operational checks, etc. Refer to Chapter 98 for systems wiring diagrams.
OIL TEMPERATURE BULBS
Electrical temperature sensitive resistance bulbs are used in engine oil and transmission oil temperature
indicating systems. Each bulb is part of a resistive bridge circuit connected in series with indicator. As
engine or transmission oil temperature changes, resistance of bulb will change, causing indicator
movement. The resistance elements of bulbs are hermetically sealed in metal wells. The transmission oil
temperature bulb (Z3) is located on left side of transmission and the engine oil temperature bulb (Z2) is
mounted in a line near engine oil reservoir.
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TACHOMETER GENERATORS
Tachometer generators (G2 and G4) are three-phase alternating current generators that generate
signals to drive dual and gas producer tachometer indicators. Rotor tachometer generator is located on
forward left side of transmission. The power turbine tachometer generator is mounted on forward left
side and gas producer tachometer generator is mounted on forward right side of power and accessory
gearbox.
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BLEED AIR PRESSURE SWITCH (HELICOPTERS S/N 4 THROUGH 583)
The bleed air pressure switch, located beneath the service deck below the forward left side engine, is a
pressure operated switch (Figure 96-1 and Chapter 98). It closes when gas producer speed falls below
60% completing circuits to engine out warning horn (Figure 96-2) and ENG OUT warning light (Figure
96-12) simultaneously to alert pilot of engine failure.
TRANSMISSION OIL PRESSURE TRANSDUCER (HELICOPTERS S/N 254 THROUGH 913)
The transmission oil pressure transducer is connected to the cross fitting at the left side of the
transmission. It is a pressure operated potentiometer that varies input voltage to the transmission oil
pressure indicator (Chapter 98).
96-105. ENGINE OIL PRESSURE TRANSDUCER (HELICOPTERS S/N 254 THROUGH 913)
The engine oil pressure transducer is connected to the tee fitting at left side of the engine accessory
section. It is a pressure operated potentiometer that varies the input voltage to the engine oil pressure
indicator (Chapter 98).
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TORQUE PRESSURE TRANSDUCER (HELICOPTERS S/N 254 THROUGH 913).
The torque pressure transducer is connected to the engine accessory section. It is a pressure operated
potentiometer that varies the input voltage to the torque pressure indicator (Chapter 98 and TB BHT206-08-73).
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FUEL QUANTITY SYSTEM
FUEL QUANTITY INDICATOR (HELICOPTERS S/N 4 THROUGH 2211)
The fuel quantity indicator, located on instrument panel, on helicopters S/N 914 through 2211, and in
instrument cluster on helicopters prior to S/N 914, is calibrated in gallons. Indicator is part of bridge
circuit, which includes two tank units, two float elements, two calibration variable resistors, necessary
terminal blocks, indicator, and 28 VDC that serves a common bus inside instrument cluster unit.
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96-109. FUEL QUANTITY INDICATOR (HELICOPTERS S/N 2212 AND SUBSEQUENT)
The fuel quantity indicator, located on instrument panel, is calibrated in gallons. The indicator is part of
the bridge circuit, which includes two fuel level transmitter (resistive float elements), two variable
calibration resistors, and terminal blocks. Both fuel level transmitters are mounted in tank. One monitors
fuel level up to the horizontal surface of tank under the seat; the other monitors fuel level in upper
section of tank behind the seat.
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FUEL PRESSURE SYSTEM.
The fuel pressure system is composed of two electrically operated fuel boost pumps, submerged in the
fuel cell, accessible from the bottom of the fuselage. Both pumps are connected to a common fuel line
and either will furnish sufficient flow for engine operation. The pumps are energized from separate circuit
breakers in the overhead console (figure 96-2) and may be operated separately or together.
FUEL PRESSURE TRANSDUCER
(Helicopters S/N 254 and subsequent).
The fuel pressure transducer is connected to the tee fitting at aft end of fuel pressure switch in the
engine compartment. On helicopters S/N 1904 and subsequent, the fuel pressure transducer is located
near the fuel shutoff valve. It is a pressure operated potentiometer and varies the input voltage to the fuel
pressure indicator.
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LIGHTING SYSTEMS.
Lighting system comprises equipment for illumination of instruments, and switches, operation of
interior/exterior, and landing lights. Refer to Chapter 98 for individual system wiring diagrams.
INTERIOR LIGHTING SYSTEM.
The interior lighting system includes the following components and related wiring:
1. Two edge or integrally lit control panels, located on instrument panel.
2. One integrally lit control panel, located in forward section of overhead console.
3. A cockpit light is located on the control post between the crew seats or below the pilot seat at
centerline of helicopter.
EXTERIOR LIGHTING SYSTEM.
The exterior lighting system includes the following components and related wiring:
1. Two landing lights are in the lower section of the nose.
2. Helicopters S/N 4 through 103 utilize one relay for the landing lights; helicopters S/N 104 and
subsequent utilize two. These relays are located in the nose compartment below the instrument
panel. An option of both landing lights ON, or forward landing lights ON only, is available.
3. The landing light switch is located on the pilot collective stick switchbox.
4. The position lights are located on the tips of the horizontal stabilizers.
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CAUTION AND WARNING SYSTEMS
OPERATIONAL CHECK - CAUTION AND WARNING LIGHTS.
Caution and warning systems consist of caution lights (segments) located across top of instrument
panel, and their respective sensing devices. A caution light test switch allows testing of caution lights. A
bright/dim switch (located on an edge-lit panel of instrument pedestal) in conjunction with INST LT
control varies intensity of caution lights. An ENG OUT warning horn is located on overhead console.
Refer to figure 96-12 for caution panel.
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ENGINE OUT WARNING SYSTEM.
Engine out warning system provides a visual and audible indication of an engine out condition. System
includes ENG OUT warning light on caution panel, engine out warning horn in overhead console, and
engine rpm sensor on equipment shelf forward of instrument panel. On helicopters S/N 4 through 583
the ENG OUT warning light and horn are operated by the bleed air pressure switch. Refer to indicating
system for description. Engine rpm sensor is connected to gas producer tachometer generator. When
gas producer rpm drops below 55 (±3) percent, engine rpm sensor completes an electrical ground to
ENG OUT warning light and engine out alarm (horn) circuit.
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ROTOR LOW RPM CAUTION SYSTEM
(Helicopters S/N 4 thru 583 with S1206-74 installed and Helicopters S/N 584 and subsequent).
Rotor low rpm caution system provides visual and audible indication of rotor low rpm condition. System
includes ROTOR LOW RPM caution light on caution panel, rotor low rpm alarm in right side plastic
headliner, and rotor low rpm sensor on equipment shelf forward of instrument panel. Rotor low rpm
sensor is connected to rotor tachometer generator. When rotor rpm drops below 90 (±3) percent, rotor
low rpm sensor completes electrical ground to ROTOR LOW RPM caution light and rotor low rpm
warning circuit. Rotor low rpm warning alarm disable switch is installed under copilot seat, slightly
forward of collective jackshaft. When collective stick is in extreme down position, a lever on the jackshaft
opens switch and deactivates rotor low rpm alarm.
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TRANSMISSION OIL PRESSURE/TEMPERATURE CAUTION SYSTEM.
Transmission oil pressure/temperature caution system is comprised of piping (wet line), pressure switch,
temperature sensing switch, caution lights (segments), and associated wiring.
TRANSMISSION OIL PRESSURE CAUTION SYSTEM.
Transmission oil pressure switch (S4) is connected with a T-fitting into oil pressure piping (wet line).
Switch is located centered, forward, and below instrument panel. Contacts of switch are kept open by
transmission oil pressure, unless pressure drops to 28 psig (193 kPa) or below. When switch is allowed
to close at 28 psig (193 kPa), it completes transmission oil pressure caution circuit; transmission oil
pressure segment on caution panel illuminates.
TRANSMISSION OIL TEMPERATURE CAUTION SYSTEM.
Transmission oil temperature switch (S3) is a hermetically sealed, temperature sensitive component.
Switch will close when transmission oil temperature rises above safe operating limit. This completes
transmission oil temperature caution circuit; TRANS OIL TEMP segment on caution panel illuminates.
Switch is located adjacent to transmission oil temperature bulb on left side of transmission.
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BATTERY TEMPERATURE SENSING SYSTEM (Helicopters SIN 716 and subsequent).
CAUTION
THE FOLLOWING PARAMETERS AND PROCEDURES ARE FOR ONE SPECIFIC TYPE OF
BATTERY. SENSING SYSTEM CHARACTERISTICS AND PERFORMANCE DATA MAY VARY WITH
TYPE OF BATTERY USED. REFER TO VENDOR MANUALS TO ENSURE COMPATIBILITY OF
EQUIPMENT.
Battery temperature sensing system consists of battery overtemp sensor module (S103), BATTERY
TEMP caution light (DS42), BATTERY HOT warning light (DS43), and related wiring. Switch (S1) in
overtemp sensor module (S103) closes when battery case temperature reaches 130°F (54.4°C) which
illuminates BATTERY TEMP caution light (DS42). If temperature reaches 140°F (60°C), switch (S2) in
overtemp sensor module (S103) closes which illuminates BATTERY HOT warning light (DS43). When
BATTERY TEMP caution light illuminates, the battery charging circuit must be disengaged to allow
battery case temperature to drop below 130 ° F (54.4 ° C).
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ENGINE CHIP DETECTOR CAUTION SYSTEM (HELICOPTERS S/N 914 AND SUBSEQUENT)
Engine chip detector caution system is comprised of ENG CHIP caution light, two engine mounted (part
of engine) magnetic drain plug/chip detectors, and related wiring. If metal particles should segregate
from engine into oil, the magnet will attract these particles. If sufficient metal has been attracted to
complete circuit between pole (core of chip detector) and ground, ENG CHIP caution light will illuminate.
TRANSMISSION CHIP DETECTOR CAUTION SYSTEM
Transmission chip detector caution system includes TRANS CHIP caution light, two transmission chip
detectors (mounted 90° apart), tail rotor gearbox chip detector, and a mast bearing chip detector on
helicopters S/N 3905 and subsequent, and related wiring.
FUEL LOW CAUTION SYSTEM
The fuel low system, which is independent of the fuel quantity system, illuminates a FUEL LOW caution
light when there is approximately 20 U.S. gallons (206A/B and B3 helicopters S/N 4 through 4052 Post
TB 206-84-94 or Post TB 206-85-113) or 12 to 19 U.S. gallons (206B3 helicopters S/N 4053 and
subsequent) of usable fuel remaining.
FUEL FILTER DIFFERENTIAL PRESSURE SWITCH
Filter pressure switch (S10) is part of fuel filter assembly, which is mounted on lower engine firewall.
