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DISCLAIMER
THIS INFORMATION CONTAINED HEREIN WAS PREPARED AND INTENDED,
ONLY, AS A MARKETING AID. IT CANNOT BE USED AS A SUBSTITUTE FOR THE
ROTORCRAFT FLIGHT MANUAL, MAINTENANCE MANUAL, ILLUSTRATED PARTS
BREAKDOWN OR OTHER PUBLICATIONS APPROVED BY AIR REGULATORY
AUTHORITIES FOR THE OPERATION, AND/OR MAINTENANCE OF ANY BELL
AIRCRAFT.
PUBLISHERS NOTICE
The data presented herein is general in nature and is intended for the use of BHTI
Sales Personnel and International Sales Representatives as an aid in describing
the features and benefits of Bell Helicopters to prospective customers.
This data includes proprietary material of BHTI. Disclosure, reproduction, or use
of any of the material herein by persons other than BHTI employees, and BHTI
Independent Representatives [International Dealers] is forbidden without written
permission from BHTI.
The PHOTOGRAPHS and DRAWINGS appearing herein are for ILLUSTRATION
only, and may differ from the actual aircraft, as the BHTI continuing product
improvement process causes some components, specifications, and procedures to
be changed or replaced.
THE WEIGHTS, DIMENSIONS, AND PERFORMANCE DATA SHOWN MAY NOT
BE CURRENT. For the most current tabular data, and performance and fuel flow
charts refer to the current year’s Product Specification Booklets.
©Copyright, Bell Helicopter Textron, Inc., 2003
PRODUCT DATA
®™, Bell Helicopter, Bell Helicopter Textron Inc., Bell
206B3, 206L-4, 407, 427, 430, 412, 609, JetRanger and
LongRanger are registered trademarks of Bell
Helicopter Textron Inc.
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CONTENTS
SECTION ONE
GENERAL DESCRIPTION OF
THE STANDARD CONFIGURATION
SECTION TWO
OPTIONAL EQUIPMENT
BHTC FACTORY INSTALLED
& VENDOR STC KITS
SECTION THREE
DIMENSIONS,
TABULAR PERFORMANCE
& KIT WEIGHTS
INDEX
ALPHABETICAL LIST OF
FEATURES & BENEFITS
PRODUCT DATA
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GLOSSARY
INTERNATIONALLY APPROVED ABBREVIATIONS
Atmosphere
C
ISA
ISA+20 C
ISA+30 C
SL
Degrees Celsius (° is not required if meaning is clear.)
International standard atmosphere. (Standard temperature is defined
as 15°C at sea level, with a lapse rate of -1 °C for each 150 meters (500
feet) increase in pressure altitude.)
International standard temperatures increased at each altitude by a
constant 20 C or 30 C.
Sea level.
Rate of Climb
R/C
ft/min
m/s
Rate of climb.
Feet per minute.
Meters per second. (One meter per second is 3.28 ft/s )
Speed and Distance
h
km
km/h
M
mph
Hour/hours.
Kilometer/kilometers.
Kilometers per hour. (One kilometer per hour equals 0.54
knots or 0.62 statute miles per hour.)
Knots. (Nautical miles per hour.)
Meter/meters. (One meter equals 3.28 feet.)
Statute miles per hour. (One mile per hour equals 0.87
knots. )
nm
Nautical miles.
sm
Statute miles.
kg
Kilogram/kilograms.
lb
Pound/pounds. (One pound equals 0.45 kilograms.)
Weights
Performance
Definitions
IGE
OGE
max
Maximum Endurance
Speed
Long Range Cruise
Speed
Pressure Altitude
PRODUCT DATA
In ground effect. (Hovering near the ground creates a
cushion of air that increases the weight carrying capability
of the helicopter (the ground effect). This effect generally
is present up to one rotor diameter above the ground.)
Out of ground effect. (Since the ground effect extends up to
one rotor diameter out of ground effect refers to flight that is
higher than this height above the ground.)
Maximum.
The speed at which power required to fly is a minimum.
The higher of the two speeds at which 99% of maximum
range occurs. Cruising at this speed is a tradeoff of one
percent in maximum range for 20 to 30 kilometers per hour
(10 to 15 knots), higher speed.
The altitude indicated on an altimeter or barometer when
the sea level standard pressure reading is the reference
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SECTION ONE
GENERAL DESCRIPTION OF
THE STANDARD CONFIGURATION
PRODUCT DATA
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PRODUCT DATA
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SYSTEMS OVERVIEW
INTRODUCTION
The Bell 407, a light SEVEN place single turbine engine powered helicopter, is the
latest entrant in the classic 206 series -- the most popular helicopter in history -that has achieved over 30 years of sales.This latest model in the series insures its
continuing popularity by blending the attractive lines of the 206 family with the well
established rotor system from the US Army's OH-58D armed scout helicopter.
The Bell 407 rotor is a soft-in-plane flex beam type hub with four individually
interchangeable blades. The blades and the hub are all composite. The pitch change
bearings and lead / lag dampers are elastomeric elements that require no scheduled
maintenance and have benign failure modes. The hub plate and the pitch horns are
made of aluminum forgings for light weight and strength.
PRODUCT DATA
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The engine in the Bell 407 is the Rolls Royce 250-C47B. This is a new version of the
well proven 250-C30 engine series from Rolls Royce Engine Company. The Rolls
Royce 250 engine series has accumulated over 100 million hours of flight time,
with deliveries totaling over 30,000 engines. Controlling the engine is a Full Authority
Digital Electronic Control (FADEC). Through the use of electronics, the engine is
able to maintain rotor speed with closer tolerances, provide cool starts, lower direct
operating costs, and automatic start capability.
The passenger cabin is wider than current versions of the Bell LongRanger to give
each of the three rear seats more width, making the accommodations more comfortable
and the ride more pleasant. Some things have not changed: Bell Helicopter's
commitment to safety, low cost of operation, and reliable, responsive helicopters
that are supportable and economic through their lifetimes.
PRODUCT DATA
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BELL HELICOPTER CANADA
The Bell 407 is assembled at Bell Helicopter Canada, a division of Textron Canada
Ltd. The manufacturing facilities are situated in Mirabel, Quebec, 30 kilometers (19
miles) north of Montreal. The plant is a short drive from Dorval Airport for United
States and Canadian arrivals, and only minutes from Mirabel International Airport.
MANUFACTURING AUTHORITY
The Bell 407 received its Canadian Type Approval Certificate [H-92, Issue 9] on
February 9, 1996. The Bell 407 was also certificated by the FAA in accordance with
provisions of 14 CFR Section 21.29 on February 23, 1996. Initial deliveries of the
new Bell 407 began in February 1996.
BASIS OF CERTIFICATION
The Bell 407 is certified under FARs Part 27 and 36, Appendix J. Additionally, it is
approved under Canadian Airworthiness Manual Chapters 516 (ICAO Ch. 11) and
527, Sections 1093 (b) (1) (ii) and (iii), 1301-1, 1557 (c) (3), 1581 (e) and 1583 (h).
INTERNATIONAL CERTIFICATION
A new Bell 407 is delivered with a Canadian Certificate of Airworthiness for Export ,
which may be readily exchanged for a Certificate of Airworthiness in the destination
country.
PRODUCT DATA
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EXTERIOR PAINT SCHEMES
The overall eye appeal of the Bell 407 is aided by the sparkling Polyurethane paint.
Not only is it durable and easy to keep clean, but it maintains and, in fact, exemplifies
the high standard of Bell quality and dependability. The standard Bell 407 comes in
three basic color scheme designs illustrated below, each in three colors. The colors
are selected from the many shown in the IMRON ® 6000 paint chip chart available
from the Bell representative.
A
B
C
Special company colors, paint design, logos,
etc., can be negotiated.
PRODUCT DATA
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GENERAL DESCRIPTION
OF THE STANDARD CONFIGURATION
SERIAL NUMBER 53000 AND SUBSEQUENT
The Bell 407 is the newest of the Bell light helicopters and is designed to carry
seven adults comfortably, including the pilot.
407 SEATING
The Bell 407 is approved for single pilot operation (standard seven place seating)
and is certified for land operation under day or night VFR non-icing conditions. The
pilot must occupy the right front seat. A passenger seat is provided on the left,
which may also carry a copilot with the addition of the optional dual flight control kit
[provisions for this kit are standard]. Five additional passengers can be
accommodated in the spacious main cabin. They can enter from either or both sides.
Two seats, centrally located, face aft. The remaining three seats are in a row, facing
forward, on a bench at the rear of the compartment.
PRODUCT DATA
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PILOT AND FORWARD PASSENGER SEATING
Two individual ergonomically designed seats with adjustable lumbar support, each
equipped with seat belt, double strap [four point] shoulder harness and inertia reel,
with "Quick Release" mechanism, are located in the cockpit. The color and upholstery
material for the seats, and interior trim of the cockpit match that which is selected
for the cabin.
OCCUPANT RESTRAINT
Each crew and passenger seat is equipped
with a restraint assembly that consists of
an inertia reel, a shoulder harness [pilot
& left front double straps / cabin single
strap], and an adjustable seat belt. The
inertia reel is provided with an antirebound lock feature and is capable of
retracting 56 centimeters (22 inches) of
web belt.
DOCUMENT STOWAGE
Provisions for stowage of documents /
equipment during flight are included. The
following list of stowage requirements are
provided for: Operators Manual, Maps,
Approach Plates / Airport Directories,
Flight Computer / Portable GPS.
Additionally, provisions for the aircraft log
book are available.
PRODUCT DATA
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STANDARD SEATING & INTERIOR TRIM
STANDARD SEATING & INTERIOR TRIM - The standard cabin seating consists of
five ergonomically designed fabric covered seats with individual seat belts, with
"Quick Release" buckles and single strap shoulder harness, arranged with three
equal width forward facing seats across the rear of the cabin and two individual
rearward facing seats aft of the cockpit. Available with Red, Blue, Tan, or Grey
upholstery fabric with Black seat belts. All vinyl or vinyl surround / fabric insert
upholstery is available as an option. The standard interior trim consists of color
coordinated, rugged Kydex ® covering for the cabin sidewalls and doors, with
lower door panel colored accent, vinyl covered arm rests, and Kydex ® headliner.
The floor is covered in nylon blend loop carpet.
The standard seating and interior trim weighs 154.9 lbs. (70.3 Kg.) and is included
in the basic aircraft weight.
STOWABLE HEADREST
A new initiative with the Bell 407 is a
redesign of the two aft-facing passenger
seats in conjunction with the wider cabin
and the larger capacity forward fuel tanks,
as compared to the Bell 206L-4
LongRanger IV. The notable features
of these seats is the head rest that stows
near the top of the seat back when there
is no passenger seated in them to enhance
rear-row passenger field of view, and the
seats are not canted.
PRODUCT DATA
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CORPORATE SEATING & INTERIOR TRIM
CORPORATE TRIM DOOR DETAILS
PRODUCT DATA
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CORPORATE SEATING & INTERIOR TRIM
The corporate cabin seating consists of five ergonomically designed fabric covered
seats with individual seat belts and single strap shoulder harness, arranged with
two extra wide forward facing outboard seats and a middle seat for occasional use
across the rear of the cabin (with a fold down arm rest between the outboard seats)
and two individual rearward facing seats aft of the cockpit. Available with Pewter,
Mushroom, Dark Blue, Saddle, Burgundy, Smoke Grey, or Forest Green upholstery
fabric with color coordinated seat belts in the passenger cabin (Crew seat belts are
Black). All leather or leather surround/fabric insert upholstery is an available option.
The corporate interior trim consists of color coordinated Kydex ® covering
for the cabin sidewalls and doors, with lower door panel fabric / leather accent, leather
covered arm rests, and leather covered Kydex ® headliner. The floor is covered
in plush deep pile wool carpet.
Selection of the corporate interior trim also specifies installation of Corporate
Soundproofing, which provides a reduced interior noise level. The corporate seating,
interior trim, and corporate soundproofing increase the basic aircraft empty weight
18.5 lbs. (8.4 kg.).
ALL SEATING CONFIGURATIONS FEATURE
EXTRA WIDE 3-ACROSS SHOULDER & HIP
ROOM FOR THE AFT ROW OF SEATS
PRODUCT DATA
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CORPORATE SEATING AND INTERIOR TRIM
CABIN
LIGHTS
HAND
GRIPS
FOLD DOWN
ARM REST
BEVERAGE
HOLDERS
OUTBOARD
ARM RESTS
PRODUCT DATA
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SOUNDPROOFING
The soundproofing blanket consists of two layers of fiberglass with a lead vinyl filler
and a quilted ''duon'' cover. It is sealed securely in position. The blanket is covered
with a plastic headliner. The parcel shelf is enclosed on all sides by a soundproofing
blanket. In addition, fiberglass batting is used in the area of the aft bulkhead. The
standard soundproofing installation weighs 5.8 lbs. (2.6 Kg.) and is included in the
basic aircraft weight.
CARGO
In addition to its primary role as a passenger transport, the 407 can be quickly converted to carry
bulky cargo internally. When the five passenger seats and seat backs are removed, the cabin
offers 2.4 cubic meters (85 ft 3) of cargo space. With the forward passenger seat removed this
can be increased to 3.0 cubic meters (105 ft3). It is simple to load cargo from one or both sides
because of the low sills, the extra wide doors and the relatively flat benches and floor. Both the
floor and seat benches are capable of withstanding a loading of 370 kilograms per square
meter (75 Ib/ft2). The large doors are approximately 91 centimeters wide and 99 centimeters
high (36 x 39 inches). The panel door increases the width of the opening on the left side to 155
centimeters (61 inches).
PRODUCT DATA
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DOORS
The Model 407 helicopter has FIVE entrance/exit doors for crew and passengers.
A crew and passenger door is installed on the left and right sidebody panels. The
right sidebody panel has hinge points for the crew door and passenger door. The
left sidebody panel has hinge points for the crew door and litter door. The left passenger
door is hinged on the litter door. The sidebody panels are molded carbon epoxy
structures that close off the cabin area and complete the aerodynamic structure of
the helicopter. The crew, passenger and litter doors are all interchangeable (between
407s), molded carbon-epoxy doors hinged to open out and forward. The caution/
warning and advisory panel, annunciators warn the pilot if the litter and/or passenger
door is not closed.
INTERIOR HANDLE
EXTERIOR
HANDLE
The two forward doors give direct entry to the crew compartment; the pilot right,
forward passenger left. The two aft doors each give access to the whole passenger
compartment. The four large doors each have a slam closing feature and a dualpoint latching system. The exterior handles are recessed so that they are flush
with the door surface. They are long and are hinged at the top so that they are easy
to operate. The interior handles are set into the window trim. A pullout motion opens
the door. All of the FIVE cabin doors may be removed if mission requirements so
dictate [the litter door is removable only if the left cabin door is removed first].
Any of the four large doors may be removed individually or in pairs. Any flight limits
imposed are detailed in the flight manual.
PRODUCT DATA
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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 constructed of carbon fiber with fiber glass skin and
equipped with a latch assembly, which may be operated from either side of the door.
Each crew door incorporates a sliding window and a stationary window.
PASSENGER DOORS
The passenger doors are installed on each side of the fuselage to give access to the
cabin area. Each door has a latch assembly that you can operate from the interior or
exterior side of the door. The latch assembly has interior and exterior door handles
and a lock that is installed near the exterior door handle.
CREW DOOR [PILOT'S SIDE]
PASSENGER CABIN DOOR
[ OPTIONAL SLIDING WINDOW INSTALLED]
DOOR SILLS
The low door sills, 50 centimeters high (20 inches) from ground level [standard
low skid gear], make entry or exit comfortable without additional steps. All door
sills are trimmed in metal.
OPTIONAL SLIDING WINDOW
Each cabin door has a panoramic stationary window (sliding windows are a NO
COST option).
BATTERY AND BAGGAGE DOORS
TWO additional doors are provided. A door in the nose compartment allows access
to the aircraft battery (and hourmeter), while a door on the left aft fuselage permits
access to the separate 0.45 cubic meter (16 cubic foot) baggage compartment.
PRODUCT DATA
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LITTER DOOR
Between the doors on the right side is a panel approximately 63 centimeters (25
inches) wide, with a large, single-pane window. Between the large doors on the left
side is the fifth, narrower door. It is hinged forward and carries the hinges for the
large aft door. It also has a large single pane window. When the aft door is open and
the narrower panel door is unlatched, the two open together exposing the whole left
side of the passenger compartment. The exterior handle for the narrow door is also
flush mounted and is equipped with a keyed lock. It is installed vertically to the right
of the panel window and turns 64° to the left to open. The narrow panel door cannot
be opened from the inside since it has no interior handle.
LITTER DOOR
HANDLE
&
KEYED
LOCK
[CLOSED] LITTER DOOR HANDLE [OPEN]
LITTER DOOR INSTRUCTIONS
LITTER DOOR CLOSED
PRODUCT DATA
LITTER DOOR OPEN
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CREW, LITTER, AND PASSENGER DOOR SEALS
The Bell 407 cabin doors are effectively sealed to prevent the entrance of water
either in flight or on the ground, and external gutters are not required. The seals for
the crew doors, the litter door and the passenger doors are made of “V” shaped
silicone rubber. They are installed on the left and right fairing assemblies. The seal
for the forward edge of the left passenger door is installed on the aft edge of the
litter door. The seals are attached to the fairing assemblies, nose and litter door
with double-sided adhesive tape. When you open the doors, the “V” shaped seal
supplies a path for water that leads away from the inside of the cabin.
CABIN AND COCKPIT DOOR SEALS
BATTERY ACCESS DOOR
The battery access door is installed on
the nose of the helicopter. It gives access
to the battery, hourmeter and one circuit
breaker. The access door is hinged and
has three 1/4 turn turn-lock fasteners
attached to the forward edge of the door,
and two additional fasteners at the mid
points on either side of the door. Plastic
coated stainless steel restraint cables
limit upward travel of the door to prevent
contact with the windshield. The battery
access door has a plastic honeycomb
core with an external fiberglass skin.
PRODUCT DATA
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BAGGAGE DOOR
The baggage compartment door is aft of the passenger door on the left side of the
helicopter. The door is hinged and is held closed by two push-button latches and
a keyed lock. The door has a composite honeycomb core with an external fiberglass
skin. Weather stripping along the edges seals the door. The caution/warning panel
baggage door annunciator warns the pilot if the door is not closed.
WARNING LIGHT
MICROSWITCH
PUSH BUTTON
LATCHES
[ WEB STRAP IS PART OF
THE OPTIONAL LEFT HAND
FUSELAGE MOUNTED
MAINTENANCE STEP KIT]
WEATHER
STRIP SEAL
MICROSWITCH
CONTACTOR
KEYED
LOCK
CARGO LOADING INFORMATION
DOOR LOCKS
All five cockpit and cabin doors and the baggage door can be locked from the outside
with the same key. The key must be used for each. Once locked, the door cannot be
opened from the inside. Two keys are provided.
DOOR OPENERS / PROTECTORS
Pneumatic strut door openers/supports are available for all five cabin doors and the
baggage door as an optional kit , and help protect the doors when operations are
conducted in high or gusty wind conditions.
PRODUCT DATA
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WINDOWS
All the windows and chin bubbles are made of gray-tinted transparent acrylic.
VISIBILITY is excellent. The two large windshields are separated by a narrow center
post assembly. The door window, a pair of large chin bubbles and a pair of skylights,
complete the ''all around'' visibility. The two skylights are of acrylic and
have a heavier tint. The tint provides comfortable shielding from direct sunlight without
interfering with the excellent visibility from all seats.
SKYLIGHTS
WINDSHIELDS
LOWER WINDOWS
[CHIN BUBBLES]
WINDSHIELDS
The two piece windshields are made of light grey tinted acrylic plastic. They
supported by the windshield center post, cabin roof, side body fairings and the nose
structure. The windshields are held in position by formed aluminum alloy retainers.
The retainers on the windshield center post and nose structure are installed with
screws and nutplates. For water tightness, sealant is applied to the mating surfaces
of the windshields and retainers.
PRODUCT DATA
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WINDOWS
The Bell 407 windows are made of light grey tinted acrylic plastic. The crew
doors contain a three component window with the lower forward section as a sliding
window for crew ventilation. The litter door window, cabin panel window and passenger
door windows are single piece fixed windows. Sliding ventilation windows similar
to those in the crew doors are available as a no-cost option for the passenger doors.
The crew doors, passenger doors, litter door, and cabin panel windows are flush
mounted and sealed for water tightness. Screws and self-adhesive tape are used
to attach the flush mounted windows to their frames.
CABIN DOOR WITH
OPTIONAL SLIDING WINDOW
STANDARD CABIN
WINDOW [NON-VENTED]
LITTER PANEL WINDOW
PILOTS DOOR WINDOW
PRODUCT DATA
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LOWER WINDOWS
The two lower windows are made of light grey tinted acrylic plastic. They
are installed in the lower cabin nose section and supported by the nose skin, floor
panel and console structure. The lower windows are beaded and are held in position
with a rubber seal for water tightness. Drain holes with plastic plugs are provided in
both chin bubbles to aid in keeping the bubbles clean.
VENT INLET DRAINS
LOWER WINDOWS
[CHIN BUBBLES]
DRAIN HOLES
SKYLIGHT
The two skylights are made of dark grey
tinted acrylic plastic. The skylights
are flush mounted on the forward cabin
roof and part of the sidebody fairings. The
skylights are installed with double-sided
self-adhesive tape and four screws. The
skylights are sealed for water tightness.
SKYLIGHT
PRODUCT DATA
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THE PILOT'S STATION
The Pilot's Station is on the right. The pilot enters the crew compartment by the
right front door. The entire pilot's work space is engineered for efficiency and comfort.
It is roomy. The seat is well designed. Visibility through the large windshields, chin
bubble and door windows is excellent. In spite of the expansive curved surfaces,
distortion is minimal. The flight controls are placed for optimum convenience.
FLIGHT CONTROLS
Flight controls are conventional. They consist of a cyclic and a collective flight
control stick and directional (antitorque) pedals. The flight control system is a
mechanical linkage, using push-pull control tubes and bell cranks [NO CABLES
are used]. When adjustments are required, they are simple and easily made. All
self-aligning and rod-end bearings are the spherical type and require no lubrication.
To the pilot, good flying qualities mean responsive control, good stability, ease of
handling and excellent maneuverability. The Bell 407 has these qualities.
CYCLIC & COLLECTIVE
HYDRAULIC SERVOS
OPTIONAL COPILOT
"QUICK DISCONNECT"
COLLECTIVE STICK
PILOT'S CYCLIC STICK
OPTIONAL COPILOT
"QUICK DISCONNECT"
CYCLIC STICK
OPTIONAL COPILOT
TAIL ROTOR CONTROL PEDALS
PRODUCT DATA
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TAIL ROTOR
PITCH CHANGE
MECHANISM
TAIL ROTOR
HYDRAULIC SERVO
PILOT'S COLLECTIVE STICK
& TAIL ROTOR PEDALS
The flight controls are routed beneath the pilot’s and passenger seats aft to the
center of the helicopter then up to the cabin roof through the control column that
also serves as a primary cabin structure. Access panels on the forward and aft side
of the control column and the belly of the helicopter and seat panels are provided for
inspection of control components and maintenance accessibility. Aluminum alloy
control tubes are used throughout the collective, cyclic, and tail rotor controls. Some
control tubes are fixed in length with bonded end fittings, while others have adjustable
fittings that are readily replaceable.
Dual controls may be installed as an option on any Bell 407 helicopter to provide
dual flight control capability either for operations requiring a pilot and a copilot or
for pilot training operations. The copilot cyclic and collective controls may be removed
by use of "Quick Disconnect" fittings when not required. The dual controls may be
selected as a factory option or added later, as the provisions for this optional kit
are part of the standard configuration.
PRODUCT DATA
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CYCLIC CONTROLS
The cyclic controls consists of a control stick, torque tube, hydraulic servo actuators,
control tubes, and bellcranks. Movement of the control stick is transmitted through
linkage and hydraulic servo actuators to the swashplate, which actuates the rotating
controls to the main rotor. Servo actuators are incorporated to minimize the effort
required to move the controls and to reduce main rotor feedback forces. The cyclic
stick extends upward and forward from the front of the pilot’s seat. The torque tube
connects to the cyclic stick support. The cyclic fore and aft movement is fed through
a cam assembly that automatically adds an amount of lateral cyclic input that is a
percentage of the fore and aft cyclic movement. A spring canister is provided in
line with the cam input to permit cyclic movement in the event that the cam assembly
becomes jammed. A stick centering indication light is attached to the cyclic stick
bellcrank. When the aircraft is on the ground this switch will cause the cyclic centering
light to display on the caution/warning panel when the stick is not centered. A balance
spring is used in the system to minimize the cyclic stick mass imbalance forces in
the longitudinal and lateral control system.
CYCLIC CONTROL STICK
PILOTS' CYCLIC WITH OPTIONAL
EQUIPMENT PROVISIONS
The pilot's cyclic grip has provisions for radio/ICS and other equipment control
switches. A cyclic friction adjustment wheel is located at the base of the pilot's
control stick. For helicopters with dual controls, the copilot’s cyclic stick is installed
in front of the copilot seat with all VFR control functions. A quick disconnect feature
permits rapid removal of the stick (when the electrical connector is disconnected,
an adapter must be installed to provide circuit continuity). Also, a spring pin assembly
is provided to ensure positive engagement of the stick.
PRODUCT DATA
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COLLECTIVE PITCH CONTROLS
The collective pitch controls consist of a collective stick, jackshaft, control tubes,
bellcranks, and a hydraulic servo actuator. Movement of the collective stick is
transmitted by means of linkage and the servo actuator to the swashplate collective
lever. Collective pitch control is transmitted to the main rotor controls by vertical
movement of the swashplate. The servo actuator is mounted on a support, which is
located on the cabin roof directly forward of the transmission, along with two servo
actuators for the cyclic system. The collective stick is installed to the left of the
pilot’s seat, and it extends upward and forward through a flexible cover. The collective
jackshaft provides a mounting point for the collective stick. The pilot's collective
stick mounts the engine "twist-grip" throttle, the engine start switch, and landing
light switch.
COLLECTIVE CONTROL STICK
COLLECTIVE WITH TWIST THROTTLE AND
OPTIONAL EQUIPMENT PROVISIONS
MANUAL THROTTLE PERCENT SCALE
There is adequate space for other equipment control switches. An adjustable friction
bearing mounted on the jackshaft allows the pilot to adjust the friction to his own
requirements. A minimum friction adjustment clamp, which is located at the left end
of the jackshaft, ensures that the collective stick will always have a preset minimum
friction. The collective trunnion and lever is installed between the collective jackshaft
and control tube. It ties the collective controls to the mixing lever of the cyclic controls.
When the collective stick is moved to change rotor pitch, the cyclic control servo
actuators and linkage will move to maintain the swashplate in its relative plane. For
helicopters with dual controls, the copilot collective stick is installed at the left of
the copilot’s seat with a fully functioning throttle control. A quick disconnect feature
permits rapid removal of the copilot collective stick.
PRODUCT DATA
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DIRECTIONAL [TAIL ROTOR] FLIGHT CONTROL SYSTEM
The directional control system is a conventional push pull type system. The pilot
control input from the pedals is routed, through the center pedestal side panels, to
a bellcrank located beneath the center pedestal. The input is routed aft to the control
column tunnel. At the bottom of the control column tunnel a bellcrank routes the
input to a bellcrank at the top of the control column just below the cabin roof. The
input is then routed inside the main beam assembly to a walking beam located in the
hat rack area. The walking beam then lowers the horizontal push pull motion to a
hydraulic servo located in the intermediate section of the helicopter. The boosted
input is routed to a walking beam and then into the tail boom to a bellcrank under the
tail rotor gearbox. The control input changes direction and is routed to the pitch
change mechanism located on the right hand side of the tail rotor gear box.
TAIL ROTOR
GEARBOX
HYD. SYSTEM
PRESSURE &
RETURN LINES
TAIL
ROTOR
SERVO
T / R PITCH CHANGE
BELLCRANKS
TAIL ROTOR SERVO LOCATION
TAIL ROTOR SERVO DETAIL
PRODUCT DATA
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AIRSPEED ACTUATED PEDAL STOP [AAPS]
The directional control system incorporates an Airspeed Actuated Pedal Stop which
acts to automatically restrict pilot pedal input during cruise flight , while allowing
maximum pedal authority for low airspeed and hovering flight.
EMERGENCY RELEASE
AREA A
DETAIL OF AREA A
CAM & SOLENOID
PRODUCT DATA
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The Airspeed Actuated Pedal Stop electrical components include a Pedal Restrictor
Control Unit (PRCU), an actuating rotary solenoid, a position sensing microswitch
and a status/press-to-test switch / annunciator. System power is provided from a 5
Amp overhead panel circuit breaker. The aircraft pitot and static system interfaces
with the PRCU. The PRCU calculates airspeed from the pitot and static inputs. When
the airspeed is greater than 55 ±5 KIAS, the PRCU solenoid extends the pedal stop
restrictor into the left pedals range of travel. The PRCU will trigger the solenoid to
retract the pedal stop when calculated airspeed falls below 50± 5 KIAS. Upon extension
of the pedal stop, the position sensing microswitch is activated and the green ENGAGED
message on the lower panel annunciator illuminates. This message is extinguished
when the pedal stop is retracted. If the system malfunctions the MASTER W/C
annunciator will illuminate, a single "bong" tone will sound and the Caution Panel
will display a PEDAL STOP caution message. A red emergency pedal stop release
handle located under the pilots collective will disengage the pedal stop to restore
full pedal authority.
PEDAL STOP
WARNING
LIGHT
SYSTEM TEST
&
ENGAGED LIGHT
PRODUCT DATA
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TAIL ROTOR PEDALS
The tail rotor control pedals mounted on the pilot’s compartment deck include a
pedal adjuster, which provides manual adjustment, in and out, of up to 4.5 inches
(11.4 cm) pedal position according to the pilot’s needs. For helicopters with dual
controls, the copilot’s fully functional tail rotor control pedal assembly is installed
on the floor in front of the copilot’s seat to provide a means for the copilot to control
the tail rotor assembly. The control pedals are linked to the pilot’s by means of control
tubes and a bellcrank. They also include a pedal adjuster.
PEDALS
ADJUSTER
HEEL PLATES
OPTIONAL COPILOT PEDALS
PILOTS' PEDALS
PEDAL DISTANCE ADJUSTMENT WHEELS
PRODUCT DATA
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HYDRAULIC SYSTEM DESCRIPTION
The main rotor cyclic, collective and yaw flight controls are assisted by single stage
servo actuators utilizing a single hydraulic system. Normal operating system
delivers a maximum volume flow of 2.85 gpm at operating pressure of 1000±25 Psi
at rotor Rpm of 100% (NR ).
SYSTEM OPERATION
Gravity supplies fluid from the hydraulic reservoir to the pump. Pressurized fluid
from the pump flows through a quick disconnect to the pressure filter. The pump is
also connected to the return side of the system to provide cooling to the pump. The
pressure filter will collect contaminants contained in the fluid. A differential pressure
indicator alerts in the event of a possible blockage of the filter element but does not
bypass the filter. The solenoid valve then routes the fluid to the servos, pressurizing
the system or routes fluid to the return side of the hydraulic system, depressurizing
the system. Activation of the hydraulic system is accomplished electrically. Energizing
the solenoid valve will deactivate the hydraulic system. De-energizing the solenoid
will activate the hydraulic system. This creates a fail safe mode in the event of
electrical failure. Hydraulics activate automatically during total loss of electrical
power. The hydraulic pressure solenoid switch is located on the upper control panel
in the cockpit. Pressurized fluid flows to the four flight control actuators assisting
the pilot. Return fluid flows through the pressure filter to collect contaminants contained
in the fluid. A differential pressure indicator alerts in the event of possible filter
element blockage and a bypass relief valve relieves in the event of filter blockage.
Fluid flows through a quick disconnect to return to the reservoir. A safety relief
valve between the pressure and return sides of the system protects against excess
pressure conditions.
CYCLIC [OUTBOARD]
AND COLLECTIVE
[CENTER] SERVOS
FLUID
RESERVOIR
HYDRAULIC
PUMP
TO TAIL
ROTOR
SERVO
"QUICK DISCONNECT"
SERVICE FITTINGS
ACTUATOR SUPPORT
PRODUCT DATA
SOLENOID
& RELIEF VALVE
ASSEMBLY
FILTER &
"BYPASS"
INDICATOR
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MAIN ROTOR SERVOS
TAIL ROTOR SERVO
HYDRAULIC SYSTEM SCHEMATIC
PRODUCT DATA
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HYDRAULIC SYSTEM COMPONENTS
[ FOR COMPONENT LOCATIONS REFER TO THE PRECEDING ILLUSTRATION AND SCHEMATIC DIAGRAM S]
HYDRAULIC RESERVOIR
The hydraulic reservoir and cover are constructed from magnesium alloy. The
reservoir is mounted on a brace and support forward of the main transmission and
above the hydraulic pump assembly. Fluid is cooled in the reservoir by means of
atmospheric radiation.
HYDRAULIC PUMP
The hydraulic system uses a main transmission mounted variable delivery, selflubricated, constant pressure pump. The transmission output is protected by the
shaft assembly. The shaft assembly has a shear strength of 150 - 225 inch pounds.
Pump discharge pressure is rated at 1000±25 Psi (6985±170 kPa). The pump will
run at 900 Psi (6287 kPa) full flow, 2.85 gpm (10.8 liter per minute). The pump
contains four marked ports: inlet, outlet, case, and seepage drain.
FILTER ASSEMBLY
Two hydraulic filters are installed on a bracket near the hydraulic reservoir ,on the
left side of the cabin roof. One filter is installed in the pressure line and one in the
return line. Both filters have a red mechanical clogged condition indicator to alert
possible blockage of the filter element. If the filter element in the return line becomes
clogged, the bypass valve will open and allow fluid to bypass the filter and return to
the reservoir. The hydraulic pump has an internal pressure compensator which
prevents application of excessive pressure to the system and to the filter in the
pressure line. Therefore, no bypass valve is required in the pressure line filter This
indicator rises when the differential pressure across the filter is 70 ± 10 psi (482 ±
69 kPa). The indicator will not work at hydraulic fluid temperatures less than about
35°F (2°C). If the hydraulic fluid is cold, the viscosity will change and a high differential
pressure across the filter will be normal. The clogged indicator will not work at low
fluid temperatures to stop false clogged indications. If the hydraulic fluid temperature
is more than 35°F (2°C) the indicator will give the correct condition of the filter, even
if the ambient temperature is below 35°F (2°C).The return line filter is a bypass type
with the bypass valve inside the return line filter body. When the differential pressure
across the filter is more than 100 ± 25 psi (690 ± 172 kPa), the bypass valve opens.
Pressurized hydraulic oil is then sent directly to the reservoir. When the differential
pressure falls to 60 psi (450 kPa) the bypass valve will close. The hydraulic oil will
now pass through the filter on the way to the reservoir.
HYDRAULIC SYSTEM LIGHT
The hydraulic system light [caution warning annunciator] will be illuminated when
the hydraulic pressure switch detects a decreasing pressure of less than 650 psi
(45.7 kPa), and will extinguish on an increasing pressure at 750 psi (52.7 kPa). The
hydraulic pressure switch is mounted on the hydraulic manifold, forward of the hydraulic
actuator support.
SERVO ACTUATOR SUPPORT
The cyclic and collective servo actuator support is installed on the cabin roof
forward of the transmission. It serves as a mount for the hydraulic servos and associated
bellcranks. The collective servo actuator is mounted in the center position and the
two cyclic servo actuators are mounted on the outboard positions.
PRODUCT DATA
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SOLENOID VALVE
The solenoid valve is incorporated in the hydraulic system for turning the system
ON and OFF. The solenoid valve is installed on the left hand side of the upper deck.
The solenoid valve is normally open, and when the HYDRAULIC SYSTEM circuit
breaker is in and HYDRAULIC SYSTEM switch is OFF, electrical power is applied to
the solenoid which closes the valve and removes hydraulic pressure from the four
servo actuators. In case of a total electrical failure, the system is fail-safe ON.
RELIEF VALVE
The relief valve is incorporated in the hydraulic system for controlling system pressure.
The relief valve is installed on the left side of the upper deck. The relief valve is
normally closed, but if system pressure should exceed the normal pressure range
of 1000 Psi (6985 kPa) and rise to 1075 to 1375 Psi (7412 to 9480 kPa), the relief
valve will open to protect the system from damage.