Switch is pressure operated and connected to fuel filter caution light. If fuel filter pressure drops below
safe operating limit due to clogged filter, switch closes, and fuel filter caution light illuminates. Fuel is
bypassing filter at this time.
AIRFRAME FUEL FILTER CAUTION SYSTEM (AFTER INCORPORATION OF TB 206-82-75)
Airframe fuel filter (impending bypass) switch is part of engine inlet fuel filter assembly, which is mounted
on left side of engine compartment. Switch is pressure operated and is connected to A/F FUEL FILTER
caution light. If filter element becomes too clogged to allow normal fuel flow, bypass valve opens and
allows fuel to circumvent filter. Impending bypass switch closes prior to bypass occurring; caution light
illuminates. Circuit incorporates a press-to-test button on filter assembly; depressing button will illuminate
caution light and verify circuit integrity.
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ENGINE CONTROL AND ACCESSORY SYSTEMS
ENGINE CONTROL AND ACCESSORY 96-207. FUEL DRAIN SYSTEM (Helicopters 716 and
subsequent).
Engine control and accessory systems include fuel boost pumps, fuel shutoff valve, fuel dump system,
governor control switch, governor actuator, and engine anti-icing control unit.
FUEL BOOST PUMPS.
Two electric fuel boost pumps (B1 and B2) are mounted submerged in fuel cell. They are accessible
from bottom of fuselage. Both pumps are connected to a common fuel line. Either pump can furnish
sufficient fuel flow for engine operation. Pumps are controlled by separate circuit breakers in overhead
console (figure 96-2) and may be operated separately or together. (See figure 96-1 for equipment
location and Chapter 98 for wiring diagram.)
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FUEL SHUTOFF VALVE.
Fuel shutoff valve, located above fuel cell on right side of helicopter, is an electric solenoid valve and
provides the means to shut off fuel to engine. Valve is controlled by a switch on instrument panel.
FUEL DRAIN SYSTEM
Fuel drain system consists of solenoid actuated fuel drain valve (1315) and fuel drain switch (S91). Valve
is located at lowest point of lower fuel cell, and vents directly overboard when energized. Drain switch is
spring-loaded to off position, and is mounted flush with fuselage. Switch has a rubber cover, and is
located on right side of helicopter, directly above rear skid mount. As long as switch is depressed, 28
Vdc is applied to valve solenoid; this opens drain valve and allows venting. Drain system provides for
quick drainage of moisture accumulation from tank system prior to flight. System receives power through
fuel valve circuit breaker (CB2) when the fuel shutoff valve is in the OFF position.
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GOVERNOR RPM SWITCH.
Governor rpm switch (S5), located in pilot collective switchbox, is double-pole, double-throw, springloaded to center, momentary contact switch. It enables pilot to increase or decrease governor rpm
actuator setting. With switch in INCR position, circuit to actuator motor is completed and allows motor to
move arm in one given direction. With switch in DECR position polarity to actuator motor is reversed,
allowing actuator arm to move in opposite direction. When switch is at center position, circuit is deenergized
GOVERNOR ACTUATOR.
Governor actuator (133) is located on forward left side of engine. It is a reversible motor and provides
increase or decrease of governor setting. Unit is controlled by governor switch on collective stick
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ENGINE ANTI-ICING ACTUATOR.
Engine anti-icing actuator (134) is located on upper forward section of engine; its position determines
condition (open or closed) of engine anti-icing valve. Actuator is controlled by anti-icing switch on
instrument panel pedestal
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HYDRAULIC CONTROLS
Hydraulic servo actuators are used as power boost source for the flight control system. Hydraulic control
system reduces operating load on flight controls.
HYDRAULIC BYPASS SOLENOID AND SWITCH.
Hydraulic bypass SOLENOID (1-1) is located on service deck forward of transmission; it is controlled by
control boost switch (S7) in instrument panel pedestal. With switch OFF, SOLENOID is energized,
allowing boost system to be bypassed. Setting switch to ON activates control boost system.
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FIRE PROTECTION
HAND FIRE EXTINGUISHER
A manually operated fire extinguisher is furnished with each helicopter. The extinguisher is located on
the forward side near the top of the control tube vertical tunnel or on the aft end of the console (Figure
26-1). Mounting brackets are the quick-opening type for rapid removal of extinguisher and are secured
with screws, spacers, washers, nuts, and nut plates. The extinguisher consists of two major parts. The
operating head embodies a discharge nozzle, operating lever, safety catch or pin and red discharge
indicator disc. Body of extinguisher is the agent container and incorporates a pressure gauge.
Extinguisher is charged with Halon 1211 and is rechargeable at the factory. Replacement cylinders are
available. Engine fire protection consists of engine firewalls for containment.
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INSTRUMENT
AND AVIONICS
SYSTEM
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INSTRUMENT SYSTEM
The instrument system is divided into four separate categories: flight, navigation, propulsion, and
miscellaneous. All indicators are installed in the hinged instrument panel (figures 95-1 through 95-4)
except the pilot standby magnetic compass, hourmeter, and free air temperature indicator. The pilot
standby magnetic compass is mounted in a support attached to the rightside of the cabin structure,
slightly forward of the instrument panel. The hourmeter is mounted in the nose compartment, and free air
temperature indicator is mounted in upper left corner of the pilot windshield.
INSTRUMENTS.
1. The flight instrument system includes the pitot-static system and the following instruments:
a. Airspeed indicator
b. Altimeter
c. Inclinometer.
2. The navigation instrument system consists of the pilot standby magnetic compass.
3. The propulsion instrument system includes the following instruments:
a. Dual tachometer indicator
b. Gas producer tachometer indicator
c. Engine oil temperature/pressure indicator
d. Transmission oil temperature/pressure indicator
e. Engine torquemeter
f. Turbine outlet temperature indicator.
4. The miscellaneous instrument system includes the following instruments:
a. Hourmeter
b. Eight day clock
c. Free air temperature indicator
d. DC loadmeter
e. Fuel quantity indicator
f. Fuel pressure indicator.
5. There have been significant changes to the layout of the control panel throughout the evolution of the
206A/13 series as seen in figure 95-1 through 95-4. The changes in the individual instruments are as
follows:
a. Helicopters S/N 4 through 153 are not equipped with an inclinometer (figure 95-1).
b. Helicopters S/N 154 and subsequent are equipped with an inclinometer (11, figure 95-2) (4, figure
95-3), (5, figure 95-4).
c. Helicopters S/N 4 through 913 are equipped with separate fuel pressure indicators and DC loadmeters
(7 and 8, figure 95-1) (15 and 1, figure 95-2).
d. Helicopters S/N 914 and subsequent are equipped with a dual fuel pressure indicator/DC loadmeter.
(11, figure 95-3), (12, figure 95-4).
e. Helicopters S/N 4 through 913 are equipped with separate engine oil temperature and engine oil
pressure indicators (11 and 12, figure 91-1) (4 and 13, figure 95-2).
f. Helicopters S/N 914 and subsequent are equipped with dual engine oil temperature/pressure
indicators (1, figure 95-3), (1, figure 95-4).
g. Helicopters S/N 4 through 913 are equipped with separate transmission oil temperature and
transmission oil pressure indicators (9 and 10, figure 95-1) (3 and 14, figure 95-2).
h. Helicopters S/N 914 and subsequent are equipped with dual transmission oil temperature/pressure
indicators (14, figure 95-3), (16, figure 95-4)
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FLIGHT INSTRUMENTS.
The flight instrument system consists of the pitot-static system, airspeed indicator, altimeter, and
inclinometer
PITOT STATIC SYSTEM.
The pitot tube (1, figure 95-13) is mounted on a support (2), located on the most forward part of the cabin
nose structure just right of the helicopter centerline. This tube supplies impact air to the airspeed
indicator. Static air pressure for instrument operation is obtained from two static vents which are located
immediately forward of the crew doors and just below the windshields.
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AIRSPEED INDICATOR
The airspeed indicator (2, Figure 95-1, 6, Figure 95-2, 3, Figure 95-3 and 3, Figure 95-4) is a standard
pitot-static instrument. This single scale indicator provides an airspeed reading in miles per hour, and
knots by measuring the difference between impact air pressure from the pitot tube and the static air
pressure from the static vents.
NOTE
Maintenance on the pitot-static system (paragraph 95-10 through paragraph 95-14) is the only
recommended maintenance for the airspeed indicator. If the indicator itself is suspected of malfunction it
must be replaced.
ALTIMETER
The altimeter (14, Figure 95-1, 5, Figure 95-2, 5, Figure 95-3, and 6, Figure 95-4) provides a direct
reading of helicopter height in feet above sea level. This indicator is connected to the static air system to
sense atmospheric pressure. An external knob is provided to make compensation for variations of
prevailing barometric pressure.
INCLINOMETER
The inclinometer (11, Figure 95-2, 4, Figure 95-3 and 5, Figure 95-4) is a simple instrument consisting of
a covered glass tube, ball, and damping fluid. The ball indicates when the helicopter is in directional
balance, either in a turn or straight and level flight. If the helicopter is yawing or slipping, the ball will
move off of center.
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NAVIGATION INSTRUMENT.
The sole navigation instrument installed is the pilot standby magnetic compass (figure 95-14).
PILOT STANDBY MAGNETIC COMPASS.
The pilot standby magnetic compass (1, figure 95-14) is a standard, nonstabilized, magnetic type
instrument mounted on a support which is attached to the forward cabin right side. The compass is used
in conjunction with a compass correction card located below the compass.
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PROPULSION INSTRUMENTS
The propulsion instruments consist of the dual tachometer, gas producer tachometer, engine oil
temperature, engine oil pressure, transmission oil temperature, transmission oil pressure, engine
torquemeter, and turbine outlet temperature indicators.
PROPULSION INSTRUMENTS — DUAL TACHOMETER
The dual tachometer (Figure 95-15), indicating in percent, furnishes both rotor RPM and power turbine
RPM information. This instrument is powered by the rotor tachometer and power turbine tachometer
generators. The generators are self-generating andare not connected to the electrical system. Normal
operation of the helicopter is when rotor RPM and power turbine RPM needles are synchronized and in
the green arc. Refer to Figure 95-15 and Chapter 11 for instrument placards and markings.
GAS PRODUCER TACHOMETER
The gas producer tachometer (Figure 95-15), indicating in percent, furnishes gas producer RPM
information. This instrument is powered by the gas producer tachometer generator. This generator is a
self generating component. Refer to Figure 95-15 and Chapter 11 for instrument placards and markings.