RESERVOIR
HYDRAULIC
CIRCUIT BREAKER
QUICK DISCONNECT
HYDRAULIC PUMP
HYDRAULIC SWITCH
HYDRAULIC SYSTEM CONTROLS
QUICK DISCONNECT FITTINGS
The "Quick-disconnect" fittings in the pressure and return lines of the hydraulic pump
allow the attachment of a ground service unit [hydraulic mule] to perform system
checks and maintenance without starting the helicopter's engine. When the fittings
are disconnected, each half of the connector self-seals to keep the system free from
contamination and to prevent leakage of fluid. Ground hydraulic operation instructions
are found in the aircraft maintenance manual.
PRODUCT DATA
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HYDRAULIC SERVO ACTUATOR DESCRIPTION
The flight control servo actuator consists of a servo head assembly mounted on a
cylinder assembly. The servo head assembly contains the servo valve, connected
to an input arm assembly which is pivoted on a rod end assembly. Check valve,
sequence and thermal relief functions are incorporated in the servo head assembly.
The piston is connected to the airframe at a fixed point which is pivoted on a rod end.
Input from the controls is assisted by movement of the servo head assembly and not
by the piston.
HYDRAULIC SERVO SCHEMATIC DIAGRAM
CYCLIC AND COLLECTIVE SERVOS
& SERVO ACTUATOR SUPPORT
PRODUCT DATA
TAIL ROTOR SERVO
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COCKPIT INSTRUMENTS
Cockpit instrumentation is the latest technology and incorporates state-of-the-art
individual liquid crystal displays (LCD) which provide greatly improved reliability
compared to conventional electromechanical instruments. Three of the LCD
instruments incorporate event and exceedance recording capability which can be
used to monitor over-limit power usage. Monitoring aids in determining the proper
level of maintenance action. The displays also include caution/warning as well as
the instrument functions.
The engine instruments selected by Bell Helicopter for the Model 407 are derivatives
of Litton Special Devices solid state LCD production that is trademarked as EASI™.
The 407 engine instruments are "next generation" design that combines advanced
LCD and microprocessor technologies to provide the Model 407 with the most reliable,
accurate and day and night readable engine instruments available in the aerospace
industry. The Litton engine instruments introduce the following leading edge
technologies:
First Minima Twisted Nematic (FMTN) Display Technology - The main advantage
of this display medium is that it offers ten times greater contrast over a wider viewing
angle than the early LCD's. The greater contrast enhances brightness and readability
from either cockpit seating positions. FMTN displays are based on the same technology
that is utilized in active matrix panels.
Dual Backlight System - A dual day/night backlight system is an integral part of
the Litton engine instruments for the Model 407. The dual backlight system has full
dimming capability which will provide optimum overall performance in all lighting
conditions. The dual system emits variable low level lighting for night operations
and high level illumination for daytime operations, the latter of which produces excellent
readability in situations of direct sunlight and forward glare.
Totally Solid State Design - There are no moving parts to break down or wear out.
The entire instrument is built around advanced microprocessor technology which
offers superior reliability compared to conventional electromechanical instruments
for low DOC and less downtime.
PRODUCT DATA
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In fact, Litton has demonstrated, in accordance with MIL-STD-781 and Bell Helicopter
specifications, a 79,800 hour MTBF! This demonstrated MTBF is over 10 times
greater than conventional gauges. Additionally, there is no scheduled maintenance
required for the Model 407 engine instruments.
DATA DOWNLOAD CONNECTORS
LCD TORQUE / WITH TEST SWITCH
Performance Monitoring System - The engine instrument package in the Model
407 includes a complete engine performance monitoring system.
Each engine instrument in the Model 407 contains a unique software program that
meets the particular operating requirements of that instrument. Configured in the
software program is a self-test function as well as data acquisition, computational
and communication routines. Three of the indicators [Torque, MGT, and NG] are
also programmed to monitor for predetermined events and, when these preset levels
are exceeded, the event magnitude and duration will be stored in the indicator's
nonvolatile memory. All of the engine instruments on the Model 407 are connected
to a serial bus in the panel harness. Exceedence events can be downloaded using
Litton supplied software and an IBM compatible Laptop computer. To operate the
system, the computer, via its serial port, is connected to the serial bus in the panel
harness which allows for single point downloading of exceedence information. The
maintenance system software programmed in the computer contains a communications
routine that polls each indicator for events recorded in the nonvolatile memory. The
events, including magnitude and duration, are then downloaded into the personal
computer for display, storage and analysis.
Warning for the pilot of impending exceedance events is also a benefit of the Model
407 Engine Performance Monitoring System. As an engine instrument detects an
impending operating exceedance, the CHECK INSTRUMENT caution annunciator
flashes and, a flashing warning of the trend arc and digital display occurs on that
instrument's dial facing.
PRODUCT DATA
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LCD INSTRUMENTS
FUEL PRESS.
& AMPS
MEASURED
GAS TEMP.
FUEL QTY.
N G (RPM)
GAS PROD.
ENGINE
PRESS.
& TEMP.
XMSN. TEMP.
& PRESS.
VOLT METER /
CLOCK / OAT
PRODUCT DATA
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INSTRUMENT PANEL
The instrument system is divided into four separate categories: flight, navigation,
propulsion, and miscellaneous. All indicators are installed in the hinged instrument
panel except the standby compass, and hourmeter. The standby compass is mounted
in a support attached to the right side of the cabin structure slightly forward of the
instrument panel. The hourmeter is mounted in the nose compartment.
PEDAL
STOP
123
123
123
123
123
123
123
123
123
123
LEGEND FOR ABOVE ILLUSTRATION
[BASIC INSTRUMENT PANEL GAUGES & SWITCH LOCATIONS]
CAUTI ON / WARNING / ADVISORY PANEL
FWD
FUE L
OAT / VOLTS
CLOCK
TORQUE
FUEL PRESS/
AMPS
MEASURED GAS
TEMPERATURE
FUEL
QUANTITY
RPM GAS
PRODUCER
C/W TE ST
LCD
I NS TRUME NT
RPM HORN
CHECK
CHE CK
MUT E
AIRSPEED
INCLINOMETER
[PROV FOR ATT. I ND. ]
ALTIMETER
DUAL TACH
POWER TURBINE
& ROTOR RMP
PROVISIONS FOR
DIR. GYRO. OR HSI
VERTI CAL
S PEED IND.
PE DA L STOP
PROV. FOR
TURN IND.
QTY TEST
FUEL
FADEC
GA UGE
HORN
TES T
TEST
X MSN OIL
PRESS/TEMP
PRODUCT DATA
E NG OIL
PRESS/TEMP
P ROV. FOR
ADF INDICATOR
42
PROVISIONS FOR
COURSE DEV. IND.
PTT
WARN
FA DE C
FUEL
MODE
VALVE
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The flight instrument system includes the pitot/static system and the following
instruments and panels:
Airspeed indicator, Altimeter, Vertical Speed, Inclinometer, and Airspeed
Limitation Panel.
The navigation instrument system consists of the standby magnetic compass;
but depending on the kit installation, may include others.
The propulsion instrument system includes the following:
Dual NR and NP tachometer indicator, Engine torque indicator, Measured Gas
Temperature indicator, NG gas producer tachometer indicator, Engine
Oil pressure and temperature indicator, Transmission Oil pressure and
temperature indicator, Fuel quantity indicator, Fuel pressure and
Ampmeter.
The miscellaneous instrument system includes the following:
Clock/Outside Air Temperature/Voltmeter, Engine Hourmeter.
• SOME INSTRUMENTS & EQUIPMENT SHOWN BELOW ARE OPTIONAL KITS •
PRODUCT DATA
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PITOT/STATIC SYSTEM
The pitot tube, which is mounted on the STATIC PORT
most forward part of the cabin nose
structure just right of the helicopter
centerline, supplies impact air to the
airspeed indicator. Static air pressure for
instrument operation is obtained from two
static ports on the left and right panels
aft of the cabin lower window. Both the
static ports and the pitot tube are
electrically heated, controlled by a switch
in the overhead control panel. The pitot
static system supplies input to the Pedal
Restrictor Control Unit [PRCU] of the
Airspeed Actuated Pedal Stop.
ALTIMETER
VERTICAL
SPEED
PITOT TUBE
AIRSPEED
INDICATOR
PEDAL RESTRICTOR
CONTROL UNIT
[PRCU]
STATIC PORTS
SYSTEM
DRAINS
PITOT TUBE
PRODUCT DATA
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FLIGHT INSTRUMENTS
AIRSPEED INDICATOR
The airspeed indicator is a standard pitot/
static instrument. This indicator provides
airspeed reading in KNOTS [Nautical
Miles per Hour] by measuring the
difference between the impact air pressure
from the pitot tube and the static air
pressure from the static vents.
ALTIMETER
The altimeter provides a direct reading
of helicopter height in feet above sea level.
It is connected to the static air system to
sense atmospheric pressure. An external
adjustment knob is provided to
compensate for variations of prevailing
barometric pressure [setable in millibars
or inches of mercury] .
VERTICAL SPEED INDICATOR
[RATE OF CLIMB]
The VSI is an instrument sensitive to the
rate of change in atmospheric pressure.
The scale gives an indication of positive
or negative vertical speed of from zero
to four thousand feet per minute. It is
connected to the static air system
TURN AND SLIP INDICATOR
[INCLINOMETER]- NOT ILLUSTRATED
The inclinometer is a simple instrument
consisting of a curved glass tube, ball,
and dampening fluid. The ball indicates
when the helicopter is in directional
balance, either in a turn or straight and
level flight. If the helicopter is in a yaw or
slip, the ball will move off center.
PRODUCT DATA
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NAVIGATION INSTRUMENT
STANDBY MAGNETIC COMPASS
The magnetic compass is a standard,
nonstabilized, magnetic type instrument
mounted on a support which is attached
to the right side of the forward cabin. The
compass is used in conjunction with the
compass correction card that is located
below it.
PROPULSION INSTRUMENTS
DUAL TACHOMETER (NP/NR)
The dual tachometer, which indicates in
percent, furnishes both rotor and power
turbine rpm information. It indications are
derived form signals generated by
monopole magnetic pickups on the engine
and transmission. Normal operation is
indicated when both the rotor rpm and
power turbine rpm indicator needles are
synchronized and in the green arc. In a
departure from tradition dating to piston
engines, the NR is shown on the outer
scale.
ENGINE TORQUEMETER
The engine torquemeter, which is
calibrated in percent, receives its signal
from a transducer attached to a wet line
on the forward left side of the accessory
gearbox. It provides direct pressure
readings from the torquemeter sensor to
the gage.
PRODUCT DATA
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MEASURED GAS TEMPERATURE
(MGT) INDICATOR
The Measured gas temperature indicator
receives temperature indications from four
bayonet type thermocouples mounted
between the NG and NP turbines. It is
graduated in degrees Celsius, and
electrical power is supplied from the 28
Vdc bus through the MGT circuit breaker.
GAS PRODUCER TACHOMETER (NG)
The gas producer tachometer, which
indicates in percent, provides gas
producer rpm information. The inner scale
reads in 1% increments and the outer in
2%. This instrument receives it's signal
from the gas producer monopole speed
pickup.
ENGINE OIL TEMPERATURE/
PRESSURE INDICATOR
The engine oil temperature/pressure
indicator is a dual instrument. The
pressure side of the instrument is
precalibrated in pounds per square inch
against a standard, and receives its signal
from a transducer attached to the output
from the engine oil pump. Temperature
indication in degrees centigrade is
provided from a thermobulb in the engine
oil reservoir.
TRANSMISSION OIL TEMPERATURE/
PRESSURE INDICATOR
The transmission oil temperature/
pressure indicator is a dual instrument.
The pressure side of the instrument is
precalibrated in pounds per square inch
against a standard and receives it’s signal
from a transducer attached to the output
from the transmission oil pump.
Temperature indication in degrees
centigrade is provided from a thermo bulb
in the transmission oil manifold.
PRODUCT DATA
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PROPULSION INSTRUMENTS (continued)
FUEL QUANTITY INDICATOR
The fuel quantity indicator provides
readings of the fuel supply in pounds. It
indicates total fuel quantity from the three
quantity probes located in the fuel cells.
The capacitance type system includes a
fuel signal conditioner, and is self
calibrating for varying fuel temperatures.
FUEL PRESSURE INDICATOR
and DC AMPMETER
The fuel pressure indicator and DC
ampmeter is a dual instrument that
provides both ampmeter and aircraft fuel
pump pressure indications. Electrical
circuitry is provided from the fuel pressure
transmitter to the indicator.
MISCELLANEOUS INSTRUMENTS
CLOCK / O.A.T. / VOLTMETER
The clock/outside air temperature/
voltmeter is a multifunction indicator
mounted in the upper left area of the
instrument panel. The OAT temp sensor
is located on the aircraft belly, just forward
of the front cross tube. A button located
on the center top of the instrument
changes the upper display between OAT
and Volts. The instrument defaults to the
voltmeter function. The lower half of the
display shows Flight time, Elapsed time,
Local time, or Universal time (G.M.T.). The
flight time function is controlled by the
landing gear microswitch and is activated
when the aircraft landing gear leaves the
ground (microswitch open).
PRODUCT DATA
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MISCELLANEOUS INSTRUMENTS (continued)
HOURMETER
The engine hourmeter is mounted in the
nose compartment on the aft bulkhead.
For the hourmeter to record time, the
hourmeter circuit breaker must be in, the
engine NG must be greater than 55% and
the aircraft landing gear must be off the
ground (landing gear microswitch open).
INSTRUMENT OPERATION FEATURES
All of the propulsion instruments have a built in test (BIT) capability.
POWER-ON BIT
The power-on BIT starts when the power is applied to the instrument. The power-on
BIT does an integrity check of the electronic components of the indicator.
On all indicators except the dual tachometer, during the power-on BIT, the trend
arc display shows the full scale for 6 to 8 seconds. Torque and NG digits display
8188.8. MGT and Fuel digits display 81888. NR and NP needles of the dual tachometer
move to 107% and 100% (upper redline limit) respectively.
If an LCD indicator fails the BIT, the trend arc and the digits, if applicable, will show
blank. If faults are detected in the dual tachometer during the BIT, pointer movement
will not occur. Any failure during the BIT found on torque, MGT and NG indicators
will be recorded in nonvolatile memory (NVM) of the instrument. Instruments that
successfully complete the BIT will have the first LCD of the indicator lit or will display
the appropriate indication. On the torque, MGT and NG indicators, a letter E will
appear in the digital display when the instrument has recorded an exceedance.
Pressing the INSTR CHK button will display the value of the last exceedance recorded.
Releasing the button will clear the E until the next power up of the instrument.
LCD TEST
BUTTON
PRODUCT DATA
INSTRUMENT
CHECK BUTTON
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COMMANDED BIT
The commanded BIT starts when the LCD TEST button is pushed. The commanded
BIT turns on all the LCDs so that the pilot can verify that all the LCDs are working.
On the dual tachometer indicator the individual Rotor and Turbine pointers are driven
to indicate their respective upper redline limits.
BIT CHECK DISPLAY
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LCD EXCEEDANCE MONITORING FEATURES
TORQUE, MGT and N G instruments have an ability to record exceedances.
Exceedances have been predefined at the design stage and preprogrammed into
the microprocessor in each indicator. Exceedances are defined as limits of operation
above which there may be some maintenance action required. Exceedance monitoring
is provided as an aid to determining required maintenance action. Only the person
in command of the aircraft at the time the exceedance occurs can verify that the
exceedance recorded reflects the actual occurrence. The above instruments have
built in NVM(non volatile memory) that allows them to store up to 50 exceedance
events. For each exceedance the memory records the date, duration and peak
value during the exceedance. If exceedance events recorded are greater than 50,
the earliest dated exceedance that was in memory is deleted and the latest exceedance
is added in its place. To provide advance notice to the pilot that an exceedance is
about to be recorded, these indicators also have other preprogrammed advisory
points at which they will flash the LCD display. Specific values are discussed in
applicable systems descriptions in the Bell 407 Manufacturers Data. When an
advisory is displayed by the instrument, the instrument will also turn on the CHECK
INSTR light on the caution panel. If the pilot makes control inputs to reduce the
instrument readings below the advisory values, the advisory will no longer be displayed
and the CHECK INSTR light will be turned off and no exceedance will be recorded.
When no pilot corrective action is taken and the indicator exceeds specifically
preprogrammed values an exceedance will be recorded. When an exceedance is
about to be recorded, the CHECK INSTR light will be turned on by the instrument. In
addition, when an exceedance is recorded the letter E will be displayed at the left
digit on the digital display. The pilot can acknowledge the exceedance and cause
the peak value to be displayed on the analog and digital display by pushing the INSTR
CHK button on the instrument panel. The exceedance will display for a maximum of
11 seconds; less if the pilot pushes on the INSTR CHK button for a shorter period of
time. If there have been exceedances recorded on different indicators each indicator
will display its last exceedance. Once the pilot has pushed the INSTR CHK button
and the exceedance(s) has been displayed, the E will disappear from the digital
display. The E will not display until the power is removed from and applied to the
indicator (such as by pulling its circuit breaker or turning aircraft power off and on).
The last exceedance will continue to display each time the indicator is powered up
until the exceedance(s) are removed from the NVM of the indicator using a computer
with maintenance download software.
LCD CHECK INSTRUMENT (CHECK INSTR) CIRCUIT
If an exceedance has not been detected by the TORQUE, MGT or NG indicator, the
CHECK INSTR light and the “E” on the indicators digital display will not illuminate
when the INSTR CHK button is pressed. The CHECK INSTR light circuit is designed
to alert the pilot that either the TORQUE, MGT, or NG indicator is about to, or has
detected an exceedance. When the indicators exceed preset values the indicator’s
display will begin to flash and the CHECK INSTR light will turn on. If an E is displayed
by an indicator the CHECK INSTR light will remain on until the pilot acknowledges
the exceedance by pushing the INSTR CHK button.
PRODUCT DATA
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CAUTION AND WARNING SYSTEM
The caution and warning system includes the caution light panel, engine rpm sensor
and warning horn, low rotor rpm sensor and warning horn, FADEC warning horn,
associated components and interconnecting wiring.
CAUTION LIGHT PANEL
The caution light panel is a
multisegmented panel energized by the
28 Vdc bus and protected by the 5 amp
CAUTION LIGHTS circuit breaker. Each
segmented indicator light is in series with
its respective caution system.
ENGINE OUT-[RED]
The engine out warning system provides both visual and audible indications of an
engine out condition. It includes the ENG OUT warning light located on the caution
panel, the engine out warning alarm located on the overhead console, and the engine
rpm sensor located on the equipment shelf forward of the instrument panel.
The engine rpm sensor is connected to the gas producer tachometer circuit; and
when the gas producer rpm drops below 55 ± 3%, the engine out warning light will
illuminate and the engine out warning alarm will sound.
LOW ROTOR RPM-[RED]
The low rotor rpm caution system provides both visual and audible indications of a
low rotor rpm condition. It includes the ROTOR LOW RPM caution light located on
the instrument panel, the rotor rpm warning alarm installed on right side plastic
headliner, and the rotor rpm sensor located on the equipment shelf forward of the
instrument panel. The ROTOR rpm sensor is connected to the rotor tachometer circuit;
and when rotor rpm drops below 95 % the rotor rpm caution light will illuminate and
the rotor rpm warning alarm will sound.
HIGH ROTOR RPM-[RED]
The high rotor rpm caution system provides a visual warning of a high rotor rpm. The
Rotor RPM sensor is connected to the tachometer circuit; when rotor rpm exceeds
107%, the rotor rpm caution light will illuminate.
PRODUCT DATA
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RPM WARNING HORN MUTING SYSTEM
The rpm warning horn muting system disables the warning horns when a pushbutton
switch mounted on the instrument panel is activated. The rpm warning horns are
automatically reset when operational limits are reached.
CAUTION PANEL SEGMENTS
CAUTION/WARNING/ADVISORY ANNUNCIATORS LEGEND
POSITION ANNUNCIATOR COLOR POSITION
ANNUNCIATOR
1
FLOAT TEST
GREEN
19
FADEC FAIL
2
SPARE
AMBER
20
FADEC DEGRADED
3
ENGINE ANTI-ICE
WHITE
21
FADEC MANUAL
4
FLOAT ARM
AMBER
22
ENGINE CHIP
5
AUTO RELIGHT
WHITE
23
XMSN CHIP
6
START
WHITE
24
T/R CHIP
7
BAGGAGE DOOR
AMBER
25
GEN FAIL
8
LITTER DOOR
AMBER
26
XMSN OIL PRESS
9
HEATER OVERTEMP AMBER
27
CHECK INSTR
10
L/FUEL BOOST
AMBER
28
BATTERY RELAY
11
L/FUEL XFR
AMBER
29
XMSN OIL TEMP
12
FUEL FILTER
AMBER
30
HYDRAULIC SYSTEM
13
R/FUEL BOOST
AMBER
31
BATTERY HOT
14
R/FUEL XFR
AMBER
32
ENGINE OVSPD
15
FUEL VALVE
AMBER
33
CYCLIC CENTERING
16
FADEC FAULT
AMBER
34
ENGINE OUT
17
RESTART FAULT
AMBER
35
PEDAL STOP
18
FUEL LOW
AMBER
36
RPM
COLOR
RED
AMBER
AMBER
AMBER
AMBER
AMBER
AMBER
RED
AMBER
AMBER
RED
AMBER
RED
RED
AMBER
RED
AMBER
RED
SHADED SEGMENTS INDICATE OPTIONAL EQUIPMENT PROVISIONS
PRODUCT DATA
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BATTERY HOT-[RED]
The battery hot warning system includes the BATTERY HOT warning light and internal
battery overtemp sensor switches. Both switches are activated when the battery
temperature of 140 to 150 ° F (60 to 65.5 ° C). When the warning light illuminates,
the battery switch must be turned off to allow the battery temperature to decrease.
ENGINE OVERSPEED-[RED]
The engine overspeed warning system provides a visual indication of an engine
overspeed condition. The OVSPD warning light is red and is located on the caution
panel. The light is activated when either NG or NP exceeds normal operating limits.
FADEC FAIL -[RED]
The FADEC FAIL caution light system utilizes a red caution light to warn the pilot of
a failure of the FADEC during start or during flight.
XMSN OIL PRESS -[RED]
The transmission oil pressure caution system includes the TRANS OIL PRESS caution
light and an oil pressure switch. The pressure switch, which is installed in the
transmission oil pressure line to the indicator, serves to activate a TRANS OIL PRESS
caution light on the instrument panel to advise of low oil pressure. The light will
illuminate at a decreasing pressure of 32 ± 2 psig, and it will extinguish at a maximum
increasing pressure of 36 psig.
XMSN OIL TEMP-[RED]
The transmission oil temperature switch, which is located next to the temperature
bulb on the transmission oil filter assembly, will close when the temperature of the
oil rises above 110 ± 5 C and the caution light will illuminate.
AUTO RELIGHT-[WHITE]
Engine igniter is operating. This light will illuminate when the FADEC is in MANUAL
mode. The light will also be on during start.
PEDAL STOP- [AMBER]
When illuminated indicated the failure of the Pedal Restrictor Control Unit of the
Airspeed Activated Pedal Stop.
BAGGAGE DOOR-[AMBER]
The baggage door caution system includes the BAGGAGE DOOR OPEN caution
light, micro switch and related wiring. If the baggage door is not securely fastened,
the caution light will illuminate.
BATTERY RELAY-[AMBER]
The battery relay caution system includes the BATTERY RLY caution light and related
circuitry through the battery relay and the battery switch. If the battery relay remains
energized when the battery switch is placed in the OFF position, the caution light
will illuminate. The battery is still connected to the DC BUSS.
CHECK INSTRUMENT-[AMBER]
Torque, MGT, or NG is about to or has detected an exceedance. Flashing of the
LCD trend arc and digital display of the effected instrument indicates an impending
exceedance. The letter E in the digital display indicates an exceedance has occurred.
PRODUCT DATA
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CAUTION PANEL SEGMENTS (continued)
CYCLIC CENTERING-[AMBER]
A proximity sensor below the pilots cyclic is energized when the cyclic is NOT centered.
When there is weight on the landing gear (landing gear microswitch closed) this
illuminates the light [regardless of operating RPM].
ENGINE ANTI-ICE-[AMBER]
The anti-ice switch is on. The engine is receiving anti-icing bleed air, which is required
to be ON in visible moisture when ambient temperature is at or below 5°C (40°F).
ENGINE CHIP DETECTOR-[AMBER]
The engine chip detector caution system includes the ENG CHIP caution light, two
engine furnished magnetic drain plug/chip detectors, and related wiring. If ferrous
metal particles are found in the oil, the magnet will attract them; and when there are
sufficient metal particles to complete the circuit, the ENG CHIP caution light will
illuminate.
FADEC DEGRADED-[AMBER]
The FADEC degrade light is an amber light that receives its signal from the Electronic
Control Unit (ECU) and indicates a degraded condition either during flight or during
engine shutdown.
FADEC FAULT-[AMBER]
The Fadec fault light is an amber light that receives its signal from the Electronic
Control Unit. It indicates that MGT, TRQ, or NG automatic limiting circuit(s) not
functional [FADECs/w Ver. 5.1] .
FADEC MANUAL-[AMBER]
The FADEC manual light is an amber light that receives its signal from the Electronic
Control Unit when no automatic governing is available. The Auto Relight caution
light will also illuminate.
FUEL BOOST PUMPS-[AMBER]
The fuel pump caution system includes the R/FUEL PUMP caution light, the L/FUEL
PUMP caution light, the right and left side fuel pressure switches, and related wiring.
The left or right fuel pump caution lights are controlled by the left and right fuel
pressure switches respectively, which are located in the base of each pump and
activated by fuel pressure. The lights will illuminate when that fuel flow is reduced.
FUEL FILTER (AIRFRAME)-[AMBER]
The fuel filter caution system includes the FUEL FILTER caution light, fuel filter
pressure switch, and related wiring. Should a fuel pressure differential of approximately
1 psi develop at the airframe filter, the filter pressure switch will close and energizes
the FUEL FILTER caution light, thereby alerting the pilot that the fuel filter is beginning
to clog.
FUEL LOW-[AMBER]
The fuel low system, which is independent of the fuel quantity system, will illuminate
the FUEL LOW caution light when approximately 110 ± 15 pounds (52.2 ± 6.8 Kg.)
of fuel remain in the aft tank.
PRODUCT DATA
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FUEL TRANSFER PUMPS-[AMBER]
The fuel transfer caution system includes the R/FUEL TRANSFER caution light, the
L/FUEL TRANSFER caution light, the right and left side pressure switches, and the
related wiring. The left or right transfer caution lights are controlled by the left and
right pressure switches, which are located in the base of each pump and are activated
by fuel pressure. The lights will illuminate when fuel flow is reduced.
FUEL VALVE-[AMBER]
The fuel shutoff valve, located above the aft fuel cell on the right side of the helicopter,
is an electrically operated, motor driven valve and provides a means of shutting off
fuel to the engine. The amber light is activated when the fuel valve position differs
from the Fuel Valve switch indication.
GENERATOR FAILURE-[AMBER]
The generator failure caution system includes the GEN FAIL caution light, line control
relay, starter relay, and related wiring. When the generator is not on line, the GEN
FAIL caution light will illuminate.
HYDRAULIC SYSTEM-[AMBER]
Hydraulic pressure is below operating limits.
LITTER DOOR-[AMBER]
The litter door caution system includes the LITTER DOOR OPEN caution light, microswitches for the upper and lower door strikers, and related wiring. If the litter door is
not securely fastened, the caution light will illuminate.
RESTART FAULT-[WHITE]
The Restart Fault light is a white light which detects a fault in the ECU and will not
allow a start in the auto mode.
START-[WHITE]
The start light is a white advisory light and will illuminate when the start relay is in
start mode.
TRANSMISSION CHIP DETECTOR -[AMBER]
The transmission chip detector caution system includes the TRANS CHIP caution
light, two transmission chip detectors, a freewheeling unit chip detector, and related
wiring. If ferrous metal particles are found in the oil, the magnet will attract them;
when there are sufficient particles to complete the circuit, the XMSN CHIP caution
light will come on.
TAIL ROTOR GEARBOX CHIP DETECTOR-[AMBER]
The tail rotor gearbox chip detector caution system includes the T/R CHIP caution
light, tail rotor gearbox chip detector, and related wiring.
PRODUCT DATA
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OVERHEAD CONTROL GROUP
The overhead console groups most of the electrical switches and controls in one
convenient easy-to-reach panel. It contains: 1. Switches for interior and exterior
lights, defog blowers, battery and generator. 2. Breakers for all the standard electrical
circuits (They are all the trip-free type which cannot be overridden). 3.Space for
circuit breakers and switches for optional kits.
There is an airspeed limitation placard in knots mounted between the console
and the windshield. Just behind the overhead console, on the forward face of the
vertical control tunnel, are two headset outlets and the loud speaker for the pulsating
engine out warning and FADEC horn. The speaker for the steady low rotor RPM
warning is mounted on the ceiling.
WARNING HORNS
CIRCUIT BREAKERS & SWITCHES
MAIN SWITCH PANEL
HEADSET OUTLETS
PRODUCT DATA
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WARNING
HORNS
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RADIO PEDESTAL
The pedestal provides space for the installation of optional communications and
navigation radios. A case for the airworthiness and registration certificate is riveted
to the right forward side of the pedestal. The radio pedestal slopes down from the
instrument panel and extends horizontally to the seat bench. The standard 407 has
no radios, however, it can accept the standard width radio boxes in both the sloping
and horizontal surfaces [ 159 centimeters (6.25 inches) in width]. The design gives
the pilot the best view of all the radio controls.
DOCUMENT CASE
The airworthiness, registration
certificate, and radio license case i s
riveted to the right forward side of the
instrument panel console. These
documents are required to be carried in
the aircraft at all times.
***DRAWING FOR ILLUSTRATION ONLY***
DOCUMENT CASE
FIRE EXTINGUISHER
SLANTED AREA
HORIZONTAL
AREA
CABIN 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. A mounting bracket is the quick
RADIO PEDESTAL [SHOWN WITH OPTIONAL opening type for rapid removal of the
extinguisher.
EQUIPMENT]
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VENTILATION SYSTEM
Air for cabin ventilation is obtained by opening the sliding windows in each of the
cockpit doors. Additional ventilation to the cockpit may be obtained through the ram
air grills located on the nose of the helicopter by pulling out on the left and right
VENT control knobs. At lower airspeeds when the ram air flow is reduced or when
the windshield has condensation ventilation may be assisted by positioning the DEFOG
BLOWER switch on the overhead console to the ON position.
OPTIONAL CABIN SLIDING WINDOW
[AVAILABLE AS A NO-COST OPTION]
COCKPIT SLIDING WINDOW
[STANDARD EQUIPMENT]
RAM AIR SYSTEM
When the VENT control knob below the
instrument panel is pulled out, ambient
ram air will be forced into the air intake
grills and directed through the plenum and
turn vane. The turn vane incorporates a
flapper valve assembly which is open or
closed by use of the VENT control knob
through the control cable.
RAM AIR INLETS
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DEFOG SYSTEM
Two electrically driven axial flow blowers are installed in the inlet end of the defroster
nozzles. The blowers are controlled by DEFOG BLOWER circuit breaker type switch
on the forward end of the overhead console. The defog system is primarily used for
ventilation and defogging during ground operation of the helicopter. When the system
is used for ventilation or defogging, it is recommended that both VENT control knobs
be extended to the full open position.
VENT
CONTROL
KNOB
RAM AIR INLET
AXIAL
BLOWER
DEFOG AIR
OUTLET
WATER DRAIN
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AIRFRAME
The airframe consists of three major sections; the fuselage, the cowlings, and the
landing gear.
COWLINGS
LANDING GEAR
FUSELAGE
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CONSTRUCTION
The Bell 407 has extended contour side panels containing two crew doors, the two
left cabin passenger doors and the fixed panel and one passenger door on the right
side. The panels incorporate all occupant doors and provide a smooth faired surface
from the forward edge of the cockpit doors to the forward edge of the baggage door
at the rear to reduce aerodynamic drag. These panels bulge outward to a maximum
of three and one-half inches, 8.9 centimeters, on both sides, resulting in the three
aft seats having a width of approximately 19 inches, 48 centimeters, each.
The bulged side panels effectively provide more shoulder room in the critical areas
of the cockpit and the aft cabin seat while having minimal impact on the primary
structure.
The two left side doors can be opened simultaneously to provide a 61 inch opening,
155 centimeters to more easily load litter patients or bulky cargo.
MODEL 407 AIRFRAME
RIGHT SIDE
PANEL AND
DOORS
LEFT SIDE
INTERMEDIATE
PANEL
LEFT SIDE PANEL AND DOORS
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FUSELAGE
The fuselage has three primary structural assemblies: the forward section extends
from the cabin nose to the bulkhead aft of the passenger compartment, the intermediate
section extends from the bulkhead aft of the passenger compartment to the tailboom,
and the aft or tailboom section.
407
FUSELAGE
ASSEMBLY
FIXTURE
RIGHT SIDE
PANEL
AND DOORS
INTERMEDIATE
SECTION
LEFT SIDE PANEL
AND DOORS
FORWARD SECTION
The forward section is primarily made up of a aluminum honeycomb structure which
provides the major load carrying elements of the forward cabin. The forward section
provides for pilot and passenger seating, fuel cell enclosures, and pylon support.
The side body panels are made of carbon fiber plys and core with fiberglass skin.
Under the fiber glass skin is a layer .005" thick of copper mesh for bonding and
lighting protection. The entrance and exit doors are also made of carbon fiber plys.
PRODUCT DATA
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OCCUPANT PROTECTION
Two rugged bulkheads are built into the cabin, one behind the crew seat bench
and one behind the aft passenger bench. A hollow vertical control tunnel is from
floor to roof immediately behind the forward bulkhead. These all add structural
integrity to the cabin and provide rollover protection to the occupants. A pair of
longitudinal beams in the cabin roof provide strength for the support of the transmission.
The cabin structure extends forward, up through the nose section to the windshields,
forming a compartment for the landing lights and one for the battery. The external
power receptacle and pitot tube are mounted between the two compartments.
Ventilation grilles are low on the nose; one on each side of the battery compartment.
NOSE AREA
CABIN ROOF
BEAMS &
TRANSMISSION
MOUNTS
AFT ROLLOVER
BULKHEAD
VERTICAL
CONTROL
TUNNEL
FORWARD ROLLOVER
BULKHEAD
PRODUCT DATA
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FUSELAGE FEATURES
A recessed step is provided aft of the left passenger door for access to the roof.
The light weight corrosion free carbon fiber side body fairings add stiffness to the
fuselage, with the added benefits of increasing cabin internal volume, and ease of
field repair.
The battery compartment is sized to accept either the standard 17 ampere-hour
battery, or the optional 28 ampere-hour [heavy duty] battery. Plastic tracks, and an
extra wide door make installation and removal of the battery quick and convenient.
COROSION FREE CARBON FIBER SIDE BODY PANEL
The lower fuselage at the forward edge of the aft passenger seat is strengthened to
support increased cargo hook capacity of 2646 pounds, 1200 kilograms. The core
material of the lower fuselage shell assembly and the nose panel from the earlier
206 series has changed to a new Aluminium alloy for increased corrosion protection
when operating in a saltwater environment.
PRODUCT DATA
65
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INTERMEDIATE SECTION
The intermediate section uses an aluminum internal structure and two one piece
carbon fiber skin assemblies, which do not need stiffeners or stringers, to form an
semimonocoque construction. Also provided by this section is a deck for engine
installation, a spacious baggage compartment, and a compartment under the engine
for environmental, electrical, and tail rotor hydraulic equipment.
LEFT HAND COMPOSITE PANEL
RIGHT HAND COMPOSITE PANEL
IN ASSEMBLY FIXTURE
PRODUCT DATA
TAIL ROTOR SERVO LOCATED BENEATH
ENGINE DECK
LOWER FUSELAGE & AFT CROSS TUBE MOUNT
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FIREWALLS & ENGINE PAN
AFT FIREWALL
FWD. FIREWALL
A titanium engine pan, located below the
engine, acts as both a drip pan and firewall.
It is curved to provide sufficient clearance
to allow the removal of accessories without
having to remove the engine. The engine
pan along with the titanium fore and aft
firewalls are fitted within the top part of
this section to form the engine
compartment.
AFT
FIREWALL
FWD.
FIREWALL
ENGINE PAN
FORWARD FIRESEAL COVER
TAIL ROTOR DRIVESHAFT FIRESEAL
COVER
FORWARD FIRESEAL COVER
The forward cover makes a fireseal around the main driveshaft and the optional
rotor brake. The cover is made of fiberglass composite and is attached to the forward
side of the forward firewall. The cover is made up of two cover assemblies, which
allows easy access for inspection or maintenance.