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ENGINE OIL TEMPERATURE INDICATOR (HELICOPTERS S/N 4 THROUGH 913)
The engine oil temperature indicator (Figure 95-15) indicates the engine oil temperature in degrees
Celsius. This instrument is included in a bridge circuit with a resistor element and temperature bulb
located in the engine oil tank. The indicator and bulb are matched electrically in the bridge circuit and
require no calibration. Refer to Figure 95-15 and Chapter 11 for instrument placards and markings.
ENGINE OIL PRESSURE INDICATOR (HELICOPTERS S/N 4 THROUGH 913)
NOTE
Helicopters S/N 254 through 913 are equipped with electrically operated engine oil pressure indicators
with a transducer mounted in the piping circuit adjacent to the oil pressure switch. On helicopters S/N 4
through 913, the engine oil pressure indicator (11, Figure 95-1 and 13, Figure 95-2) provides indications
of engine oil pressure in pounds per square inch (PSI). The indicator reading is supplied by the
transducer in the piping circuit. Refer to Figure 95-15 and Chapter 11 for instrument placards and
markings.
TRANSMISSION OIL TEMPERATURE INDICATOR (HELICOPTERS S/N 4 THROUGH 913 AND
PRIOR)
The transmission oil temperature indicator (10, Figure 95-1 and 3, Figure 95-2) provides transmission
oil temperature reading in degrees Celsius. The indicator is used in the 28 VDC bridge circuit with a
temperature bulb that is located in the left side of the transmission. Refer to Figure 96-15 and Chapter 11
for instrument placards and markings.
TRANSMISSION OIL PRESSURE INDICATOR (HELICOPTERS S/N 913 AND PRIOR)
The transmission oil pressure indicator (9, Figure 95-1 and 14, Figure 95-2) is included in the piping
circuit to the transmission oil pressure disconnect at the lower firewall. This indicator indicates
transmission oil pressure in pounds per square inch (PSI) and is precalibrated (bench) against a
standard. Refer to Figure 96-15 for instrument placards and markings.
ENGINE OIL TEMPERATURE/PRESSURE
SUBSEQUENT)
INDICATION
(HELICOPTERS
S/N
914
AND
Engine oil temperature/pressure indication (Figure 95-15) is a dual instrument providing both
temperature and pressure indications. The temperature side of the instrument is part of the bridge circuit
with the resistor element of the temperature bulb located in the engine oil tank; it indicates engine oil
temperature in degrees Celsius. The indicator and temperature bulb are matched electrically in the
bridge circuit and do not require calibration. The pressure side of the instrument is precalibrated in
pounds per square inch (PSI) and is part of the piping installation, which provides direct (wet line)
readings from the engine to the indicator. Calibration of the system is not required. Bleeding of the
pressure gauge lines is required if air is allowed to enter the pressure lines. Refer to Figure 96-15 for
instrument placards and Markings.
TURBINE OUTLET TEMPERATURE (TOT) INDICATOR
Turbine Outlet Temperature (TOT) system consists of indicator (Figure 95-15), related wiring, and
engine thermocouples. Indicator is graduated in degrees Celsius and receives temperature indications
from thermocouples mounted in turbine exhaust outlet. On helicopter S/N 913 and prior, electrical power
for the indicator is generated by thermocouples and external or battery power is not required for
indications. On helicopters S/N 914 and subsequent electrical power for the indicator is supplied from the
28 VDC bus through the TOT circuit breaker. Refer to Figure 96-15 and Chapter 11 for instrument
placards and markings.
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CAUTION
TOT INDICATION ON S/N 914 AND SUBSEQUENT MAY NOT BE ACCURATE ON BATTERY
STARTS WHEN BATTERY POWER IS LOW OR DROPS TO 10 VOLTS.
Helicopters S/N 914 through 3366 (Post TB 206-82-61) and S/N 3367 and subsequent are equipped with
a red warning light on face of TOT indicator, which will illuminate if temperature of 812°F (+3, -0) is
exceeded for more than 10 seconds (+2, -0). This light will remain illuminated until circuit is reset with
TOT OVERTEMP LIGHT reset switch. The reset switch is a lockswitch that requires a key to actuate. It
is located adjacent to engine hourmeter in nose of helicopter. If circuit is not reset, light will illuminate any
time power is applied to indicator.
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MISCELLANEOUS INSTRUMENTS
The miscellaneous instruments installed consist of the DC loadmeter, free air temperature indicator,
eight day clock, and the hourmeter.
FUEL PRESSURE INDICATOR
The fuel pressure indicator (7, Figure 95-1, and 15, Figure 95-2) provides indication of amount of fuel
pressure in pounds per square inch. On helicopters S/N 913 and prior: circuitry for the indicator (Chapter
96) is located in the piping circuit to the fuel pressure disconnect. On helicopters S/N 914 and
subsequent: electrical circuitry is provided from the fuel pressure transducer to the indicator (Chapter
96). The newer indicator (11, Figure 95-3) is part of a dual indicator with the DC loadmeter.
DC LOADMETER
The DC loadmeter (8, Figure 95-1 and 1, Figure 95-2) measures and indicates generator output in
percentage. On helicopters S/N 914 and subsequent, the DC loadmeter (11, Figure 95-3 and 12, Figure
95-4) is part of a dual indicator with the fuel pressure indicator.
FREE AIR TEMPERATURE INDICATOR
The free air temperature indicator (1, Figure 95-17) is a simple probe type thermometer inserted through
a hole in the windshield and provides indication of outside ambient air temperature.
FUEL QUANTITY INDICATOR
The fuel quantity indicator (6, Figure 95-1, 2, Figure 95-2, 13, Figure 95-3 and 14, Figure 95-4) provides
indication of available fuel in fuel cell. It is calibrated in gallons and is part of the fuel system electrical
bridge circuit.
EIGHT DAY CLOCK.
The eight day clock (13, figure 95-1, 12, figure 95-2, and 9, figure 95-3) provides accurate indication of
time in hours and minutes, and has a sweep second hand pointer.
HOURMETER.
The engine hourmeter (2, figure 95-17) provides indication of engine operating hours throughout the life
of the engine.
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AVIONICS SYSTEMS.
Customer requirements vary so extensively that a basic avionics system is not provided. Various
systems are available in Bell installed kits. These kits are installed when ordered by the customer.
Intercommunications, navigation, communication, and identification functions are performed by the
avionics equipment. Operational and maintenance procedures are available from the various equipment
manufacturers.
Four basic systems exist: intercommunications, navigation, communication, and identification.
Intercommunications system (ICS) equipment provides communication capability among helicopter
crewmembers, and between crew and passengers. Navigation system equipment provides
VHFomnidirectional range (VOR) and localizer (LOC) indications, and provides automatic direction finder
(ADF) navigation indications. Communication system equipment provides two-way, VHF voice
communications between helicopter and ground or other aircraft. Identification system equipment
provides identification signals for air traffic control purposes.
CONFIGURATIONS.
Four basic systems exist: intercommunications, navigation, communication, and identification.
Intercommunications system (ICS) equipment provides communication capability among helicopter
crewmembers, and between crew and passengers. Navigation system equipment provides VHF
Helicopter equipment configurations are dependent on the Bell systems kits installed or customized
systems installed by other equipment manufacturers. Specific kit configurations and locations are
provided in figures 97-1 and 97-2. Only available Bell kit configurations are covered within each system
section.
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INTERCOMMUNICATIONS (ICS) SYSTEM
ICS SYSTEM.
Intercommunications system (ICS) equipment provide communications capability among crewmembers
and between crew and passengers. Additionally, ICS equipment provides audio signal control for all
avionics systems. Audio outputs to pilot/copilot headsets and cabin speakers are controlled by an audio
panel, if installed, or by avionics equipment volume controls. When not equipped with an audio panel,
audio-signal switching is provided by internal relays within helicopter ICS circuitry. Headset and handheld microphones are keyed for ICS communication by depressing cyclic-stick grip switches to the ICS
position, or by depressing the copilot ICS foot switch. ICS system components and circuitry provisions
are available in kits for installation. ICS components include the KX-170 and KX-170B transceivers or the
KMA-24H-52 and KMA-24H-71 audio panels. Internal relays are installed for use with the KX-170 and
KX-170B VHF transceivers to provide audio signal switching control.
COMMUNICATION SYSTEM
VHF COMMUNICATION SYSTEM.
Two-way voice communications are available on 720 channels at frequencies ranging from 118.000 to
135.975 MHz. VHF transceivers available in kits include the KX-170/KX-170B and KX-155. Both
transceivers are dual function providing VHF navigation capabilities as well as voice communications
(refer to Navigation Systems). On the KX-155 transceiver front panel, two communications frequencies
are displayed on the far left side: one active and one standby.
VHF NAVIGATION SYSTEM
VHF navigation capability provides VHF omnidirectional range (VOR) and localizer (LOC) information to
the pilot and copilot. Two VHF transceivers are available in avionics kits: the KX-170/KX-170B and the
KX-155. These transceivers are dual purpose providing two-way voice communications as well as
navigation capability. Both transceivers provide 200 navigation channels at frequencies ranging from
108.00 to 117.95 MHz. If the KX-170 or KX-170B VHF transceiver is installed, the KI-201 C indicator is
used to display VOR/LOC information. If the KX-155 transceiver is installed, the KI-208 indicator is used
to display course deviation information. On the KX-155 transceiver front panel, two navigation
frequencies are displayed on the far right side: one active and one standby.
AUTOMATIC DIRECTION FINDER (ADF) NAVIGATION SYSTEM
ADF navigation capability provides relative bearing to the selected transmitting station which allows
distance-to-station calculations to be made. Available kits provide two ADF receivers: the KR-85 and KR87. These receivers operate in a frequency range from 200 to 1699 KHz in 1.0 KHz steps. ADF
frequency, antenna, and beat frequency oscillator modes are the three ADF receiver operational modes.
If the KR-85 ADF receiver is installed, the KI-225 ADF indicator is used to display bearing indications. If
the KR-87 ADF receiver is installed, the KI-227 ADF indicator is used to display bearing indications.
IDENTIFICATION SYSTEM
Identification system transponders are designed to fulfill Air Traffic Control Radar Beacon System
(ATCRBS) requirements for an airborne beacon. Three transponders are provided for installation in
available avionics kits: the KT-75R and the KT-76176A and KT-79. A KFS-575 control unit is utilized by
the KT-75R transponder. These transponders receive interrogation signals from ground radar at 1030
MHz. A coded response is automatically generated to the ground receiving station at 1090 MHz. Specific
pulse sequences for 4096 preselected codes are assigned to helicopters or reserved for specific
occasions. These codes enable the air traffic controller to accurately and quickly identify the helicopter.