TAIL ROTOR DRIVESHAFT FIRESEAL COVER
The tail rotor driveshaft fireseal cover makes a fireseal around the tail rotor driveshaft.
It is made of stainless steel. The cover is attached to the forward side of the aft
firewall and has a doubler installed on the right side.
PRODUCT DATA
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ENGINE COMPARTMENT INSTALLATION
INTERMEDIATE SECTION / TAIL BOOM
AIR DAM
EQUIPMENT COMPARTMENT
WITH TAIL ROTOR SERVO
EQUIPMENT COMPARTMENT
The equipment compartment is the area behind the parcel shelf and above the baggage
compartment. It contains the electrical relays and regulators and has space for the
optional heater or environmental control unit in a convenient
location near the
source of engine customer bleed air.
AIR DAM
At the junction of the tail boom and the intermediate section an AIR DAM is installed
to prevent the flow of high pressure air behind the fuselage into the equipment
compartment and the inside of the cabin.
PRODUCT DATA
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BAGGAGE COMPARTMENT
Located directly below the engine area is the baggage compartment which has a
volume capacity of 16 cubic feet and can carry a maximum of 250 pounds. The
cargo door, located on the left side of the fuselage and hinged at the forward end,
opens the full width and height of this compartment. It is secured by means of two
push-button latches and a keyed lock. Within the compartment are nine tie-down
loops to secure cargo and equipment. There is also provisions for the optional 76
Liter (20 U.S. Gallon) auxiliary fuel cell in the baggage compartment.
AUX. FUEL
TANK MOUNT
PROVISIONS
TIE DOWN
LOOPS
LEVELING
PLATE
AUX. FUEL
LINE PROV.
BAGGAGE COMPARTMENT WITH OPTIONAL
AUXILIARY FUEL TANK
PRODUCT DATA
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TAILBOOM SECTION
The tailboom is an aluminum alloy full monocoque structure, except for the forward
10 inches, where the loads are redistributed by means of four intercostal load carrying
members. The tailboom supports the tail rotor drive shafting, tail rotor gearbox, vertical
fin, and horizontal stabilizer. Covers are provided to protect and provide a fairing
for the tail rotor driveshaft and gearbox.
TAIL BOOM ASSEMBLY FIXTURE
VERTICAL FIN
GEARBOX COVER
DRIVESHAFT COVERS
[TWO SECTIONS]
TAIL SKID
HORIZONTAL STABILIZER
407 TAIL BOOM
END PLATES
[AUXILIARY FINS]
PRODUCT DATA
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LEADING EDGE SLAT
HORIZONTAL STABILIZER
The horizontal stabilizer (with end plates) is a one piece aluminum honeycomb
structure, and does not have a movable elevator. The incidence of the horizontal
stabilizer is fixed to avoid the necessity of additional control system components.
AUXILIARY
FINS
The stabilizer assembly is a fixed horizontal tail surface, set to give an aerodynamic
down force on the helicopter when it is in flight at normal cruise speed. The stabilizer
has slats installed on the leading edge which make nozzle-shaped passages between
the stabilizer and the slats. This design improves the airflow conditions at high angles
of attack and slow speeds. Auxiliary fins are installed on the outboard edges of the
stabilizer to help the helicopter stay stable in flight. Both finlets are angled 5° [leading
edge] outboard.
PRODUCT DATA
71
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VERTICAL FIN
ANTICOLLISION
LIGHT
The vertical fin is constructed of aluminum
honeycomb core with aluminum skins. The
leading and trailing edge caps are formed
aluminum alloy and bonded in place. The
vertical fin contains a top fairing for
mounting the anticollision light and a
rubber bumper and tail skid are installed
on the lower end.
The vertical fin is mounted to a one piece
cast fitting and set to 9° leading edge
outboard (right) in order to relieve the tail
rotor in forward flight and prevent flapping.
TAIL SKID
VERTICAL FIN
TAIL ROTOR DRIVESHAFT COVERS
The two section tail rotor driveshaft covers are fiberglass and honeycomb composite
structures. When removed, the sections convenient size permits easy storage in
the cabin during routine inspections and maintenance.
TWO SECTION DRIVESHAFT COVERS
PRODUCT DATA
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TAIL ROTOR GEARBOX FAIRING
The tail rotor gearbox fairing covers the tail rotor gearbox. It is attached to the tailboom
and to the vertical fin. The fairing is a two piece installation. The upper half has two
screens and a lubrication service access door. The screens supply air to the tail
rotor gearbox for cooling. You can examine the tail rotor gearbox oil level through
the aft screen. The lower half of the fairing has a screened door that gives access to
the chip detector/gearbox drain for the tail rotor. The lower half also contains the
white aft position light and a drain hole. The fairing is made of a fiberglass composite
construction.
OIL FILLER DOOR/SCREEN
TAIL LIGHT
CHIP / DRAIN DOOR
TAIL ROTOR GEARBOX
DRAIN AND CHIP PLUG ACCESS DOOR
PRODUCT DATA
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COWLINGS AND FAIRINGS
FORWARD COWLING
The cowlings and the fairings are divided
into the five sections that follow:
the forward fairing assembly,
the transmission cowling assembly,
the engine air inlet cowling assembly,
the engine cowling assembly,
the aft fairing assembly.
TRANSMISSION COWLING
FORWARD FAIRING
ENGINE AIR INLET COWLING
ENGINE COWLING
[THREE SECTIONS]
TRANSMISSION COWLING
AFT FAIRING
The forward and aft fairings are made from fiberglass and honeycomb while the
engine side cowling is made from graphite composite. The transmission, engine
inlet and the upper engine cowlings are made from aluminum alloys and fiberglass
and honeycomb materials. Where aluminum alloy is used, structural attachments
are made either by bonding or mechanical fasteners [there are no spot welds to
promote corrosion]. Quick release fasteners hold the cowlings and the fairings in
position. All fasteners provide visual indication of proper closure. The inspection
and the access doors have flush-type latches which permit inspection and servicing
without having to remove the cowlings and the fairings.
PRODUCT DATA
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FORWARD FAIRING ASSEMBLY
FORWARD FAIRING IN RAISED POSITION
The forward fairing is installed on the cabin
roof forward of the transmission. The
fairing is made of composite material. A
single point hinge fitting assembly is
installed at the forward end of the fairing.
Two rods are kept inside the fairing on
both sides to be used as a support for the
fairing when it is raised. Two toggle hook
latch fittings hold the forward fairing when
it is closed. A seal is installed around the
perimeter to keep out water.
TRANSMISSION
COWLING
ENGINE AIR
INLET COWLING
TRANSMISSION COWLING
ASSEMBLY
The transmission cowling is installed on
the cabin roof over the transmission and
the controls. A cutout is given so that you
can check the hydraulic reservoir oil level.
ENGINE AIR INLET COWLING
ASSEMBLY
The engine air inlet cowling is installed on the cabin roof over the aft half of the main
transmission. There are inlet ducts on each side of the air inlet cowling that direct
the flow of air to the inlet screen or to the particle separator and into the engine inlet.
There is a hinged access door on each side of the cowling to permit inspection
procedures to be done without having to remove the cowling. On the right hand side
there is a hole so you can see the transmission oil level.
ENGINE AIR INLET
PRODUCT DATA
INSPECTION DOOR
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ENGINE COWLING
The engine cowling, which includes the upper cowling and the (left and right ) side
panels. Is constructed of fiberglass and aluminum alloy; it contains two screen
vents forward of the cutout for the engine exhaust stack. The upper cowling is aluminum
alloy and the side panels are constructed of fiberglass and incorporate five screen
vents and three flush latches and four wing/dzus 1/4 turn fasteners. The right engine
side cowling includes a screen vent for the starter/ generator cooling duct. The side
cowling may be hinged open and supported by a folding brace.
BLEED AIR VENT
EXHAUST
HINGE SUPPORT
OIL COOLER
EXHAUST
ENGINE OIL FILL DOOR
AFT COWLING
The aft cowling is constructed of fiberglass
and incorporates two screen vents, oil
cooler and oil tank access doors, and a
cutout for the oil cooler blower exhaust.
ENGINE OIL
RESERVOIR LEVEL
OIL DRAIN
ENGINE COWLING
PRODUCT DATA
AFT COWLING
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LANDING GEAR
The Bell 407 skid landing gear is simple, light weight, and practically maintenance
free. It is designed to absorb energy and to preclude damage to the airframe in the
event of a hard landing. The gear is designed for energy attenuation of vertical
descent speeds up to 2.0 meters per second (6.55 feet per second) by elastic
deflection of the forward and aft cross tubes. Yielding of the crosstubes, beyond
normal operating limits, in combination with elastic deflection, further attenuates
energy up to 2.44 meters per second (8.02 feet per second). The entire skid landing
gear assembly may be removed for shipping.
LOW SKID LANDING GEAR
The standard configuration Model 407 is
supplied with low skid landing gear. High
skid landing gear, and emergency float
landing gear are available as optional kits.
Each skid tube is fitted with a forward end
step, a tow fitting, saddles with sockets
for crosstubes, seven replaceable skid
shoes along the bottom, a rear cap and
two eyebolt fittings to mount the ground
handling gear.
STANDARD LOW SKID GEAR
GROUND WHEEL EYEBOLTS
SUPPORT BEAM
CROSS
TUBES
ATTACH FITTINGS
& STRAPS
EYEBOLTS
NONSLIP COATING
SKID
TOW RING
SKID SHOES
The landing gear is made from 7075 aluminum alloy tubes and consists of two
skids which are attached to the ends of two arched crosstubes. The full length of the
inside diameter and the outside diameter between left hand and right hand butt line
18 of the crosstubes are shotpeened for improved material properties. The landing
gear is attached to the fuselage by four straps which bolt to fittings located on the
two main bulkheads.
PRODUCT DATA
77
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AFT CROSSTUBE SUPPORT BEAM
The aft crosstube has a support beam
that touches the mating supports in the
helicopter structure. The function of the
support beam on the aft crosstube is to
reduce the possibility of 4 per rev ground
resonance.
SUPPORT
BEAM
LANDING GEAR SWITCH
The landing gear microswitch is activated
when weight is on the landing gear. It
provides a signal to control operation of
the cyclic centering light, the flight time
function of the clock, and the hour meter.
LOW SKID FAIRINGS (OPTIONAL)
The optional airfoil type landing gear fairings go over the forward and aft crosstubes
of the standard low skid gear. They are made of white thermoplastic and have aluminum
alloy supports. Fairings are not available for the high skid gear or emergency flotation
gear.
SKID SHOES
Each skid is fitted with seven steel replaceable skid shoes to absorb the wear caused
by normal ground contact. Also, the top surface of each skid is covered with wear
resistant BLACK non-slip coating.
TAIL SKID
A spring steel tail skid and rubber bumper are attached to the lower portion of the
vertical fin. It is designed to absorb shock, and reduce excessive nose down pitching
during hard tail low landing. Also, a tuning weight is mounted on the tail skid.
VERTICAL
FIN
RUBBER
BUMPER
PRODUCT DATA
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GROUND HANDLING
To move the Bell 407 on the ground, a set
of manually operated hydraulic dual
wheels is provided as loose equipment.
A tow bar is NOT provided.
TOW RING
STANDARD GROUND WHEELS
Two eyebolts are provided on each skid, near the helicopters center of gravity, for
attachment of the dual ground handling wheels. The wheels are snapped and pinned
in place and the wheels are then extended or retracted manually using a built-in bar
which activates a small hydraulic mechanism. The wheels weigh a total of 40.8
kilograms (90 pounds) and must be removed for flight. A tow ring is provided near
the forward end of each skid for ground handling. The maximum approved weight
for towing with the standard ground wheels is 1973 Kilograms (4350 pounds). If
the ground wheels from the Bell 205/212/412 series are used the 407 may be ground
handled at it's normal maximum gross weight.
PRODUCT DATA
79
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AIRCRAFT IDENTIFICATION & REGISTRATION
The airworthiness, registration certificate, and radio license case is riveted to
the right forward side of the instrument panel console. These documents are required
to be carried in the aircraft at all times.
DOCUMENT
CASE
FUEL
NOZZLE
GROUND
FUEL FILLER
DATA PLATE
FUEL
FILLER
CAP
AIRCRAFT DATA PLATE
The aircraft data plate is attached to the
right aft intermediate section skin between
the fuel nozzle ground and the fuel filler
cap.
A
S
PRODUCT DATA
80
M
P
E
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MAIN ROTOR
The Bell 407 main rotor hub and blade assembly consists of four composite blades
mounted to a main rotor hub assembly. Each blade is attached to a grip assembly by
two fixed diameter bolts. The main rotor diameter for the Bell 407 is 35 feet (10.7
meters) and has a tapered chord equivalent to 10.75 inches (25.5 centimeters).
The main rotor blades have a built in twist of 13°. With a design RPM of 413, the tip
speed is a low 757 feet per second (231 meters per second) . Elastomeric bearings
provide feathering, help dampen out vibrations and provide lead-lag for the main
rotor blades. The hub yoke is fabricated with a new epoxy matrix which has
demonstrated the capability to increase yoke fatigue life.
AERODYNAMIC COVER
FRAHM
DAMPER
MAIN ROTOR HUB
The main rotor hub assembly is attached
to the mast with a mast nut, an upper cone,
center cone set and a lower cone. The
main rotor hub assembly is aligned on the
mast by a master spline that also indicates
where the red blade will be.
MAST NUT
[LIFTING EYE]
BLADE GRIP
COMPOSITE YOKE
PRODUCT DATA
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ELASTOMERIC BEARINGS
The main rotor blade grip connects to the inboard and outboard elastomeric shear
bearings. The elastomeric shear bearings permit the main rotor blade to turn [feather]
around its pitch axis. The inboard elastomeric shear bearings attaches to the
elastomeric damper. The elastomeric damper is attached to the upper and lower
hub plate assemblies. The damper accommodates lead-lag, while the bearings and
yoke permit blade cone, droop, and blade pitch change. The pitch horns attach to
the blade grips and supply an attach point for the pitch links. The pitch horns turns
[feathers] the blade on the elastomeric shear bearings. This changes the angle of
attack (pitch) of the main rotor blade.
BLADE GRIP
UP & DOWN STOPS
PITCH HORN
SHEAR [FEATHER]
BEARINGS
YOKE
ELASTOMERIC
BEARING
UP & DOWN STOP ASSEMBLY
The main rotor blade up and down stops are attached to the upper and lower hub
plate assemblies. The main rotor blade stops keep the yoke flex (cone and droop) in
operational limits.
PRODUCT DATA
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MAIN ROTOR BLADES
The main rotor blades are a composite all-bonded assembly consisting of three
structural members: a fiberglass spar, nomex honeycomb core, and a trailing
edge strip. Skins, stabilized by the honeycomb core, are a 2-ply fiberglass wrap
with fibers that run at a ±45 degree angle to the span of the main rotor blade. The
outboard 18 inches of each blade has a bonded stainless steel erosion shield on
the lower surface. Conductive paint protects the main rotor blade against lightning
strikes and static charges. Nickel plated stainless steel erosion strips cover the
leading edge for resistance to abrasion. The main rotor blades are dynamically
balanced and are interchangeable. An adjustable trim tab is installed on the main
rotor blade trailing edge at blade station 144.00. Each blade weighs approximately
53 LBS (28.3 Kg). Each main rotor blade is color coded according to position. The
red blade is in the position next to the master spline on the mast. An index hole in the
hub also identifies the position of the red blade. The rotor blade retention bolts, the
optional expandable bolts, the grips and the rotating control components all have
the applicable color coded markings.
1. BLADE BOLT RETENTION BUSHINGS
2. M/R BLADE BONDED ASSEMBLY
3. NICKEL PLATED 301 STAINLESS STEEL
EROSION STRIP
4. TRIM TAB
5. FRETTING PAD
6. 301 STAINLESS STEEL EROSION SHIELD
7. 2-PLY UPPER SKIN [CLASS B GLASS/EPOXY
FABRIC & E GLASS/EPOXY TAPE
PRODUCT DATA
8. UNIDIRECTIONAL S-2 GLASS/EPOXY
TRAILING EDGE STRIP
9. 2-PLY LOWER SKIN [CLASS B GLASS/EPOXY
FABRIC & E GLASS/EPOXY TAPE
10. 2 LB/CU. FT. NOMEX HONEYCOMB CORE
11. UNI-DIRECTIONAL S-2 GLASS/EPOXY
ROVING SPAR STRUCTURE [+/- 45 DEG S-2
GLASS/EPOXY OUTER AND INNER TORSION
WRAPS]
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OPTIONAL M/R BLADE FOLDING
DYNAMIC BALANCE WEIGHTS
A Main Rotor Blade folding kit is an
option [Vendor STC]. The kit permits two
of the rotor blades to be folded (one aft,
and one forward, with each folded blade
held in a support fixture). This kit requires
the addition of a M/R Blade Fold
Provisions Kit which replaces two fixed
diameter bolts with the expandable bolts.
STANDARD BOLT
EXPANDABLE BOLT
[OPTIONAL W/FOLDING]
WEIGHT POCKET
FRAHM DAMPER
The frahm damper helps attenuate
vibrations created by aerodynamic forces
acting on the main rotor and hub, before
the vibrations are able to enter the
transmission and airframe. It is installed
on top of the main rotor hub, under an
aerodynamic fairing.
PRODUCT DATA
FRAHM DAMPER [COVER REMOVED]
84
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TAIL ROTOR
The Bell 407 tail rotor and tail rotor drive system are the same as for the OH-58D.
This simplification will reduce maintenance cost. The diameter is 5.4 feet, 1.7 meters.
Blade chord is 6.35 inches, 16.1 centimeters. The airfoil section is the BHT 12083, a 12% thick version of the well proven Wortmann FX-083 design. With a design
RPM of 2500, the tip speed is 709 feet per second, 216 meters per second. The
tail rotor yoke is made of stainless steel to resist corrosion. There are four Teflon
impregnated bronze lined monoball type pitch change [feather] bearings for low
pedal forces. The trunnion/flapping bearing is constructed out of elastomeric
material, requiring no lubrication, and is mounted to the yoke at a 30 degree angle
to control blade flapping.
TAIL ROTOR BLADE
1. TRAILING EDGE STRIP
2. 301-STAINLESS
EROSION SHIELD
3. KEVLAR SKIN
4. BALANCE WEIGHT
5. BALANCE SCREW
6. BUSHING
7. 301-STAINLESS
ABRASION STRIP
TAIL ROTOR BLADE
The two tail rotor blades are constructed with Kevlar skins and a nomex honeycomb core
for corrosion avoidance. Inboard, there are stainless steel reinforcing plates bonded to the
blade root for strength at the attachment points. The full length of the blades leading edge has
a stainless steel abrasion strip, and the blade tip is covered with a nickel plated stainless
steel shield to protect the blade from erosion.
PRODUCT DATA
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NOTE: PHOTOS AND DRAWINGS SHOWN
ARE FOR EXAMPLE ONLY
TAIL ROTOR HUB
The tail rotor yoke assembly is mounted
on the 90° gearbox output shaft. The yoke
is forged steel and houses the feathering
bearings for the tail rotor blades. The
feathering bearings are staked into the
yoke and are replaceable. The trunnion
consists of two elastomeric bearings that
are bonded to the spindles of a splined
steel trunnion making it one piece.
TRUNNION
TAIL ROTOR GEARBOX
OUTPUT SHAFT
YOKE
FEATHER
BEARINGS
COUNTER
WEIGHTS
BALANCE
WEIGHTS
Balance weights, on each side of the two
blades, are installed on the blade bolts
to counteract out of balance forces from
the rotor. This is done with the adjustment
of the balance of the blades along their
spanwise and chordwise axis. The amount
of weights installed is dependent on the
hub and blade assembly static or dynamic
balance. Counterweights are added to
brackets attached to the yoke along the
chordwise axis, to reduce control forces
during a hydraulic off condition.
PRODUCT DATA
86
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NOTE: PHOTOS AND DRAWINGS SHOWN
ARE FOR EXAMPLE ONLY
TAIL ROTOR PITCH CHANGE
MECHANISM
The tail rotor pitch change mechanism mounts on the right side of the tail rotor gearbox.
It consists of a housing assembly, trunnion assembly, cap assembly, lever assembly,
idler assembly, rod assembly, pitch links, driver plate cross head, and control tube
assembly. Movement of the tail rotor control system is transmitted through the gearbox
by the tail rotor by the pitch change mechanism. The pitch change tube is located in
the output shaft and rotates at the same rpm.
TAIL ROTOR CROSSHEAD
& PITCH LINKS
TAIL ROTOR GEARBOX & PITCH CHANGE
MECHANISM
TAIL ROTOR DRIVE SYSTEM
The tail rotor driveshaft system delivers
torque from the freewheeling assembly
to the tail rotor gearbox. It is mounted
along the rear portion of the engine and
along the top of the tailboom. It is covered
by a two piece composite drive shaft cover.
The system consists of one steel shaft,
one steel oil cooler shaft, and five
aluminum shafts of which four are
interchangeable. These shafts are
connected at each end by a flexible
coupling. The system is supported by six
hanger bearings, each being equipped
with a grease fitting for lubrication.
PRODUCT DATA
FLEXIBLE COUPLING DETAIL
87
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TAIL ROTOR DRIVESHAFT
HANGER BEARING
TAIL ROTOR DRIVESHAFT SEGMENTS
PRODUCT DATA
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ROTOR TRACK & BALANCE PROVISIONS
Provisions have been included for the use of electronic track and balance systems
for the main and tail rotor systems. Power and data connection for the primary
components is on the radio pedestal.
UNIVERSAL TRACKING DEVICE
PROVISIONS FOR TRACKING EQUIPMENT
OIL COOLER BLOWER ASSEMBLY
The oil cooler blower assembly is mounted aft of the aft engine fire wall. The blower
housing is mounted on the helicopter structure and supports the engine and
transmission oil cooler. The oil cooler blower shaft is mounted in two sealed bearings
and is connected to the forward tail rotor drive shaft. The blower shaft passes through
the blower assembly housing, and the blower impeller is bolted to the shaft. This
shaft serves to drive the tail rotor and also to drive the blower impeller which forces
air to the engine and transmission oil cooler.
BLOWER DETAIL
OIL COOLER BLOWER
PRODUCT DATA
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MAIN ROTOR DRIVE SYSTEM
The main rotor drive system provides a means of transmitting power and torque
from the engine to the main rotor. The main rotor drive train system consists of the
following assemblies: main driveshaft, transmission, main rotor mast assembly,
pylon assembly, freewheeling unit, and oil system.
FREE WHEELING UNIT
MAIN DRIVESHAFT
MAST
SPACER
OPTIONAL
ROTOR BRAKE
TRANSMISSION
PYLON ASSEMBLY
MAST
OIL FILLER CAP
MOUNTS
OIL SIGHT GLASS
LOWER RESTRAINT
FITTING
PRODUCT DATA
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MAIN DRIVESHAFT (Kaflex)
The engine to transmission Kaflex drive shaft provides a flexible drive connection
between the freewheeling unit and the transmission. This flexible shaft allows smooth
transfer of torque even though the transmission and freewheeling unit are not always
in perfect alignment. Flexibility of the shaft is provided by rectangular plates, three
in each coupling. Each plate flexes providing both angular misalignment and length
changes to accommodate movement of the transmission on its mounts. Each coupling
can be considered a truss work in which torque loads are carried as axial loads in
straight members of each plate. A fail-safe feature exists which enables uninterrupted
drive of the shaft after a failure has occurred in one of the dual load paths provided
by the plate couplings.
FLEXIBLE PLATES
PRODUCT DATA
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TRANSMISSION ASSEMBLY
The transmission assembly is made up of a top case, support, lower case which
contains a bevel gear and shaft arrangement, an input pinion, a planetary gear train,
sun gear and an accessory gear drive. The components that are attached to the
transmission assembly are the mast assembly, swashplate, hydraulic pump, engine
to transmission driveshaft, monopole sensor, two chip detectors, four oil jets, oil
pump and the oil filter manifold and housing.
The oil filter manifold and housing contain the thermoswitch, thermostatic valve,
temperature bulb and the bypass valve with indicator.
PLANETARY
ASSEMBLY
PINION GEAR
[ONE OF FIVE]
TOP CASE
SUN GEAR
COLLECTOR GEAR
[FIRST REDUCTION STAGE]
INPUT GEAR SHAFT &
SPIRAL BEVEL GEAR
RING GEAR
MAIN CASE
PRODUCT DATA
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TRANSMISSION TOP CASE AND INTERNAL COMPONENTS
TOP CASE The transmission top case is made from an aluminum forging. The pylon
mounting studs on the top case are individually replaceable.
DEBRIS COLLECTOR The debris collector will collect metal particles from the upper
portion of the top case and give warning through the No. 2 electric chip detector
before falling into the main case and entering the oil system.
RING GEAR The steel ring gear, which is part of the second stage of reduction,
allows the planetary assembly to rotate.
PLANETARY ASSEMBLY The planetary assembly is made up of a carrier assembly
and five pinion gears. The top half of the carrier assembly provides the drive for the
mast assembly.
SUN GEAR The sun gear lower splines are splined into the input gear shaft. The
upper splines of the sun gear provides drive for the five pinion gears of the planetary
assembly.
GEAR SHAFT CASE The magnesium gear shaft case houses a roller alignment
bearing that provides alignment for the input gear shaft.
TRANSMISSION MAIN CASE AND INTERNAL COMPONENTS
MAIN CASE The main case provides support for the spiral bevel gear assembly,
and alignment for the lower mast bearing support accessory bevel gear, and the
main input quill. The case is made from magnesium casting. The studs on the main
case are individually replaceable.
INPUT GEAR SHAFT AND INPUT SPIRAL BEVEL GEAR The steel spiral bevel
gear assembly consists of the input gear shaft and input spiral bevel gear.
LOWER MAST BEARING SUPPORT The aluminum lower mast bearing support houses
a rolling alignment bearing. It also provides support for the lower portion of the
mast assembly.
FIRST REDUCTION STAGE [COLLECTOR GEAR] The first reduction stage gear
(COLLECTOR GEAR) is driven by the spiral bevel gear, and in turn drives the sun
gear. Also, it powers the accessory drive gear, which drives the transmission oil
pump and hydraulic pump.
MAIN INPUT QUILL The main input quill supports a triplex bearing set and provides
drive for the spiral bevel gear assembly.
PRODUCT DATA
93
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MAIN ROTOR MAST
The mast assembly is made up of a pole,
bearing assembly and the necessary
hardware to make the installation. The
mast pole is made as follows from the lower
end up:
•a running surface for the alignment of the
mast bearing,
•splines that mesh with the planetary
pinion gears,
•splines for the swashplate drive,
•three cone seats with the middle seat
coated with E-nickel,
•a wear strip bonded at the lower cone
seat,
•splines for the main hub engagement and
•threads for the main rotor hub retention
nut.
MAST POLE
MAST ASSEMBLY
The mast assembly is installed in the transmission assembly and attached to the
transmission top case studs. The swashplate and support assembly are mounted
on top of the mast bearing plate and attached to the transmission top case studs.
The mast pole turns at 413 RPM and absorbs torsion, tension and bending loads
that are made by the engine torque, helicopter weight and flight motion.
PRODUCT DATA
94
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SWASHPLATE ASSEMBLY
The purpose of the swashplate and support assembly is to convert non-rotating
control inputs into rotating control inputs. The assembly encircles the mast directly
above the transmission. The swashplate is mounted on a (collective sleeve) support
which permits it to be tilted in any direction. Movement of the cyclic control results
in a corresponding tilt of the swashplate and the main rotor. Movement of the collective
control actuates the sleeve assembly which raises or lowers the swashplate and
transmits collective control to the main rotor. The cyclic controls are properly
coordinated with collective control by action of the mixing lever at the base of the
control column.
PITCH
HORN
CYCLIC
INPUTS
SWASHPLATE
ASSEMBLY
COLLECTIVE
INPUT
[ROTATING]
SWASHPLATE
ASSEMBLY
[NON-ROTATING]
COLLECTIVE
SLEEVE ASSEMBLY
TRANSMISSION NR MONOPOLE
PICKUP [RPM SENSOR]
The main rotor rpm measuring system measures rotor rpm in percent. This system
measures rpm with a single mag pickup mounted through the main case, taking its
reading from the spiral bevel gear and displaying it on the dual tachometer mounted
on the instrument panel. An RPM warning light illuminates when the upper or lower
limit of the main rotor RPM is reached. The lower limit is accompanied by an aural
warning in the crew headsets. The aural warning is disabled with the collective full
down. The warning light will remain lit until rotor RPM returns to the normal range
when in flight.
PRODUCT DATA
95
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TRANSMISSION GEAR RATIO
The main transmission, located directly forward of the powerplant, provides a two
stage reduction of 15.29 : 1 (6317 to 413 rpm ). The first stage is a bevel gear
arrangement with 3.26 to 1.0 reduction (6317 to 1936 rpm); the second stage reduction
is obtained with a planetary gear train providing 4.69 to 1.0 reduction (1936 to 413
rpm).
The accessory drive gear provides a 1.42 to 1.0 reduction (6317 to 4445 rpm), which
drives the oil pump and hydraulic pump.
GEAR TRAIN SCHEMATIC
NOTE: For location 4 in the above illustration, the ring gear is stationary.
PRODUCT DATA
96
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PYLON ASSEMBLY
[TRANSMISSION MOUNTING]
ELASTOMERIC MOUNTS
A soft mounted System for Attenuating Vibration Independent of Tuning and Dampening
(SAVITAD) pylon assembly is used to decrease the vibration transfer to the structure.
The soft mounted pylon reduces the vibration made by the main rotor. This is done
by the installation of the elastomeric corner mounts and restraint springs with a
natural frequency that is less than the excitation frequency made by the main rotor.
This permits the inertia forces of the pylon to oppose the excitation force of the rotor
and decrease the net force transmitted to the fuselage. More safety is given by the
two deck stop fittings that are attached to the forward top deck and the roof beam of
the helicopter that limit the movement of the two restraint spring fittings attached to
the lower case of the transmission.
The pylon beam assemblies are attached to the transmission top case by two
elastomeric mount assemblies on each side of the lower transmission case and to
each pylon support arm.
SIGHT
GLASS
RESTRAINT
PRODUCT DATA
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TRANSMISSION OIL SYSTEM
The lubrication system is made up of the following:
•oil pump, •pressure regulating valve, •oil cooler, •oil filter assembly, •pressure
transducer and switch, •temperature bulb, •bypass valve and indicator, • two chip
detectors, •and spray oil jets No. 1, No. 2, No. 3, & No. 4., •manifold.
The oil pump, driven by the accessory gear, scavenges oil from the transmission
lower case sump through a 16 x 16 gauge wire screen and a chip detector. From the
pump the oil is delivered under pressure externally to the 3 micron oil filter assembly,
oil cooler and then returns oil to the main transmission internal lubrication channels
and spray jets for lubricating internal parts.
An oil level sight gage is located on the right side of the main transmission lower
case where it can be easily inspected. A non vented filler cap is located on the
main transmission top case, venting is accomplished by an independent vent tube.
PRODUCT DATA
98
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OIL PUMP
The oil pump is a self contained, positive displacement and constant volume Grotor oil pump that has a flow rating of 6 to 6.7 gpm at 4445 RPM. The maximum
pressure is 150 psig with a rated pressure of 80 psig. The oil pump is flush mounted
on the forward left hand side of the transmission main case and is driven by the
accessory bevel gear. The outboard end of the oil pump has an accessory drive pad
for mounting the flight control hydraulic pump.
TRANSMISSION OIL FILLER CAP
The oil filler cap is a non-vented cap located on the forward right hand corner of the
top case.
OIL BREATHER AND LINE
The breather line is installed on the right hand side of the transmission top case. A
rigid line is attached to the breather and it runs along the aft side. The breather
permits the lubrication oil and air in the gearbox to expand or contract when the
temperature changes.
OIL DRAIN PLUG
A transmission oil drain is installed on the forward left hand corner of the transmission
lower case. It is made up of the drain plug, packings and a self closing valve.
FREEWHEELING ASSEMBLY
FILLER
CAP
OIL FILTER
OIL PUMP
DRAIN PLUG
PRODUCT DATA
99
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UNION FILTER
The union filter is located on the aft left hand corner of the transmission lower case.
This union filter is the outlet pressure for the freewheeling unit assembly lubrication
oil.
TRANSMISSION OIL LEVEL SIGHT GAGE
The transmission oil level sight gage is located on the right hand side of the transmission
lower case. You can monitor the oil level or oil condition through the sight gage.
TRANSMISSION OIL JETS
There are four spray oil jets incorporated in the transmission along with the oil passages
in the transmission cases. The #1 jet directs a lubricating oil spray to the alignment
roller bearing of the input pinion and also guides part of the oil flow to the annulus
slot which lubricates the gearshaft duplex bearing. The #1 jet also supplies oil to oil
jets #3 and #4. The #2 jet lubricates the planetary pinions and mast bearing, #3 jet
lubricates the entering mesh of the input bevel and the #4 jet lubricates the exiting
mesh of the input bevel for cooling.
TRANSMISSION OIL MANIFOLD & FILTER ASSEMBLY
The transmission oil filter manifold with housing assembly is mounted to the left
hand side of the transmission lower case. There are two bushings that are fitted in
the manifold to permit the oil flow transfer from the lower case to the manifold to the
lower case. A housing that contains the oil filter element is mounted to the manifold.
An oil filter differential pressure bypass valve and indicator is attached to the filter
housing button. The manifold serves as the mount point to the high pressure bypass
valve, a temperature bulb, a thermoswitch, a thermostatic valve and two fittings.
The two fittings permit the oil circulation to the oil cooler and back.
OIL FILTER ELEMENT AND HOUSING
The transmission oil filter element is located inside the oil filter housing that is attached
to the oil filter manifold. The filter sits inside the housing on the bypass valve and
indicator. The filter element is a 3 micron pleated paper type with a wire mesh
cover.
PRODUCT DATA
100
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TRANSMISSION OIL PRESSURE REGULATOR VALVE
The transmission oil pressure regulating valve is located on the left rear corner of
the transmission main case. The regulating valve is used to adjust transmission oil
pressure to normal operating limits and relieves excess oil pressure back into the
transmission case.
TRANSMISSION HIGH PRESSURE OIL FILTER BYPASS VALVE
The transmission high pressure oil filter bypass valve is located on the bottom of
the oil filter manifold. It will open when there is too much back pressure caused by
a clogged oil filter element. The bypass valve will then bypass the oil filter element
and direct the oil to the thermostatic valve and the oil cooler. The bypass valve will
open when the differential pressure across the filter reaches 29.6 to 38.6 PSID.
FILTER BYPASS BUTTON
TRANSMISSION OIL FILTER BYPASS VALVE & INDICATOR
The transmission oil filter bypass valve is installed in the oil filter housing. There
are two functions of the oil filter bypass valve and indicator. First it will give an
impending bypass indication at 14.0 PSID (±2.0 PSID). This will be indicated by the
extension of the red impending bypass button. Then at a cracking pressure of 17
PSID MIN. the relief valve will begin to open, with a full flow bypass at 27.0 PSID.
A thermal lockout is included in the system to prevent actuation of the indicator
button until the oil temperature has reached 120° F (49° C). This feature will prevent
the indicator button from activation on a cold day during start up.
TRANSMISSION OIL THERMOSTATIC VALVE
On start up when the transmission thermostat valve is open, lubricating oil flowing
to the oil cooler is residual. The main flow of lubricating oil during run up, to an oil
temperature less than 150°F (65.6°C), is back into the main case. AT 150° F the
thermostat valve begins to close, as the valve closes the main oil flow is then directed
to the oil cooler. At 178°F (81.1°C ), ± 2°F, the thermostat valve is closed on seat
and all of the oil flow is directed through the cooler. The total movement of the valve
from 150°F to 178°F is .080 inch minimum for valve to be closed.
The thermostatic valve also has bypass capabilities. The valve will open slightly to
permit the oil to bypass the oil cooler when the differential pressure across the oil
cooler reaches 40 PSID.
PRODUCT DATA
101
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TRANSMISSION OIL TEMP BULB
The transmission oil temperature bulb is located on the oil filter manifold is electrically
connected to oil temperature indicator on the instrument panel.
TRANSMISSION OIL THERMOSWITCH
The transmission oil thermoswitch is located on the filter manifold and is electrically
connected to the TRANS OIL TEMP caution light segment. When the transmission
oil reaches a temperature of 230°F (110°C) the thermoswitch will illuminate the caution
light segment.