To further assist with rapid identification, the air traffic controller may request the pilot to identify the
helicopter. This is accomplished by depressing the IDENT switch on the transponder panel which causes
the ATC radar display blip to flash or "bloom".
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NOTE
FAA designated codes shall only be used for the specific purpose identified:
0000 and 7777 - never used;
1200 - VFR below 10,000 feet;
1400 - VFR above 10,000 feet;
4000 - restricted or warning area;
7500 - hijacking;
7600 - lost communications; and
7700 - inflight emergency (MAYDAY).
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FUEL
SYSTEM
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FUEL SYSTEM (Helicopter S/N 4 through 3566).
The fuel system (figure 28-1) incorporates a single bladder type fuel cell located below and aft of the
passenger seat (figure 28-1). Installed within the fuel cell are two electrically operated boost pumps,
lower and upper tank indicating unit and sump drain valve.
Boost pumps are interconnected and supply fuel through a single hose assembly to the fuel shutoff valve
and from the shutoff valve to the engine mounted fuel filter and pump. Boost pumps incorporate pressure
switches in discharge ports and drain plugs in the pump drain port. The fuel cell is filled from the right
side and has a capacity of 76 U.S. gallons (287.66 liters usable).
Access to boost pumps, lower tank unit and drain valve is from the bottom of fuselage and access to
upper indicating unit is gained from a cover plate located on deck aft of passenger seatback. Access to
fuel shutoff valve and vent line is in the fuel compartment located on the right side of access panel above
filler cap. Provisions are also made in fuel compartment for combustion heater fuel, fuel pressure
instrument line, and fuel pump purge line.
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FUEL SYSTEM (Helicopters S/N 3567 and subsequent).
The fuel system (figure 28-2) incorporates a single crash resistant bladder type fuel cell located below
and aft of the passenger seat. Installed within the fuel cell are two electrically operated boost pumps,
lower and upper tank indicating unit and electrically operated sump drain valve. Boost pumps are
interconnected and supply fuel through a single hose assembly to the fuel shutoff valve, and from shutoff
valve to the airframe mounted fuel filter.
Boost pumps incorporate pressure switches in discharge ports and drain plugs in pump drain ports. The
fuel cell is filled from the right side and has a capacity of 91 U.S. gallons (344.44 liters usable).
Access to boost pumps, lower tank unit and solenoid drain valve is from the bottom of fuselage and
access to upper indicating unit is gained from a cover plate located on deck aft of passenger seatback.
Access to fuel shutoff valve and vent is in fuel compartment located on the right side above filler cap.
Provisions are made in the fuel compartment for a fuel purging line to be installed at tank vent fitting for
maintenance purposes
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FUEL CELL (Helicopters SIN 4 through 3566).
Fuel cell is a bladder type unit installed in the fuselage cavity below and aft of passenger seat and laced
to helicopter structure.
FUEL CELL (Helicopters S/N 3567 and subsequent).
Fuel cell is a crash resistant bladder type fuel cell located below and aft of passenger seat structure. The
fuel cell is held to structure by screws.
FUEL DISTRIBUTION.
Fuel distribution consists of all fuel cell mounted components which transfer fuel, and monitor fuel
quantity, fuel flow, or fuel pressure.
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FUEL PUMP AND FILTER ASSEMBLY.
The engine fuel pump and filter assembly are integral units mounted on the aft end of engine. Fuel
enters engine fuel system at inlet port of the pump and passes through filter before entering gear
elements of pump. Filter draining is accomplished by a drain valve mounted on filter housing. Fuel filter
is monitored by a pressure differential switch located on lower firewall and connected electrically to fuel
filter caution light. Refer to Allison Engine Company Operation and Maintenance Manual (5W2 for C-18
engine or 10W2 for C-20 engine) for detailed maintenance instructions.
FUEL BOOST PUMP.
Two electrically operated fuel boost pumps are located in the bottom of fuel cell. Pumps are
interconnected and furnish fuel through one supply line. Pumps are equipped with check and thermal
relief valve, pump drain port, seal drain port, intake screen, and pump operating pressure switch located
in discharge port of pump. Pumps are protected by circuit breakers located in overhead console.
Fuel pump motor/impeller cartridge can be removed without removing fuel boost pump assembly. Refer
to paragraph 28-22 for replacement of motor/impeller cartridge.
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FUEL SHUTOFF VALVE.
A motor operated shutoff valve incorporating a thermal relief feature is installed in main fuel supply line
and is located in fuel compartment above fuel filler cap. Valve is electrically controlled by an ON-OFF
switch located on instrument panel and protected by a circuit breaker located in overhead console panel.
In event of electrical failure valve will remain in position selected before failure.
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FUEL PRESSURE TRANSDUCER.
The fuel pressure transducer provides a means of monitoring fuel pressure. The transducer is located on
a fitting on top of aft right side of fuel cell.
FUEL QUANTITY – INDICATING UNITS.
Two float type fuel level transmitting units (tank units) are installed in fuel cell. The lower unit is mounted
in the tank bottom and monitors fuel level up to horizontal surface of cell, under seat; upper unit monitors
fuel level in upper section of fuel cell, behind seat, and is mounted to top of fuel cell. Both indicating units
are connected to a common quantity indicator.
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SOLENOID VALVE.
The solenoid valve is an electrically and manually controlled drain valve used to drain fuel from fuel cell.
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CAP.
Cap and adapter assembly covers fuel access opening at top of fuel cell.
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AIRFRAME MOUNTED FUEL FILTER.
A fuel filter is mounted to structure on left side of engine compartment (right side of forward firewall on
helicopters S/N 3387 and subsequent). Fuel filter (figure 28-15) consists of a replaceable filter element,
drain valve, bypass valve, impending bypass switch, and manual test button. The airframe mounted fuel
filter assembly eliminates requirement for adding anti-icing additive to fuel supply when temperatures are
below 40°F (4°C). Indication of impending bypass lights the A/F FUEL FILTER segment on pilot caution
panel.
NOTE
Filter element must be replaced when caution light comes on during engine operation. Replace filter
element at same hourly interval as engine fuel filter maintenance is performed or 300 hours. (Refer to
Allison 250-C20 Series Operation and Maintenance Manual, 10W2.)
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HYDRAULIC
SYSTEM
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HYDRAULIC SYSTEM.
The hydraulic system provides power to operate cyclic and collective flight control servos.
NOTE
Helicopters S/N 4 through 497 have hydraulic powered tail rotor control systems which are also powered
by the hydraulic system. Helicopters S/N 498 and subsequent do not have hydraulic powered tail rotor
systems, unless a stabilization augmentation system is installed. The hydraulic system consists of the
pump/ reservoir regulator assembly, filter, solenoid valve, tube, and hose assemblies. Pump, regulator,
and reservoir are mounted on forward side of transmission oil pump as an assembly. The pump and
regulator assembly includes a mounting pad for rotor tachometer generator. Operation of the hydraulic
system is electrically controlled by an ON/OFF switch mounted on console for pilot control of the
solenoid valve. When solenoid is energized (ON/OFF switch "OFF'), pressurized hydraulic fluid flows to
the reservoir, bypassing servo actuators. Refer to figures 29-1 and 29-2 for system components and to
Chapter 98 for hydraulic system wiring diagram. Tail rotor servo is a customer option and may be
removed.
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HYDRAULIC SYSTEM COMPONENTS
Hydraulic system components consist of pump and reservoir (1, Figure 29-6), pressure regulator valve
(2) (located within pump and reservoir assembly), solenoid valve (3), hydraulic system filter (4) with
replaceable filter element, and quick-disconnect sockets (5 and 6).
HYDRAULIC PUMP AND RESERVOIR
The hydraulic pump and reservoir (1, Figure 29-6) is located on the forward end of the transmission.
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HYDRAULIC FILTER ASSEMBLIES
Two configurations of filter assemblies are utilized on the helicopters. Both utilize elements with a
filtration rating of 15 microns absolute. The unit utilized on helicopters S/N 154 and subsequent provides
a positive indication of restricted flow through the filter element by means of a red indicator button (3,
Figure 29-4) located on top of the filter body. The button pops up at 70 ±10 PSI (480 ±69 kPa)
differential, and is inoperative below 35° ±15°F (1.67° ±9.44°C).
HYDRAULIC FILTER ELEMENT
The hydraulic filter element (8, Figure 29-4) functions at high pressure over a wide temperature range.
Filtration rating is 15 microns absolute. When red indicator button (3) on return line filter pops up,
inspect, clean, or replace hydraulic filter element.
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HYDRAULIC SOLENOID VALVE
The solenoid valve (21, Figure 29-8) is incorporated in the hydraulic system for turning the system ON
and OFF. The solenoid valve is located forward of transmission work deck area. The solenoid valve is
normally de-energized; when HYDR SYSTEM circuit breaker is in and HYDRAULIC SYSTEM switch is
OFF, electrical power is applied to energize solenoid, which closes the valve and removes hydraulic
pressure from the servo actuators (Figure 29-2).
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QUICK-DISCONNECTS
Two quick-disconnect fittings provide a convenient means of connecting a hydraulic test stand to
helicopter. Each quick-disconnect is made up of a coupling and a coupling half (15 and 16, Figure 29-4).
When disconnected, each coupling automatically closes to prevent loss of fluid and/or entry of foreign
matter.
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ROTOR BRAKE SYSTEM
The rotor brake system provides a positive means of stopping main and tail rotors. The rotor brake is a
completely self-contained hydraulic system. It is operated by a handle located on the right side of the
overhead console. Rotor brake system is shown in Figure 29-9.
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MAIN ROTOR
SYSTEM
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MAIN ROTOR HUB AND BLADES.
The main rotor assembly is a two-bladed, semi-rigid, seesaw type rotor with underslung mounting. The
main rotor blades are all metal construction with an aluminum alloy honeycomb core, aluminum skins,
spar and trailing edge strip. All the structural components are joined by means of metal-to-metal
bonding. The main rotor hub consists primarily of a forged steel yoke with two spindles, a trunnion
assembly, two blade grips with pitch horns, and two grip-retention strap assemblies. Oil or grease
lubricated bearings provide for smooth rotation of the trunnion and blade grips on the yoke.
The blades are attached to the hub grips with bolts which have hollow shanks for installation of weights
for static balance of hub and blade assembly. After balancing, the bolts must be kept with their
respective rotor hub grips. Blade alignment is accomplished by adjustment of blade latches, which
engage the root end of the blade. 206-010-100-003 through -017 and 206-011-100-001 through -021
main rotor hub assemblies, prior to compliance with T.B. 206-78-5 or 206-79-21, are oil lubricated hub
assemblies. T.B. 206-78-5 and 206-79-21 modifies these hubs to grease lubricated hub assemblies.