TRANSMISSION OIL SYSTEM HOSES & TUBE ASSEMBLIES
The transmission oil system hose and tube assemblies consist of both aluminum
hard lines and flexible steel braided hoses.
TRANSMISSION OIL PRESSURE SWITCH
The transmission oil pressure switch, located on the transmission oil manifold, is
connected to a TRANS OIL PRESS caution light segment on the master caution
panel. During normal operation, the oil pressure switch contacts open at 38 psi
maximum on increasing oil pressure and the caution light is extinguished. The switch
contacts close at 30 ±2 psi on decreasing oil pressure to cause the illumination of
the caution light segment.
TRANSMISSION &
FREEWHEELING CHIP
DETECTOR
The chip detector is made up of a self
locking bayonet type probe with a
permanent magnet at the end. If metal
particles become free in the oil system,
the magnet will pick them up until there
are enough metal particles to complete
the circuit between the pole and ground.
This in turn will activate the applicable
caution annunciator on the caution/
warning and advisory panel.
TRANSMISSION OIL PRESSURE
TRANSDUCER
The transmission oil pressure transducer,
located on the transmission oil filter
manifold, is connected to the transmission
oil pressure gauge.
PRODUCT DATA
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TRANSMISSION [& ENGINE] OIL COOLER
The oil cooler assembly contains two separate independent cores welded together.
The aft core is for the engine oil system and the forward core is for the transmission
oil system. The cooler is secured to the oil cooler blower assembly by means of a
transition duct.
HOT AIR EXHAUST
AMBIENT AIR INLET
OIL COOLER
TRANSITION
DUCT
BLOWER HOUSING
PRODUCT DATA
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FREEWHEELING ASSEMBLY
The freewheeling assembly is mounted on the engine accessory gearbox. The
input shaft of the freewheeling assembly is splined directly to the engine power takeoff
gear shaft. The engine power is transmitted to the input shaft of the freewheeling
assembly that then engages the freewheeling output shaft through the sprag elements
of the clutch assembly. The freewheeling output shaft is coupled on the forward
end to the engine to transmission driveshaft; the aft end to the tail rotor driveshaft
system. During autorotation, the main rotor assembly drives the freewheeling output
shaft. When this condition occurs, the freewheeling input shaft is disengaged from
the freewheeling output shaft through the effect of the sprag clutch. The engine is
then disconnected and the rotational force produced by autorotation of the main
rotor can drive the tail rotor drive system and transmission with all its mounted
accessories
FREEWHEELING UNIT
PRODUCT DATA
104
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POWER PLANT ASSEMBLY
The Bell 407 Helicopter is supplied with power by an Rolls-Royce model 250-C47B
turboshaft engine. The engine has a single stage centrifugal-flow compressor, a
single combustion chamber, a two stage gas producer turbine, and a two stage power
turbine. This engine has a Full Authority Digital Engine Control (FADEC). The power
plant assembly is horizontally installed aft of the transmission in the engine pan
above the fuselage. The engine is held in place by three bipod mounts and a sway
bar which are attached to the service deck. The engine is then connected to the
transmission by the freewheeling unit and main driveshaft.
PRODUCT DATA
105
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RELIABLE POWER
Manufactured by Rolls-Royce, the 250-C47B is one of the latest in the evolutionary
Model 250 series, which has amassed over 100 million engine flight hours. During
this evolution, many improvements have been incorporated. The engine is light in
weight and of minimum size. Easy to maintain, all engine adjustments and almost
all inspections and routine maintenance can easily be performed while the engine
is still installed on the helicopter.
MODULAR CONSTRUCTION
The Rolls-Royce 250-C47B turboshaft engine features modular design for easy
access to the major components. One module can be replaced quickly without disturbing
the others. Modular design offers substantial direct operating cost advantages
because maintenance or overhaul of a single module can be accomplished
independently. The compressor, for instance, may be removed for repair or overhaul
while the remainder of the engine continues in service with a replacement (overhaul
exchange unit / temporary leased) compressor.
FOUR MODULES
The engine consists of four major modules; a compressor section, an accessory
gearbox, a turbine section and a combustion section. Some of the benefits of the
modular concept are shown below.
COMPRESSOR MODULE
TURBINE MODULE
•READY FOR OVERHAUL OR REPAIR BY
REMOVING SIX GEARBOX MOUNTING
BOLTS.
•SIMPLE SINGLE STAGE IMPELLER.
•NO VARIABLE GEOMETRY.
•MOUNTED TO THE REAR OF THE
GEARBOX WITH FIVE NUTS.
•FIRST STAGE TURBINE ACCESSIBLE BY
REMOVING THE COMBUSTOR MODULE.
•ALL ACCESSORIES AND SHAFTING
REMAIN IN PLACE.
GEARBOX MODULE
•RUGGED CONSTRUCTION, COMMON MOUNTING PADS FOR ALL
MODEL 250 SERIES ENGINES.
•ALL LIP SEALS CHANGED EXTERNALLY.
•ROTATING ACCESSORIES USE COMMON
NUTS.
•BOTH FRONT AND REAR DRIVE.
•EASY ACCESS TO THE ACCESSORIES
FOR QUICK REMOVAL AND
REPLACEMENT.
PRODUCT DATA
COMBUSTOR MODULE
•SIMPLE ONE PIECE OUTER
COMBUSTOR, SINGLE COMBUSTOR
LINER.
•SINGLE FUEL NOZZLE.
•SINGLE SPLIT LINE FOR REMOVAL
FROM TURBINE.
•COMPLETE COMBUSTOR SYSTEM
HARDWARE CAN BE INSPECTED BY
REMOVAL OF COMBUSTOR MODULE.
106
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ENGINE AIRFLOW
The Rolls-Royce 250C-47B TURBOSHAFT engine consists of a compressor section,
combustion section, turbine section , and a power and accessory gearbox. Like
any internal combustion engine, it incorporates intake, compression, combustion,
and exhaust to produce power. There is no " power stroke" cycle required for the
absorption of energy from the gases of combustion in a turbine engine. In a turbine
engine, the combustion cycle is continuous and the rotation of the turbine directly
generates shaft horsepower.
250C-47B ENGINE
PRODUCT DATA
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ENGINE OPERATION PRINCIPAL
Both the absorption of energy from the
gases of combustion and the development
of shaft horsepower are accomplished
through the use of the four turbine stages
which are located between the combustion
and exhaust sections. These turbines
provide the means of extracting gas
energy and converting it into mechanical
energy in the form of horsepower.
C47B ENGINE SECTIONS AND AIRFLOW SCHEMATIC
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ENGINE BELLMOUTH
ASSEMBLY
The engine bellmouth assembly is made
of aluminum alloy and makes an airfoil inlet
for the inlet air to the engine. The engine
bellmouth is attached to the forward side
of the forward firewall and the front of the
engine compressor.
ENGINE
BELLMOUTH
ENGINE MOUNTS, ENGINE DECK, & FIREWALLS
ENGINE MOUNTS
The engine is installed in the engine pan
with three bipod type mounts and one sway
bar. The mounts are found on the left,
right and aft side of the engine. The sway
bar is installed on the front of the engine
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under the gearbox. The engine mount legs
and sway bar are made of steel tubes and
riveted to clevises and fittings. Shims are
used between the fittings and airframe to
align the engine when it is installed.
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AIR INTAKES
The air intakes are made up of an engine air inlet cowling and an engine inlet screen
assembly. The screen assembly is installed on the forward side of the forward firewall.
The engine inlet screen must be installed before the engine is started. This will prevent
possible engine damage from any unwanted material that is sucked in when the
engine is in operation. For operation in dusty environments, an optional engine air
particle separator is available.
INLET
SCREEN
COMPRESSOR
WASH TUBING
ENGINE INLET SCREEN ASSEMBLY
The engine inlet screen assembly is between the engine inlet cowling and the engine
intake. The engine inlet screen assembly is attached to the forward firewall. The
screen assembly is made up of two screens that are installed one on top of the other
and permits air to go through while it stops unwanted objects that could get into the
engine. The two screens are set apart to permit inlet air to pass between the screens
if the screens are blocked.
COMPRESSOR WASH SYSTEM
The engine compressor can be washed through a compressor wash provision that
is installed on the engine inlet screen assembly. This provision is made up of a right
angle fitting, tube, clamp, and plug. The tube and fitting are connected to a bracket
on the engine inlet screen. The provisions are conveniently located inside the right
transmission inspection door, allowing the engine wash to be accomplished without
removing cowling, or panels.
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COMPRESSOR SECTION
BLEED AIR VENT
SCROLL
The compressor assembly is a single stage, single entry centrifugal flow compressor.
The compressor takes air from the inlet through the compressor rotor. The rotor
vanes accelerate the air into a decreasing space to increase the air pressure and
temperature. The impeller discharges air into the vanes of the diffuser. The diffuser
has 25 vanes with an annular pressure equalization chamber on the rear outer diameter.
The diffuser vanes direct air into the diffuser scroll. The diffuser scroll collects the
compressor outlet flow at a constant velocity and directs the air into the transfer
tubes which direct the air into the combustion assembly. Compressor air flow capacity
has been increased by 7% over the 250-C30 series engines. Standard day air flow
is 2.7 Kilograms per second (5.92 pounds per second). This is the result of increasing
the impeller blade length at the front and rear. A new shroud housing matches the
revised impeller blade contour. A new abradable coating is applied to the rear sides
of the housing. The inducer bleed port size is increased to match the increased air
flow.
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BLEED AIR SYSTEM
An acceleration bleed air system is incorporated in the compressor assembly.
The system allows a portion of the compressed air to escape during the start/
acceleration cycle to prevent stalls, power surges, and to assure rapid response to
engine power demands. The system consists of a pneumatically operated bleed
valve, and an inducer bleed assembly. As NG increases, the bleed valve gradually
closes to reserve air for engine operation. The inducer bleed also functions to enhance
compressor efficiency when NG is above 96 %.
NO. 1 BRG. OIL IN
INDUCER
BLEED PORT
ANTI-ICING AIR
ACCELERATION
BLEED VALVE
NO. 1 BRG. OIL OUT
BLEED AIR SOURCES
Four ports along the diffuser scroll are provided to supply bleed air. One provides
air to the anti-icing system, and the remaining three are available as bleed air
sources to power airframe systems (optional bleed air heater).
ANTI-ICING AIR SYSTEM
The anti-icing air system is made up of the following components :
• anti-icing solenoid valve mounted on the diffuser scroll of the engine, • air valve
mounted on the firewall shield, • engine anti-icing switch mounted on the overhead
controls panel, • engine anti-icing circuit breaker on the overhead console.
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The anti-icing solenoid valve is powered from the 28V essential bus through the
engine anti-icing switch. In the event of a total electrical failure the solenoid is failsafe to the ON condition. When the anti-icing solenoid valve is deenergized, it opens
and hot air passes from the engine compressor diffuser through the anti-icing solenoid
valve to the engine inlet housing. The hot air helps to prevent ice from forming on
the hollow inlet vanes. Engine anti-ice must be ON when operating in visible moisture
if the ambient temperature is at or below 5°C (40°F). When anti-ice is ON an advisory
(white) caution panel segment is illuminated.
STARTER GENERATOR
The starter-generator is installed on the
bottom of the engine to the right of the
helicopter centerline. The startergenerator has a duct to give a flow of air
to cool it. The starter-generator supplies
DC power and the start functions for the
engine.
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ACCESSORY GEARBOX
The main power and accessory drive gear trains are enclosed in a single (split) gear
case. The gear case serves as the main structural support of the engine. All engine
components including the engine mounted accessories are attached to the case.
The case has two sections. The forward (HOUSING) or compressor section, mounts
accessories driven by the gas producer gear train. The aft (COVER) or turbine section,
mounts accessories driven from the power turbine gear train.
NG & NP
MONOPOLE
PICKUPS
FORWARD VIEW OF ACCESSORIES
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AFT VIEW OF ACCESSORIES
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GAS PRODUCER GEARING
[ LEFT IS FORWARD ]
TURBINE TO
COMPRESSOR
COUPLING
2
COMPRESSOR
ROTOR
GAS PRODUCER
TURBINE ROTOR
3
1
4
DRIVES FADEC H.M.U.
& ENGINE OIL PUMP
& NG TACHOMETER
5
6
SPARE
7
STARTER / GENERATOR
DRIVE
1.
2.
3.
4.
5.
6.
7.
SPUR ADAPTOR GEARSHAFT
SPUR CENTRIFUGAL BREATHER GEARSHAFT
SPUR IDLER GEARSHAFT
SPUR IDLER GEARSHAFT
SPUR HYDRO MECHANICAL UNIT & OIL PUMP & NG TACH GEARSHAFT
GENERATOR IDLER GEAR
STARTER GENERATOR GEARSHAFT
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POWER TURBINE GEARING
[ LEFT IS FORWARD ]
POWER TURBINE
SHAFT
8
TURBINE SHAFT TO
PINION GEAR COUPLING
DRIVES P.M.A.
9
12
POWER TURBINE
ROTOR
10
11
DELIVERS ENGINE OUTPUT
POWER TO MAIN DRIVESHAFT (FWD)
AND TAIL ROTOR DRIVESHAFT (AFT)
13
14
SPARE DRIVE - 15 SHP
8. HELICAL POWERTRAIN DRIVE GEAR
9. HELICAL TORQUEMETER GEARSHAFT
10. HELICAL POWER TAKEOFF GEARSHAFT
11. SPUR POWERTRAIN IDLER GEAR
12. SPUR POWERTRAIN PERMANENT MAGNET ALTERNATOR
13. SPARE DRIVE IDLER HELICAL SPUR GEARSHAFT
14. SPARE SPUR DRIVE GEARSHAFT
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TORQUEMETER
The torquemeter system in the 250-C47B engine gearbox is a hydraulic type that
uses the engine lubrication system as its oil (hydraulic) pressure source. The power
turbine gear train has two stages of helical gearing. Helix angles are such that both
stages of reduction produce forward axial thrust on the helical torquemeter gear
shaft. The axial thrust is directly proportional to the torque transmitted through the
gears. The axial thrust acts on a sliding piston which increases or decreases the
pressure in the torquemeter oil chamber. Pressure is sensed by a transducer , and
the resulting signal is transmitted to the engine torque gauge.
TORQUEMETER SCHEMATIC
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COMBUSTION SECTION
The combustion section consists of an outer case, a combustion liner, an igniter
and a fuel injector. The prechamber combustion system is designed to burn fuel at
peak efficiency and to minimize emissions. Fuel is sprayed into the chamber at a
precisely controlled rate. The single igniter produces a constant spark while the
start circuit is energized, until automatic starter disengage at 50% NG.
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AIR COOLING
Approximately 20% to 25% of the air entering the combustion chamber is required
to burn the fuel, while the remaining air is used for engine cooling and thrust balance.
The cooling air is directed into the combustion liner, to prevent the flame from
contacting the liner wall, and to cooling air passages in the gas producer turbine
support, the turbine nozzles, and the first stage nozzle shield.
Some of the high-pressure cooling air is delivered to the thrust balance chamber
at the forward face of the second stage gas producer turbine rotor to offset forward
pressure from combustion gas flow and reduce loads on the No.8 thrust bearing.
TURBINE COOLING & BALANCE AIR SCHEMATIC
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TURBINE SECTION
The turbine section consists of a gas producer turbine support, a power turbine support,
an exhaust collector support, a two-stage gas producer turbine, and a two-stage
power turbine. The gas producer, power turbine, and exhaust collector supports
are high temperature resistant structures which house support and thrust bearings,
oil and air seals, passages for lubrication, cooling air and thrust balance air,
temperature sensing thermocouples, and the outer casings which connect the
combustion section and the accessory gearbox. The two turbine units are free to
revolve independently of one another. Each turbine stage consists of a stator, or
nozzle (stationary guide vanes which direct gases) and a rotor, or turbine wheel
(reaction driven by combustion gases).
TURBINE SECTION SCHEMATIC
The first stage turbine vane assembly flow area is increased from the C30 series
engines to handle greater air flow from the compressor. The vanes are hollow with
cooling air distribution tubes inserted and welded at the outer diameter. Each
distribution tube has ten air holes on the side so the cooling air can cool the inside
of the vane leading edges. The air then flows forward through the vane interior and
exits through two slots in the vane trailing edges. The vane assembly inner band is
slotted to allow for expansion. The slots are covered by metal sealing strips to reduce
hot gas leakage. Turbine operating temperatures are increased approximately
7°C(20°F).
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EXHAUST SYSTEM
The exhaust system is made up of an
exhaust duct assembly attached to the
engine exhaust flange and is sealed to
the upper engine cowling assembly.
INDUCER BLEED
DISCHARGE
EXHAUST DUCT
[STACK]
EXHAUST DUCT ASSEMBLY
The exhaust duct assembly lets out the
exhaust gases. The duct assembly is
made of corrosion resistant steel. The
duct assembly is attached to the engine
exhaust flange with nuts and bolts.
ENGINE BLEED AIR
DISCHARGE TUBE
The bleed air discharge tube removes the
exhaust from the engine compression
chamber.
EXHAUST COLLECTOR SUPPORT
INSULATED COVER
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ENGINE CONTROL SYSTEM [FADEC]
A Full Authority Digital Electronic Control (FADEC) System is used on the 250C47B engine.
System objectives are to :
1. Improve engine performance. 2. Enhance flight safety. 3. Reduce pilot work load.
4. Increase mean time between removals/overhauls. 5. Reduce life cycle costs. 6.
Simplify engine maintenance and trouble shooting
The FADEC fuel and control system consists of :
1. Hydromechanical Unit (HMU)
2. Fuel filter-oil filter assembly
3. Electronic Control Unit (ECU)
4. Permanent Magnet Alternator (PMA)
5. Compressor inlet air temperature sensor, inlet duet mounted
6. P1 compressor air inlet pressure sensor in the ECU
7. Engine sensors of RPM, MGT andTorque
8. Aircraft supplied inputs
The HMU contains the following components :
1. Two element fuel pump
2. Fuel metering valve
3. Manual fuel control
4. Three solenoid valves
5. Two check valves
6. Fuel shut-off valve with a pilot valve(windmill bypass valve)
7. Bypassing type pressure differential regulator valve
8. Power lever input shaft
9. Two feedback potentiometers
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ELECTRONIC CONTROL UNIT
The Series IV 250-C47 engine use a digital electronic control system, based upon
a single-channel, Full Authority Digital Electronic Control (FADEC). It is also referred
to as the Electronic Control Unit (ECU). The system controls, monitors and limits
the engine while maintaining helicopter rotor speed. The control connection between
the helicopter and the engine is both electrical and mechanical. This interface includes
a combination of Power Lever Angle (PLA) [twist-grip throttle] input, hard wired
discretes, analog signals, and a RS-423 data link for ground maintenance interrogation.