Main rotor hub assemblies subsequent to 206-011-100-021 are grease lubricated hub assemblies. A flap
restraint is on some main rotor hubs. The flap restraint assembly incorporates counterweights and
springs which serve to position limited freedom flapping stops. The stops prevent excessive flapping of
the main rotor during starting and shutdown but allow normal flapping at operating rpm.
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FLAP RESTRAINT.
The flap restraint (3, figure 62-1) is mounted on the main rotor hub trunnion. It consists of counterweights
and springs which limit rotor flapping during starting and shutdown. Normal flapping is not restricted at
operating rpm.
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MAIN ROTOR BLADES
Main rotor blades are of all-metal construction with an aluminum alloy spar, spar spacer, trailing edge
strip, honeycomb core, and aluminum skins. All structural components are joined by metal-to-metal
bonding. The blades are set in hub grips at a preconed angle and are secured by a single retaining blade
bolt in each grip. An inboard trim tab and an outboard trim tab are provided on the trailing edge for
tracking adjustments. Earlier blades 206-010-200-033 have outboard trim tabs only. The blades are
individually interchangeable.
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PITCH-LINK.
The pitch link connects the pitch horn on the blade grip to swashplate outer ring, for control input from
collective and cyclic controls. A pitch link is required for each main rotor blade.
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SWASHPLATE AND SUPPORT.
The swashplate and support encircles the mast directly above the transmission and is mounted on a
universal support (pivot sleeve) which permits it to be tilted in any direction. Movement of the cyclic
control stick results in a corresponding tilt of the swashplate and the main rotor. Movement of the
collective pitch lever actuates the sleeve assembly which raises or lowers the swashplate and transmits
collective control to the main rotor. The cyclic controls are mixed with the collective control by action of
the mixing lever at the base of the control column (figure 62-15).
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COLLECTIVE LEVER AND LINK
The collective lever and link assembly is mounted to the swashplate support assembly and transfers
collective flight control inputs to the swashplate.
SWASHPLATE DRIVE ASSEMBLY
The swashplate drive assembly consists of a collar set, idler link, and idler lever. The collar set is
attached to the mast and the idler link is attached to the outer ring of the swashplate. The idler lever
connects between the collar set and idler link.
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MAIN ROTOR DRIVE
SYSTEM
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MAIN ROTOR DRIVE SYSTEM
The main rotor drive system provides a means of transmitting power from the engine to the main rotors.
The main rotor drive system consists of a transmission, main rotor mast, main driveshaft, freewheel
assembly, and oil cooler (Figure 63-1). The freewheel assembly is mounted on the engine accessory
gear case. It connects the engine to the transmission through the main driveshaft on the forward side,
and the tail rotor gearbox through related shafting on the aft side. This provides simultaneous rotation of
main and tail rotors and permits free rotation of both rotors when the engine is not operating.
MAIN ROTOR DRIVE SYSTEM — MAJOR CHANGES
Table 63-2 outlines the major changes and improvements that have occurred on the designated
helicopters main rotor drive system components. Due to the interchangeability of components,
maintenance instructions in this section shall include the component part number in the paragraph titles
and illustrations, as applicable.
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OIL CONTAMINATION OF TRANSMISSION
WARNING
IF FOREIGN PARTICLES ARE LARGE ENOUGH TO BE IDENTIFIED AS PART OF A COMPONENT
OF THE TRANSMISSION OR MAST BEARING, REPLACE THE TRANSMISSION OR MAST BEARING
AS INDICATED BY THE LOCATION OF THE PARTICLE.
CAUTION
WHEN TRANSMISSION OIL IS CONTAMINATED, THE OIL COOLER AND FREEWHEEL ASSEMBLY
MAY ALSO BE CONTAMINATED. THE TRANSMISSION OIL FILTER WILL NORMALLY PRECLUDE
THIS, BUT IF THE OIL FILTER HAS CLOGGED ENOUGH TO CAUSE BYPASSING, THEN
CONTAMINATION OF THESE COMPONENTS AND ASSOCIATED OIL LINES CAN OCCUR.
Some transmissions are equipped with a collector pan and chip detector immediately under the upper
mast bearing, and a second chip detector in the sump. As a result, it is possible to isolate the source of
metal particles to either the transmission or the main rotor mast assembly. If chips are found on the
upper chip detector, but not on the lower, a mast bearing problem is indicated. If chips are found on the
lower chip detector, but not on the upper, a transmission problem is indicated.
WARNING
IF FOREIGN PARTICLES ARE LARGE ENOUGH TO BE IDENTIFIED AS PART OF A COMPONENT
OF THE FREEWHEEL ASSEMBLY, REPLACE THE FREEWHEEL ASSEMBLY (PARAGRAPH 63-69
THROUGH PARAGRAPH 63-72).
CAUTION
IF FOREIGN PARTICLES ARE DISCOVERED IN THE
TRANSMISSION MAY OR MAY NOT BE CONTAMINATED.
FREEWHEEL
ASSEMBLY,
THE
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MAIN DRIVESHAFT
A driveshaft with splined couplings is installed between the freewheel assembly and the adapter flange
on transmission input quill.
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MAIN ROTOR MAST
The main rotor mast drives the rotor and support components required for directional changes.
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TRANSMISSION
The transmission provides a two-stage reduction of 15.23 to 1.0 (6000 to 394 RPM). The first stage is a
bevel gear arrangement with 3.26 to 1.0 reduction. The second-stage reduction is obtained with a
planetary gear train providing 4.67 to 1.0 reduction. A complete hydraulic system power pack is mounted
on the forward side of the transmission case and is driven by a transmission accessory drive gear. The
accessory drive gear provides 1.42 to 1.0 reduction. The transmission is mounted on the cabin roof
deck, forward of the power plant. The main rotor mast is secured in top of the transmission by the mast
bearing, bearing liner, and seal plate, and is isolated from the airframe by a system composed of two
pylon support links (one on each side), and a drag link secured to the bottom of the transmission and
connected to the rubber isolation support mount on the airframe. A cylindrical boss extends downward
from the forward end of drag link and fits loosely in a hole in the pylon stop mounted on the airframe,
providing a positive limit of travel of the pylon.
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TRANSMISSION OIL SYSTEM
Lubrication is provided by a system that includes an oil pump, pressure regulating valve, oil cooler, filter
element, and oil jets No. 1 and No. 2. The oil pump is a constant-volume type, driven by the accessory
drive gear, that delivers oil under pressure externally to the oil filter and housing assembly. The oil cooler
then returns oil to the main transmission and spray jets for lubricating internal parts. An oil level sight
gauge is located on the right side of the transmission lower case where it can be easily inspected. A nonvented filler cap is located on the transmission top case. The transmission oil system also provides
lubrication for the freewheel assembly mounted in the engine accessory gear case. Oil pressure for
lubrication of the freewheel assembly and to the transmission oil pressure gauge on the instrument panel
is taken from a tee fitting installed in the pressure line.
Oil temperature indications are provided by an oil temperature bulb located in the outlet side of the oil
filter housing, and a high oil temperature switch that is connected to the TRANS OIL TEMP caution
segment. The TRANS OIL PRESS caution segment light is connected to the transmission oil pressure
switch, located in the oil pressure line below and forward the instrument panel. A schematic
representation of the transmission oil system is presented in Figure 63-11.
OIL PUMP
A constant-volume 4.5 to 5 GPM transmission oil pump is mounted in the transmission inboard of the
hydraulic pump and tachometer generator, with all being driven by the accessory drive gear. The
tachometer generator and hydraulic pump must be removed to gain access to the transmission oil pump.
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OIL PUMP INLET SCREEN
The oil pump inlet screen is located in a boss, next to the oil pump on the transmission lower case, and
filters the lubricating oil before entering the oil pump. Screen material is 16 x 16 mesh, 23-gauge steel
that is brazed to an inlet sleeve and plug.
CHIP DETECTORS
NOTE
Prior to Service Letter 206-133, transmissions will contain magnetic chip detectors. Subsequent
transmissions incorporate electric chip detectors. Some transmissions will have two chip detectors and
some will have three.
NOTE
The oil monitor (31, Figure 63-10) is also a chip detector. Located on the transmission oil filter head, it
consists of the oil monitor and a removable magnetic plug (30). The chip detector is made up of the selfclosing valve and the chip detector. The self-closing valve also serves as a drain plug for the component.
The chip detector consists of a self-locking bayonet probe with a permanent magnet at the end. If metal
particles become free in the oil, the magnet will attract the metal particles, allowing for inspection. If the
chip detector is electric, when sufficient metal is attracted to complete the circuit between pole and
ground, the appropriate TRANS CHIP detector segment on the caution panel will illuminate.
OIL LEVEL SIGHT GLASS
The oil level sight glass is located on the right side of transmission and provides visual indication of oil
level.
OIL JETS
Two oil jets are incorporated in the main transmission. Oil jet No. 1 is installed in the transmission main
case and directs a lubricating spray to the spiral bevel gear and input pinion of the transmission. Oil jet
No. 2 is installed in the transmission top case and lubricates the planetary pinions and mast bearing.
OIL PRESSURE REGULATING VALVE
The transmission oil pressure regulating valve is located on the left aft corner of the transmission lower
case, just aft of the oil filter. The regulating valve is used to adjust transmission oil pressure to normal
operating limits and relieves excess oil pressure back into the transmission case.
OIL COOLER
The oil cooler contains a single core and is mounted to the top aft side of transmission. The thermal
bypass valve controlling flow of oil through the oil cooler core is located on outlet side of the oil cooler
and allows oil to bypass the oil cooler when oil temperature is low.
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FREEWHEEL ASSEMBLY
A shaft from the power turbine drives the engine power takeoff gearshaft through the engine accessory
gear case. The freewheel assembly is mounted on the engine accessory gear case and the inner race
shaft assembly is splined directly to the engine power takeoff gearshaft. Engine power is transmitted to
the outer race of the freewheel assembly, then through the full phasing sprag elements which couple the
engine power to the transmission driveshaft. The forward short tail rotor driveshaft connects, through a
flexible coupling, to a splined adapter on the aft end of the inner race shaft assembly that passes through
the engine accessory gear case. During autorotation, the main rotor drives the power input shaft. Under
thiscondition, the freewheel assembly provides a disconnect from the engine so that the rotational forces
of the main rotor are free to drive the transmission, tail rotor, and all transmission-mounted accessories.
Refer to Figure 63-14 for different configurations of freewheel assemblies.
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TAIL ROTOR
SYSTEM
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TAIL ROTOR
The tail rotor hub and blade assembly consists of two blades and hub assembly. The blades are
attached to the hub by two blade mounting bolts per blade which are inserted through spherical
bearings; spherical bearings are inserted in the blade root end, on the pitch change axis. The spherical
bearings permit pitch change of the blades. The pitch link assemblies attached to the pitch horn
assemblies and the crosshead assembly set the pitch angle of the blade assemblies through the action
of the pitch change mechanism, and flight controls. The splined trunnion assembly installed in the hub
and blade assembly enables the hub and blade assembly to mount onto the gearbox splined output shaft
and is attached to the shaft by a retaining nut. The splined trunnion also provides flapping axis
movement for the hub and blade assembly.
The tail rotor gearbox drives the tail rotor at a speed of approximately 2550 RPM. The tail rotor assembly
acts in opposition to the torque applied to the helicopter by the main rotor assembly. The tail rotor
assembly provides directional control to the helicopter around the vertical axis of the helicopter. The 206011-819 yoke assembly of the hub is made with a 4° twist for each blade. The twist in the yoke provides
additional thrust for high altitude performance. Balancing of tail rotor blades is made possible by adding
or removing balancing hardware that is attached to the balance wheel or the blade mount bolts.
TAIL ROTOR HUB AND BLADES
During flight, the tail rotor hub and blades (tail rotor) counteract the torque of the main rotor. The pitch of
the tail rotor blades is varied by means of the tail rotor control system. The tail rotor is mounted on the
left side of the tailboom and rotates at 2550 RPM. Figure 64-1 illustrates, unassembled, various
differences of the tail rotor hub and blade assemblies. Table 64-1 illustrates various configurations and
their effectivity applicable to each model.
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TAIL ROTOR BLADES
The tail rotor blade is an all-metal assembly consisting of a stainless steel shell reinforced by a
honeycomb filler and stainless steel leading edge abrasive strip. Two spherical bearings are installed in
an aluminum alloy retention block to provide for pitch change movement of the blade in the tail rotor hub.
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TAIL ROTOR DRIVE
SYSTEM
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TAIL ROTOR DRIVE SYSTEM.
The tail rotor drive system provides a means of transmitting power from the transmission to the tail rotor.
The tail rotor drive system includes the following components: tail rotor driveshaft, oil cooler blower, and
tail rotor gearbox.
OIL CONTAMINATION.
Particles of foreign material found in tail rotor gearbox electric chip detectors or in oil drained from
system may indicate that parts have failed. They are not necessarily an indication that the component is
no longer serviceable (figure 65-3). The quantity, source, form, type of material found, and service
history of component must be taken into consideration. The service time accumulated since new or since
overhaul, previous failures, and type of operation are important factors in determining further
serviceability of component. The parts may be steel, silver, aluminum, magnesium, bronze, or phenolic.
Procedure for identification of foreign material is described in the steps following.
WARNING
WHEN FOREIGN PARTICLES ARE LARGE ENOUGH TO BE IDENTIFIED AS PART OF A
COMPONENT OF THE TAIL ROTOR GEARBOX, REPLACE THE GEARBOX. WHEN SMALL
AMOUNTS OF METAL PARTICLES ARE FOUND IN TAIL ROTOR GEARBOX COMPONENTS, OR
THERE IS SOME DOUBT ABOUT THE SUITABILITY OF THE COMPONENT FOR CONTINUED
SERVICE, PERFORM A SERVICEABILITY CHECK (PARAGRAPH 65-8). SERVICEABILITY CHECKS
ARE A SPECIFIC REQUIREMENT WHEN DIRECTED IN THE TROUBLESHOOTING PROCEDURES.
OIL COOLER BLOWER.
The oil cooler blower assembly is mounted on the upper structure, aft of the aft firewall and is driven by
the tail rotor driveshaft. The squirrel cage type impeller is mounted on a flanged shaft which is mounted
in bearing hangers. The oil cooler shaft connects to the forward and aft short tail rotor shafts and is part
of the tail rotor drive system. The oil cooler blower provides cooling air for the engine oil system,
transmission oil system, and the hydraulic system. The engine oil cooler mounts above the blower
housing while a flexible duct conveys cooling air forward to the transmission oil cooler and the hydraulic
reservoir.
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TAIL ROTOR DRIVESHAFTS.
On helicopters S/N 4 through 1251, the tail rotor driveshaft is made up of the following sections: forward
short shaft, oil cooler blower shaft, aft short shaft, and the long shaft. Steel laminated flexible couplings
requiring no lubrication are used to connect the shaft sections and the tail rotor gearbox. The long tail
rotor driveshaft is designed to have a bend in the shaft between the first and second, and second and
third bearing supports.
On helicopters S/N 1252 and subsequent, the tail rotor driveshaft is made up of the following sections:
forward short shaft, oil cooler blower shaft, aft short shaft, and tail rotor driveshaft segments. Steel
laminated flexible couplings requiring no lubrication are used to connect the shaft sections and the tail
rotor gearbox.
FORWARD AND AFT SHORT SHAFTS.
The forward short shaft and aft short shaft are located on either side of the oil cooler blower assembly.
The forward short shaft is constructed of steel and is connected to the aft end of the freewheeling
assembly and forward end of the oil cooler blower shaft by splined adapters and steel laminated flexible
couplings. The aft short shaft is constructed of aluminum alloy and is connected to the aft end of the oil
cooler blower shaft and to the long tail rotor driveshaft or to the first tail rotor driveshaft segment by
splined adapters and steel laminated flexible couplings.
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LONG TAIL ROTOR DRIVESHAFT
The long tail rotor driveshaft consists of an aluminum driveshaft with five bearing hangers. The driveshaft
extends along the top of the tailboom connected between the aft short shaft and the tail rotor gearbox.
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TAIL ROTOR GEARBOX ASSEMBLY.
The tail rotor gearbox contains 90 degree spiral bevel gears providing a speed reduction of 2.35 to 1.0.
The bevel gear quill assemblies are designed to controlled dimensions to provide interchangeable
replacement of shaft assemblies without shimming. The housing is a magnesium casting attached to the
fuselage structure with four studs. A breather type filler cap, magnetic drain plug and oil level sight gage
are accessible from ground level (figures 65-14 and 65-15).
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ELECTRIC CHIP DETECTOR ASSEMBLY
One electric chip detector assembly is installed in the tail rotor gearbox. The electric chip detector
assembly is made up of one self-closing valve and an electric chip detector.
The electric chip detector consists of a self-locking bayonet probe with a permanent magnet at the end.
Free ferrous metal particles in the oil will be attracted to the magnet and when sufficient metal is
attracted to complete the circuit between pole and ground, the T/R CHIP detector segment on the
caution panel will illuminate.
The self-closing valve automatically closes and prevents loss of oil when the electric chip detector is
removed for inspection. The self-closing valve also serves as a drain plug.
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FLIGHT
CONTROLS
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FLIGHT CONTROLS.
The flight control system consists of push-pull control tubes and bellcranks actuated by conventional
helicopter cyclic, colletive, and directional controls (figure 67-1). The controls are routed beneath the pilot
seat aft to the center of the helicopter and up to the cabin roof through the control column which also
serves as a primary cabin structure. Access doors located on the aft side of the control column and
removable seats are provided for control inspection and maintenance accessibility. Cyclic and collective
controls are routed to the main rotor blades through the swashplate. The directional controls are routed
through the tailboom to the tail rotor. Fixed length control tubes and a minimum of adjustable tubes
simplify rigging. All self-aligning bearings and rod ends are spherical bearings and require no lubrication.
Cyclic, collective, and directional control systems incorporate hydraulic servo actuators. The servo
actuators on the cyclic and collective controls incorporate irreversible valves which prevent control force
feedback.
A stabilization augmentation system kit may be installed for the cyclic and directional control system.
Refer to Chapter 99 and Service Instruction for maintenance. Copilot controls (figure 67-1) are provided
as optional equipment for dual control capability and operations requirements. Dual control installation
consists of copilot collective stick, cyclic stick, and tail rotor control pedal assembly which are connected
to the pilot controls via jackshaft tube, torque tube, control tubes, and bellcranks. Copilot collective and
cyclic control sticks feature quick-disconnects for rapid removal or installation of control sticks.
CONTROL TUBES.
Aluminum alloy control tubes are used throughout the collective, cyclic, and antitorque controls. Some
control tubes are fixed length with fixed rod ends, while others have adjustable rod ends that are readily
replaceable. Removal and installation procedures relating to the individual control tubes are contained in
the applicable section for the controls.
BELLCRANKS, LEVERS, SUPPORTS, AND WALKING BEAMS.
Bellcranks, levers, supports, and walking beams are used throughout the collective, cyclic, and tail rotor
control systems. These parts transmit or control change movements in the particular system in which
they are installed. Removal and installation procedures relating to these parts are contained in the
applicable section for the controls.
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COLLECTIVE PITCH CONTROLS
The collective pitch control system consists of a jackshaft assembly with a control stick, push-pull tubes,
bellcranks, and a hydraulic servo actuator connected to a control lever on the swashplate support.
Movement of the control stick is transmitted through linkage and servo actuator to the main rotor pitch
control mechanism, causing the helicopter to ascend or descend, or to remain at constant altitude. The
servo actuator has an irreversible valve to reduce feedback, and to provide for use of controls in event of
hydraulic boost failure.
For helicopters with dual controls (Figure 67-1), the copilot collective stick is installed at left of copilot
seat. A fully functioning twist-grip throttle control is included in the copilot collective stick. A quickdisconnect feature permits rapid removal of the copilot collective stick. A spring pin assembly is provided
to ensure positive engagement of the stick. Switches are not installed on the stick.
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PILOT COLLECTIVE STICK
The pilot collective stick is installed at left of pilot seat. Stick extends upward and forward through a
flexible cover. Stick incorporates a twist grip for operation of gas producer controls. Switches are
installed on top of stick for starter, governor, RPM, landing light, and idle stop released.
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COPILOT COLLECTIVE STICK.
The copilot collective stick is installed at left of copilot seat. Stick extends upward and forward through a
flexible boot and incorporates a twist grip for operation of gas producer controls. No switches are
installed on copilot collective stick.
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OTHER COMPONENTS OF COLLECTIVE CONTROLS
COLLECTIVE JACKSHAFT.
The collective jackshaft provides a mounting point for the collective stick. An adjustable friction bearing
mounted on the jackshaft allows pilot to adjust friction for his own requirements. A minimum friction
adjustment clamp, located at the left end of the jackshaft, ensures that the collective stick will always
have a preset minimum friction.
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COLLECTIVE PITCH CONTROL LINKAGE.
Linkage between collective pitch control jackshaft and collective lever on swashplate support consists of
push-pull tubes, bellcranks, and hydraulic servo actuator assembly. Linkage from jackshaft assembly to
servo actuator is shown in figure 67-3. Linkage from servo actuator to collective sleeve lever on
swashplate support is shown in figure 67-3.
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CYCLIC CONTROLS.
The cyclic control system utilizes a linkage system to transmit movement to a swashplate, which
actuates rotating controls of main rotor to control helicopter attitude and direction. Fore, aft, and lateral
control use independent linkages from control stick to an intermixing bellcrank. From this point on,
linkage to swashplate horns cannot be considered separately. Two hydraulic servo actuators are
incorporated to reduce effort required for control and to reduce feedback forces from main rotor.
For helicopters with dual controls, the copilot cyclic stick is installed in front of the copilot seat. A quickdisconnect feature permits rapid removal or installation of the stick. A pin assembly is provided to ensure
positive engagement of the stick.
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CYCLIC STICK AND TORQUE TUBE.
The cyclic stick extends upward and forward from the front of the pilot seat. Switches are installed on the
stick grip for the intercom system and radio. The torque tube connects to the cyclic stick support.
NOTE
For helicopters equipped with dual controls, the copilot cyclic stick extends upward and forward from the
front of the copilot seat.
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OTHER COMPONENTS OF CYCLIC CONTROL
CYCLIC YOKE AND MIXING LEVER.
The cyclic system yoke extends aft from the cyclic stick support and torque tube. Movement of the cyclic
stick is transmitted by the yoke to the mixing lever. The mixing lever transmits cyclic movement to the
swashplate through mechanical linkage and servo actuators.
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TAIL ROTOR CONTROLS.
The tail rotor (antitorque) control system includes control pedals, pedal adjuster, push-pull tubes,
bellcranks, and a pitch control mechanism mounted through the tail rotor output shaft. Actuation of
pedals causes pitch change of tail rotor blades to offset main rotor torque and provides directional
control of helicopter.
For helicopters with dual controls, the copilot tail rotor control pedal assembly is installed on the floor in
front of the copilot seat. Two fully functioning control pedals are included in the assembly. Control pedals
are linked to pilot control pedals by means of control tubes and a bellcrank.
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TAIL ROTOR PEDALS AND ADJUSTER
Tail rotor control pedals mounted on the pilot and copilot compartment deck are connected under the
center console to a bellcrank pedal adjuster which provides for manual adjustment of pedal position,
according to pilot and copilot needs.
NOTE
For helicopters equipped with dual controls, the copilot pedal assembly provides a means for the copilot
to control the tail rotor assembly. The pedals can be positioned as desired by means of the pedal
adjuster.
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TAIL ROTOR CONTROL LINKAGE.
Connecting linkage consists of push-pull tubes, bellcranks, levers, and supports that connect pilot or
copilot tail rotor control pedals to the tail rotor pitch change mechanism.
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TAIL ROTOR PITCH CHANGE MECHANISM
The tail rotor pitch change control is accomplished by means of bellcrank, rod, and lever assembly
mounted on tail rotor gearbox, actuating a control tube through a hollow gearbox output shaft to the
crosshead and pitch links.
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CYCLIC AND COLLECTIVE SERVO ACTUATORS AND SUPPORT
CYCLIC AND COLLECTIVE SERVO ACTUATORS AND SUPPORT
On helicopters S/N 4 through 497, the cyclic and collective control servo actuator support is installed on
the cabin roof. It serves as a mount for servo actuators and associated bellcranks. Collective control
servo actuator is mounted in center position, and two cyclic servo actuators are mounted in outboard
positions.
Cyclic and collective servo actuators reduce the operational loads of these flight control systems. Servo
actuators in the cyclic and collective systems incorporate irreversible valves. In the event of loss of
hydraulic pressure to a servo actuator, the plunger (12, Figure 67-29) in the sequence valve (3) is
pushed up by the lower spring and poppet valve (10); the upper spring holds the valve seat (11) down.
This action closes the hydraulic return port and maintains irreversibility independent of hydraulic system
pressure. The pilot is provided with safe control of the helicopter even though hydraulic pressure is lost.
The sequence valve also serves to relieve thermal pressure buildup should this occur while the system is
inactive. The sequence valve would normally be closed when system pressure is below 100 to 180 psi
(689.00 to 1241.00 kPa). If internal pressure builds up, the valve seat is pushed up, compressing the
upper spring. The poppet valve on the lower spring is prevented from following by an internal obstruction
in the valve, shown as a line above the poppet valve in the schematic. The differential relief valve (4)
serves to relieve pressure buildup which could occur when slide and sleeve assembly (7) is centered
with both return ports closed.
On helicopters S/N 498 and subsequent, the cyclic and collective control servo actuator support is
installed on cabin roof. It serves as a mount for the servo actuators and associated bellcranks. The
collective control servo actuator is mounted in the center position, and the two cyclic servo actuators are
mounted in the outboard positions. Cyclic and collective servo actuators reduce operational loads of the
flight control systems. An irreversible valve is incorporated in each servo valve.
In event of loss of hydraulic pressure to a servo actuator, plunger (12) in the sequence valve (3) is
pushed up by the lower spring and poppet valve (10); the upper spring holds the valve seat (11) down.
This action closes the hydraulic return port and maintains irreversibility independent of hydraulic
pressure. This provides safe control of the helicopter even though hydraulic power is lost. The sequence
valve also serves to relieve thermal pressure buildup should this occur while the system is inactive. The
sequence valve would normally be closed when system pressure is below 100.00 to 180.00 psi (689 to
1241 kPa). If internal pressure builds up, the valve seat is pushed up, compressing the upper spring. The
poppet valve on the lower spring is prevented from following by an internal obstruction in the valve,
shown as a line above the poppet valve in the schematic. Differential relief valve (4) serves to relieve
pressure buildup which could occur from excessive rotor loads.
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ANTITORQUE SERVO ACTUATOR AND SUPPORT
The antitorque control servo actuator support is installed in the fuselage above and aft of the baggage
compartment. It serves as a mount for the antitorque servo actuator and associated lever assembly. The
servo actuator reduces the operational loads of the antitorque control system but does not incorporate
the irreversible feature used on the cyclic and collective servo actuators. Refer to Figure 67-35 for
schematic diagram.
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POWERPLANT
SYSTEM
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POWER PLANT SYSTEM.
The Allison Model 250-C20B and 250-C20J turboshaft engines are installed in the 206B and 206B3
helicopters. The engines consist of a single-stage centrifugal-flow compressor, a single combustion
chamber, a two-stage gas producer turbine, and a two-stage power turbine. The power plant assembly is
mounted horizontally aft of the transmission and above the fuselage. The engine is supported by three
bipod mounts attached to the service deck and is coupled to the transmission through the freewheeling
unit and main driveshaft.
The engine is rated at 470 SHP. Ai is delivered into the compressor assembly which contains a rotor
assembly that pressurizes the air. This pressurized air is delivered to the combustion chamber trough
two compressor air discharge tubes. A fuel nozzle sprays fuel into the combustion liner. This fuel is
mixed with the pressurized air and ignited. The exhaust gases expand trough the turbine rotors and into
the exhaust collector which exhaust the gas trough the two exhaust stacks and away from the aircraft.
System Limits:
Gas producer rpm (N1)
Continuous operation
Maximum
Maximum Transient
(Do Not exceed 10 seconds above 105%)
60 to 105%
105%
106%
Torque
Continuous operation
0 to 85%
Maximum continuous
85%
5 Mim Takeoff
100%
Maximum Takeoff
100%
Maximum transient
110%
(Do Not exceed 5 seconds above 100% intentional use is prohibited)
Turbine outlet temperature (TOT)
Continuous operation
Maximum continuous
5 Mim Takeoff range
Maximum Takeoff
Maximum during power transient
(Do Not exceed 10 seconds above 810ºC)
Maximum for starting
100 to 738ºC
738ºC
738º to 810ºC
810ºC
810ºC
927ºC
Power turbine
Minimum
Continuous operation
Maximum continuous
97%
97 to 100%
100%
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ACCESSORY LOCATION
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COWLING AND FAIRINGS.
The engine and transmission cowling (figure 71-3-3) consists of four sections: forward fairing, induction
fairing, engine cowl assembly, and aft fairing. The cowling is constructed of aluminum alloy, fiberglass,
and honeycomb material and is readily removable for engine and transmission changes. Cowling access
panels are provided with snap-open fasteners which permit inspection without removing the cover unit.
The forward and aft fairing assemblies are secured with fasteners.
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ENGINE MOUNTS
The engine is supported on the service deck by three bipod mounts located on the right, left, and lower
side of engine. Shims are provided at each mount leg for engine alignment.
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COMPRESSOR ASSEMBLY
The compressor is a six stage axial, single stage centrifugal flow compressor. Its primary function is to
take in ambient air, pressurize the air, and deliver it to the combustion assembly.
The compressor assembly consists of a compressor front support, case assembly, rotor wheels with
blades, centrifugal impeller, front diffuser assembly, rear diffuser assembly, diffuser vane assembly and
diffuser scroll. The five struts in the compressor inlet serve several functions. They both direct and
distribute air into the compressor inlet rotor in an efficient manner. When the anti-ice system is turned on,
hot compressor air is directed through the struts into the compressor inlet to prevent icing. The strut also
serve as a passage way to lubricate number 1 bearing
The compressor takes air from the inlet and is compressed by six stage axial and one centrifugal stage
to increase the air pressure and air temperature. The impeller discharges air into the vanes of diffuser.
The diffuser vanes direct air into the diffuser scroll. The diffuser scroll collects the compressor outlet flow
at constant velocity and directs the air into the transfer tubes which direct the air into the combustion
assembly.
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AIR FLOW SCHEMATICS
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COMBUSTION ASSEMBLY
The primary function of the combustion assembly is to mix pressurized air with fuel, then ignite the
mixture. After the ignition process, which causes the gases to expand, they are directed to the turbine
assembly.
The combustion assembly consists of two compressor discharge tubes, combustion outer case and
combustion liner. The combustion outer case is tapped with bosses for mounting a burner drain valve,
fuel nozzle, and one spark igniter. The burner drain valve threads into the boss and will ensure
appropriate draining when the engine is not operation. The fuel nozzle and spark igniter are on the rear
and extend into the center if the combustion liner. The combustion outer casing is flanged on the front for
mounting the combustion assembly to the gas producer turbine support. The combustion liner must
provide for rapid mixing of fuel and air, it must control the flame length and position to ensure that the
flame does not contact any metallic surface. The two air discharge air tubes form ducts that transfers
compressor discharge air from the scroll to the outer combustion case where it is directed to the
combustion liner. Approximately 75% of the air entering the combustion assembly is used for cooling,
while the remaining 25% is used for combustion.
N1 GAS PRODUCER GEAR TRAIN
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N2 POWER TURBINE GEAR TRAIN
TURBINE ASSEMBLY
The basic function of the turbine assembly is to take the energy (expanding gases) developed in the
combustion assembly and direct it over the two stage gas producer, two stage power turbine wheels,
and convert the expanding gas energy into the shaft horsepower.
This turbine assembly consists of a gas producer turbine support assembly, gas producer turbine rotor
assembly, power turbine support assembly, power turbine rotor assembly, and the exhaust collector
support assembly.
It incorporates the components necessary for the development of power and exhausting of gases. The
turbine has two-stage gas producer turbine and two-stage power turbine. Power to drive the compressor
rotor is furnished from the gas producer turbine rotor through direct drive. The power turbine converts the
remaining gas energy into the power which is delivered into the exhaust collector.
The gas producer turbine consists of the 1 st and 2nd turbine stages, and the power turbine consists of the
3rd and 4th turbine stages. The gas producer and power rotors are not “mechanically” coupled, but the
are “gas” coupled, in that the exhausting gases must flow through the four turbine stages. The power
turbine is a “free-turbine” since it is free to rotate at different speed than gas producer turbine rotor.
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ACCESSORY GEARBOX ASSEMBLY
The basic function of the accessory gearbox is to reduce turbine rotor speeds down to usable speed for
the output gearshaft, and various accessories mounted and driven by the accessory gearbox.
The accessory gearbox is the primary structural member of the engine as its provides mounting and
support for the compressor and turbine assemblies. The accessory gearbox contains most of the
lubrication system components and incorporates two separate gear trains. The purpose of the power
turbine gear train is to reduce the engine speed from 30,650 rpm at the power turbine rotor to 6,016 rpm
at the output shaft. The power turbine tachometer to measure engine output shaft. The power turbine
gear train incorporates a torquemeter to measure engine output torque. The power turbine tachometer
generation, power turbine governor, and spare drive gearshaft are driven by turbine gear train. The gas
producer gear train provides drive for the oil pump, fuel pump, gas producer fuel control, gas producer
tachometer generator and starter/generator. During starts, the starter/generator rotates the engine
trough the gas producer gear train.
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TORQUEMETER
TORQUEMETER
The torquemeter is incorporated in the accessory gearbox to provide a pressure signal which is directly
proportional to output torque. Pressure in the oil chamber is directed to the torquemeter pressure
sensing port on the front side of the accessory gearbox. The filtered oil is directed to the aircraft’s
torquemeter indicator located on the instrument panel.
The toque indicator is a “wet” gage that will require bleeding if air enters the system. The indicator has a
green band from 0 to 85% which is continuous operation, yellow band from 85 to 100% for takeoff power
range. The indicator also has a red line at 100% for maximum operation ( 5 minute limit).
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COMPRESSOR BLEED AIR SYSTEM
The compressor bleed air system is a entirely automatic system which permits rapid engine response.
The system consists of a bleed air control valve attached to the compressor case and necessary
plumbing between the diffuser scroll and the bleed air control valve.
The bleed control valve is open during starting an ground idle operation, and it remains open until a
predetermined pressure ratio is obtained. At this pressure ratio, the valve begins to modulate form open
to the closed position. it will be open during start cycle and ground idle, will modulate closed during
acceleration to full operational speed, and will remains closed during flight operation speeds.
Compressor discharge air pressure sensing for bleed control valve operation is obtained at sensing port
on the right front side of diffuser scroll.
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ANTI-ICING SYSTEM.
The compressor inlet guide vanes and front bearing support hub are the only engine components with
anti-icing provisions. Anti-icing is provided by the use of compressor discharge air from the engine
compressor. An air shut off valve, actuator, and bleed air valve are mounted at the 12 o'clock position of
the engine compressor to control anti-icing air.
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ANTI-ICING ACTUATOR.
The anti-icing actuator is an electrically controlled motor which positions the anti-icing valve through a
mechanical linkage. The anti-icing valve is mounted on the top of the engine on a mounting pad.
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ENGINE CONTROLS
Engine controls consist of the gas producer controls (N1) and the droop compensator controls (N2)
(Figure 76-1). The gas producer controls are operated by a twist grip on the collective stick. The droop
compensator controls are operated from a bellcrank in the collective system.
In addition, applicable information is contained in the chapter for helicopters upgraded per Service
Instruction 206-80 and 206-112. Service Instruction 206-80 upgrades 206A helicopters equipped with
250-C18 engines to a 206B helicopter by installation of a 250-C20 engine. Refer to service instruction for
specific details.
Service Instruction 206-112 upgrades 206A and 206B helicopters to a 206B-3 helicopter configuration by
installation of a 250-C20B engine. Refer to service instruction for specific details.
The engine control systems referenced in this chapter are BENDIX and CECO.
• BENDIX controls are approved for use on Rolls-Royce 250-C18, 250-C20, 250-C20B, 250-C20J, and
250-C20JN engines.
• CECO controls are approved for use on Rolls-Royce 250-C20 and 250-C20B engines.
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ENGINE FUEL SYSTEM
The fuel system consists of the fuel pump assembly, gas producer fuel control, power turbine governor,
check valve assembly and accumulators, and the fuel nozzle.
The Fuel pump assembly incorporates a single gear type pumping element and a bypass pressure
regulating valve. When the engine is in operation, the gas producer fuel control bypasses fuel back to
the pump assembly. The regulator valve controls the bypass fuel pressure.
The gas producer fuel control and the power turbine governor provide for a fuel metering system. To
regulate and to maintain fuel flow, this system senses gas producer rpm, power turbine rpm, compressor
discharge pressure, and twist grip position.
The double check valve assembly and two accumulators, located in the pneumatic line between the
power turbine governor and the governor reset section of the gas producer fuel control, are incorporated
to dampen torsional vibrations encountered in the helicopter rotor systems.
The fuel nozzle has a single entry and dual outlet orifice. This nozzle provides a finely atomized spray of
fuel at all flow conditions that are required by the engine. It is designed to provide an optimum spray
angle for starting the engine, plus even distribution of fuel into the combustion liner, and it is equipped
with a filter to minimize the possibility of contamination.
The fuel control system controls engine power output by controlling the N1 gas producer speed. These
speed levels are established by action of the power turbine governor, which senses N2 power turbine
speed. The power required to maintain this speed is automatically maintained by power turbine governor
action in metered fuel flow. Thus, the power turbine governor lever schedules the power turbine governor
requirements; and the power turbine governor, in turn, schedules the gas producer speed to a changed
power output to maintain output shaft speed.
Fuel flow for engine control is established as a function of compressor discharge pressure, engine speed
(N1 gas producer and/or N2 power turbine), gas producer lever angle and power turbine governor lever
angle. Fuel flow is a function of compressor discharge pressure as sensed in the fuel control.
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GAS PRODUCER CONTROLS (N1)
The gas producer controls (N1) consist of a flexible control cable assembly, which extends from the
throttle arm on the rear of the collective stick to a bellcrank assembly mounted on the engine deck. A
tube assembly is connected between the bellcrank and a power lever mounted on the fuel control shaft.
The twist grip, mounted on the end of the collective lever, controls the position of the gas producer fuel
control which has three positions: closed, idle (flight idle), and full open.
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DROOP COMPENSATOR CONTROLS
The droop compensator control system consists of a mechanical linkage between an idler in the
collective system and a lever mounted on the power turbine governor shaft. Movement of collective stick
results in repositioning of governor shaft. This action provides droop compensation to prevent N2 rpm
variations as power changes are made. The system incorporates a linear actuator which is controlled
electrically by a GOV RPM INCR-DECR (beep) switch mounted on collective stick. Main rotor rpm is
kept constant with power changes.
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ENGINE OIL SYSTEM.
The engine oil system is a dry sump type with an externally mounted oil supply tank and oil cooler
located on top aft section of fuselage. Oil is supplied by tank to gear type pressure and scavenge pump
mounted within the engine accessory gearbox. Return oil is routed from engine oil outlet port to oil cooler
and from cooler to the tank.
The dry sump lubrication system is furnished with a external reservoir and heat exchanger. A gear type
pressure and scavenge pump assembly is mounted within the gearbox. An assembly containing an oil
filter element, a filter bypass valve, and a pressure regulating valve, accessible from the top of the
engine, is located in the upper right-hand side of the gearbox, prevents oil from draining into the engine
from the aircraft tank when the engine is not in operation. Magnetic chip detectors are installed at the
bottom of the gearbox and at the engine oil outlet connection. All engine oil system lines and
connections are internal except the pressure and scavenge lines to the compressor front bearing and the
bearings in the gas producer and power turbine supports.
The lubrication system is designed to furnish adequate lubrication, scavenging, and cooling as needed
for bearings, splines, and gears regardless of helicopter attitude or altitude. Jet lubrication is provided to
all compressor, gas producer turbine, and power turbine rotor bearing, and to bearing and gear meshes
of the power turbine gear train with the exception of the power output shaft bearing. The power shaft
bearing and all other gears and bearing are lubricated by oil mist.
Oil from the external tank is delivered to the pressure pump which sends oil through the oil filter and then
to various points of lubrication. The check valve is open pressure, 115-130 psig, is regulated to this
relatively high value in order to balance the high axial gear thrust in the torquemeter, which is necessary
to minimize friction effects and to provide an accurate measurement of torque.
ENGINE OIL TANK.
The engine oil tank has a normal capacity of 1.5 U.S. gallons (5.68 liters) and oil level is checked with a
dipstick mounted on the filler cap. The tank provides port openings for engine supply, engine return,
vent, oil temperature bulb, oil level sight gage, and a self-locking drain valve.
ENGINE OIL COOLER.
Oil cooler is mounted on top of duct attached to oil cooler blower. When oil cooler bypass valve is open,
return oil from engine flows through oil cooler and returns to oil tank. The oil cooler is divided into 2
sections – the forward 1/3 is for transmission oil and the aft 2/3 is for engine oil. Metal contamination of
the cooler will require the cooler to be replaced.
ENGINE OIL COOLER BYPASS VALVE
The oil cooler bypass valve allows oil to either enter or bypass the oil cooler, and it will remain open until
the oil temperature reaches 71ºC. At this temperature the valve will begin to close and to allow oil to
enter the cooler. At approximately 81ºC, the valve will be fully closed, and oil will flow through the cooler.
OIL FILTER ASSEMBLY
The engine oil filter has a 10 micron, ultrasonically cleanable, metal element.
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