Any fault in the ECU which requires pilot action automaticaly initiates a transition to
MANUAL MODE [Direct Reversion To Manual-DRTM]. An ARINC-429 digital data
link is used to transmit information on engine and control status to the aircraft.
ELECTRONIC CONTROL UNIT [ECU]
HYDROMECHANICAL CONTROL UNIT [HMU]
HYDROMECHANICAL CONTROL UNIT
The Hydromechanical Unit consists of the dual element fuel pump, fuel metering
valve, manual fuel control, three solenoid valves, two check valves, a fuel shutoff
valve with a pilot valve (windmill bypass valve), bypassing type pressure differential
regulator valve, power lever input shaft, and two feedback potentiometers. During
automatic operation of the FADEC, fuel metering by the HMU is controlled
electronically by the ECU. In the manual mode the pilot’s throttle input is tied
mechanically to the fuel flow metering window in the HMU. The manual mode is
engaged by deenergizing a solenoid in the HMU via a cockpit switch, or in the event
of an inflight fault the Direct Revision To Manual System [DRTM] automaticaly
deenergizes the solenoid. This allows the pilot to vary fuel flow to the engine by
moving the twist-grip throttle (increase or decrease) as required from the normal
70% PLA FLY detent. This manual mode fuel flow is altitude compensated to allow
a consistent throttle / horsepower relationship verses altitude. At 100% throttle travel,
the manual mode will provide at least maximum continuous power. Additionally, the
fuel flow slew rate is mechanically limited to avoid blowout and to provide proper
responsiveness for aircraft operation.
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FADEC SYSTEM SCHEMATIC
Advantages during starting include :
1. Faster and cooler starts 2. Improved cold weather and low battery starts 3.
Overtemperature start abort
Advantages during operation include:
1. Improved power turbine governing 2. Torque limiting 3. MGT limiting 4. Constant
transients, insensitive to fuel type, temperature and metering consistencies. 5.
Improved autorotation recoupling minimizes rotor RPM droop 6. Surge detection,
recovering, avoidance
Advantages regarding safety include:
1. Improved overall engine reliability 2. High reliability electronics 3. Dual channel
NP. overspeed shutoff, tested at shutdown by procedure 4. PMA electrical power 5.
No pneumatics
Advantages regarding maintainability include :
1. Supportable 2. No field adjustments 3. Exceedance recording
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FADEC SYSTEM OPERATIONAL FEATURES
Features of the FADEC are listed below:
GAS GENERATOR (NG) GOVERNOR
In the auto mode the pilots PLA (throttle) controls the set point for the NG governor.
This allows the pilot to limit engine power as desired and provides smooth transition
from NG governing at ground idle to power turbine speed (NP) governing at flight
idle. Normal flight operations shall be conducted with the throttle twist grip at 100%
travel. If desired, the PLA (throttle) can be rolled back to force a power turbine speed/
helicopter rotor speed (NP/NR) needle split for autorotation training.
POWER TURBINE (NP) GOVERNOR
The ECU governs NP to maintain helicopter rotor speed (NR) at the set point. The
control utilizes isochronous speed governing with gains and compensation optimized
for the engine installation. A collective pitch position analog input signal provided
by the airframe provides load anticipation for the NP speed governor. This anticipation
initiates NG acceleration after collective pitch movement prior to actual load increase,
thus reducing rotor speed droop. The rotor speed input frequency signal provided
to the control by the aircraft enhances autorotational recovery. Using the rotor speed
input, the FADEC increases power turbine speed to match rotor speed during the
reapplication of rotor load, thus minimizing rotor speed droop/overshoot.
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FADEC SYSTEM FEATURES OPERATION (continued)
EXCEEDANCE LIMITING
Automatic limiting functions accomplished by the FADEC include turbine temperature
limiting, engine torque limiting, NG speed limiting, and NP speed limiting.
The FADEC system interfaces to the MGT harness to measure engine temperature.
When the engine is approaching its steady state or transient temperature limit, the
FADEC reduces fuel flow to prevent limit exceedance. A smooth, controlled transition
between governing and temperature limiting is accomplished by the FADEC.
The FADEC system interfaces to the TMOP (torque meter oil pressure) torque
sensor to measure output shaft torque. The FADEC reduces fuel flow in response to
measured torque to prevent limit exceedance.
SURGE DETECTION AND RECOVERY
The FADEC detects engine surge by comparing the rate of change of NG speed to a
predetermined boundary rate. If the boundary is exceeded and MGT is increasing,
the surge will be detected and recorded by the internal ECU EMS. The surge will be
recorded in the ECU’s memory relative to the NG speed at which it occurred. Without
pilot action, the FADEC will reduce fuel flow during the surge and reduce the maximum
acceleration schedule during the current acceleration in order to quickly recover
from the surge. The FADEC will then lower the acceleration schedule at the NG
where the surge occurred to avoid subsequent surge. The acceleration schedule is
reset to the original schedule at FADEC power up transition of the auto/manual switch
from manual mode to auto mode.
AUTORELIGHT
The FADEC declares an engine flameout by detecting an NG deceleration rate greater
than a predetermined flameout boundary rate. The engine flameout is enunciated
to the aircraft via the ARINC-429 data bus. Without pilot action, the flamedout engine’s
FADEC will initiate a restart sequence, which includes scheduling the appropriate
restart fuel flow as a function of operating and ambient conditions and activating
the ignition system. Relight will be detected by the FADEC and the turbine temperature
will be controlled. The engine will then smoothly accelerate back to the commanded
operating condition. The Engine Out discrete will turn off after a minimum NDOT or
increasing MGT is established. No pilot action is required unless the engine fails to
relight by 50% gas generator speed (NG). In the event of an unsuccessful relight,
the engine PLA (throttle) should be moved to cutoff.
AUTOMATIC START
The FADEC control system provides automatic start sequencing and engine control
during the engine starting cycle. This involves controlling fuel flow until stabilized
idle gas generator (NG) speed is reached. Starting is initiated by the pilot by placing
the throttle in the ground idle position and activating the start switch. Once the required
light off NG speed of 12% is achieved (10% if ambient temperature is less than 20°F),
the FADEC introduces fuel to the engine. The engine fuel flow is then regulated to
control the NG turbine rate of acceleration (NDOT) and to maintain a turbine
temperature (MGT) limit while accelerating to ground idle. Pilot fuel modulation is
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FADEC SYSTEM FEATURES OPERATION (continued)
not required. Additionally, the control can prevent most overtemperature starts by
automatically shutting fuel flow off should NG speed, battery voltage, or MGT exceed
“boundary limits” during the start.
SYSTEM BUILT IN TEST
The FADEC system incorporates logic and circuitry to perform a high level of selfdiagnostics, some of which are described herein. In general all sensors are checked
for continuity, rate, and range. Discrete inputs are checked for continuity. Output
drivers are monitored for current demand to sense failed actuators. Output drivers
are current limited. A FADEC power up check exercises all output drivers and actuators
to ensure system functionality and readiness. All failure information is recorded by
the FADEC EMS and is available to the maintenance crew via the RS423 maintenance
data bus.
ENGINE OVERSPEED PROTECTION
The FADEC ECU continuously monitors for NG and NP, or NP versus Torque (Q)
overspeed conditions in both AUTO and MANUAL modes.
The FADEC overspeed system provides two levels of protection;
OVERSPEED WARNING & EXCEEDANCE RECORDING
Activation of the ENGINE OVSPD warning light will occur in the event of an NP
overspeed, NG overspeed, or if NP versus Torque is above the continuous limit
(102.4% NP at 100% Torque to 108.6% NP at 0% Torque). The light will also momentarily
illuminate during the overspeed system test when the solenoid valve closes. If the
ENGINE OVSPD light is activated by an exceedance, and the value has been recorded
by the ECU, the pilot will be provided with a maintenance advisory on shutdown in
the form of a FADEC Degraded light.
OVERSPEED LIMITING
NP OVERSPEED SYSTEM
Power turbine overspeed limiting in both Manual and Auto modes is provided by
independent analog circuits integral to the ECU. The engine fuel flow is reduced to
minimum fuel flow (34 t0 45 pph) upon activation of the overspeed solenoid valve
(at 118.5±1%NP). Once the NP speed drops to 112.5±1%, the overspeed solenoid
valve will be deactivated and fuel flow will return to the previously commanded value.
NG OVERSPEED
In AUTO mode, a software implemented overspeed system is provided. Should the
software detect an NG speed above the NG overspeed threshold, the ECU will activate
the overspeed solenoid valve and reduce fuel flow to minimum flow. Once the NG
speed drops below the overspeed threshold, the solenoid valve opens and normal
operation may resume. In addition, if the ECU has not failed, NG overspeed protection
will also be available in MANUAL mode. When the engine reaches 110±1%NG,
overspeed limiting will occur. Once the NG speed drops to 107±1%, the solenoid
will be deactivated and fuel flow will return to the previously commanded value.
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POWER UP FUNCTION CHECK
The power up check occurs when the ECU is first turned on. This check ensures
electrical continuity of the overspeed circuit and the ability of the ECU to power the
overspeed solenoid. This test is performed by turning on each of the overspeed
solenoid drivers and measuring the voltage and current draw across the overspeed
solenoid valve. The measured voltage and current are then compared to limits.
CONTINUOUS FUNCTIONAL CHECK
Continuous checks occur during normal engine operation. These checks monitor
the functionality of the NP speed signals, which supply the overspeed system. The
two NP speed signals that supply the overspeed system are continuously compared
for differences, and should a difference become larger than a predetermined limit,
a fault is declared.
OVERSPEED SYSTEM SHUTDOWN TEST
Functionality of the overspeed system is checked during FADEC power up and
continuously thereafter by the ECU. Operation of the overspeed solenoid is checked
periodicaly by the pilot through use of the OVERSPEED SHUTDOWN TEST. The
OVERSPEED SHOTDOWN is only performed on the ground. Collective pitch must
be below 10%, throttle position at idle, NG is between 60-66% and NP is less than
75%. The Overspeed Test Button is pressed for a minimum of 1 second and no more
that 10 seconds. Once the test button is released the overspeed test sequence is
initiated. The ECU signals the overspeed solenoid valve to close and the ENGINE
OVSPD light to come on. Once the ECU senses an NG decrease greater than 0.5%,
the overspeed solenoid valve is opened, the ENGINE OVSPD light goes off, and the
engine is shut down by the FADEC ECU activation of the hot start abort feature. If
the test is unsuccessful, the engine will continue to operate at idle power, the FADEC
FAULT light will come on, and a normal shutdown must be performed.
ENGINE CONDITION MONITORING
The ECU contains a basic engine monitoring system to record and store engine and
system fault information. The information is recorded in nonvolatile memory for retrieval
by maintenance personnel via RS-423 digital data communication link to portable
terminal. The recorded parameters include:
•engine running time •compressor module running time •turbine module running
time •number of engine starts •highest peak value and time exceeding turbine speed
limit (NG and NP) •highest peak value and time exceeding turbine run temperature
limit (MGT) •highest peak value and time exceeding engine output shaft torque
limit •highest peak value and time exceeding engine start temperature limit (MGT)
•occurrence of engine surge
To permit transfer of engine ECU’s from one aircraft to another, data may be read
from and input to the ECU nonvolatile memory via the RS-423 digital data link. These
data include:
•engine serial number •ECU serial number •compressor serial number •NG turbine
serial number •NP serial number
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FADEC SYSTEM FEATURES OPERATION (continued)
MANUAL MODE
In the manual mode the pilot’s PLA input is tied mechanically to the fuel flow metering
window in the HMU. The manual mode is engaged by deenergizing a solenoid in the
HMU via a cockpit switch. This allows the pilot to vary fuel flow to the engine by
moving the PLA. This manual mode fuel flow is altitude compensated to allow a
consistent PLA horsepower relationship verses altitude. At 100% throttle travel,
the manual mode will provide at least maximum continuous power. Additionally, the
fuel flow slew rate is mechanically limited to avoid blowout and to provide proper
responsiveness for aircraft operation.
ARINC-429 DATA BUS
The ECU provides a transmit only ARINC-429 data bus. This data bus contains engine
parameters such as NG speed, NP speed, MGT, and torque.
FADEC
MAINTENANCE
SWITCH
FADEC
DATA
PORT
PERMANENT MAGNET ALTERNATOR
ENGINE FUEL/OIL FILTER
PRODUCT DATA
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ENGINE FAULT ANNUNCIATORS
There are seven lights included in the caution panel that are controlled by the FADEC.
They are :
FADEC FAIL, FADEC DEGRADED, FADEC FAULT, FADEC MANUAL, RESTART
FAULT, ENGINE OVSPD, and AUTO RELIGHT
The FADEC FAIL, FADEC DEGRADED, FADEC FAULT and RESTART FAULT lights
are controlled by faults that are detected by the FADEC/ECU.
The FADEC/ECU continuously monitors the FADEC system for faults and makes
appropriate accommodations to continue operation. Fault codes have been
preassigned to those parameters being monitored by the FADEC/ECU. When a
fault is detected by the ECU the fault code is stored in the permanent memory of
the ECU. Depending on the nature of the fault the ECU may immediately display
the fault to the pilot via a light and horn, or the fault may not be displayed until the
throttle is placed in the cutoff position. The multiple fault codes potentially generated
by the ECU have been assigned to specific lights and horn based on the affect of the
fault to the operation of the helicopter. The flight manual provides the appropriate
action required by the pilot for each light/horn indication.
All FADEC faults have been categorized into FIVE TYPES. The first four relate to
in-flight faults and the fifth relates to displayed faults with the engine shut down.
IN-FLIGHT FAULTS
FIRST are those faults that require immediate pilot action and automaticaly activate
the DRTM system. These faults will be displayed on the caution/warning panel
immediately when detected by the ECU. The light used to display these faults is
FADEC FAIL. The FADEC FAIL HORN, will also be activated in conjunction with the
FADEC FAIL light. In conjunction with activation of the FADEC FAIL light, the FADEC
DEGRADED, FADEC MANUAL, FADEC FAULT and RESTART FAULT light (s) may
also activate under certain fault conditions.
SECOND are those faults that represent a loss of some feature of the FADEC system
which may cause a degradation in performance, but operations should be continued
in AUTO mode. These faults will be displayed on the caution/warning panel immediately
when detected by the ECU. The light used to display these faults is FADEC DEGRADED.
In conjunction with activation of the FADEC DEGRADED light, the FADEC FAULT
and RESTART FAULT light (s) may also activate under certain fault conditions.
THIRD are those faults that represent a loss of redundant features that should not
cause degradation in performance. These faults will be displayed on the caution/
warning panel immediately when detected by the ECU. The light used to display
these faults is FADEC FAULT. In conjunction with activation of the FADEC FAULT
light, the RESTART FAULT light may also activate under certain fault conditions.
FOURTH are those faults that do not require immediate action by the pilot and should
not affect performance of the aircraft but may prevent a subsequent engine start.
These faults will be displayed on the caution/warning panel immediately when they
are detected by the ECU. The light used to display these faults is RESTART FAULT.
PRODUCT DATA
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ENGINE SHUT DOWN FAULTS
FIFTH are those faults that have been detected by the ECU but are considered minor
in nature. These faults will only be displayed on the caution/warning panel the next
time the throttle is placed in the cutoff position. The lights used to display these
faults are FADEC DEGRADED, FADEC FAULT and RESTART FAULT.
These faults can be investigated by using the procedure in the Bell 407 Manufacturers
Data.
FADEC FAULT DEFINITIONS
FADEC FAIL
When a fault occurs that causes a hard fault in the FADEC system the FADEC/ECU
will activate the DRTM system and the FADEC FAIL light and FADEC FAIL HORN
will be turned on.
The pilot will assume control of the throttle by increasing or decreasing PLA from
the FLY detent position as required to maintain normal rotor RPM.
FADEC DEGRADED
When a fault occurs that may cause a degradation of the performance of the FADEC
system (for example a loss of input from the collective position transducer) the FADEC
DEGRADED light will be turned on to caution the pilot. The pilot should follow the
appropriate procedures as set out in the Flight Manual.
FADEC FAULT
When a fault occurs that the ECU determines is a loss of some feature of the FADEC
system, but this error does not cause a degradation of performance, the FADEC
FAULT light will be turned on to caution the pilot. The pilot should follow the appropriate
procedures as set out in the Flight Manual.
FADEC MANUAL
In normal operation the DTRM system will revert to MANUAL MODE without any
action by the pilot without regard to the position the FADEC MODE switch. If the
pilot positions the FADEC MODE switch to MAN, the FADEC MANUAL light will be
turned on and the FADEC MODE switch will show MAN.
FADEC FAIL HORN
FADEC FAIL Horn (tone) comes on whenever FADEC FAIL light is illuminated. Horn
may be muted by positioning FADEC MODE switch to MAN.
RESTART FAULT
If an error occurs in the FADEC system that may result in a problem starting the
helicopter at the next start the RESTART FAULT light will be turned on. The purpose
is to give the pilot advance notice of a potential starting problem and allow him to
plan the balance of his flight and shutdown appropriately.
PRODUCT DATA
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CHECKING FADEC FAULT CODES
With the throttle in the closed (cutoff) position the fault codes displayed will be based
on faults present at the last engine shutdown (historical faults). With the throttle out
of cutoff or at idle position the fault codes displayed will be based on the current
faults only.
Faults shown as either current or historical are to be investigated per the procedures
outlined in the Bell 407 Manufacturers Data. This will ensure the pilot can be made
aware of the specific cause of the fault displayed and take appropriate action prior
to further flight operations.
HISTORICAL FAULT CODES [PROCEDURE FOR VIEWING]
Engine must be shutdown and the FADEC MODE switch must be in MAN. Place
throttle in the cutoff position and the collective lower than 10%.
If a historical fault exists the FADEC DEGRADED light will be ON. Depress and
release FADEC/ECU maintenance button on the left hand side of the lower pedestal
to start the fault code reporting mode.
FADEC DEGRADED, FADEC FAULT and the RESTART FAULT lights will flash 5
times to indicate that maintenance mode has been entered by the ECU. The three
lights will then turn off.
Depress and release FADEC/ECU maintenance button. Faults will be displayed via
a number of FADEC DEGRADED caution panel light segment
flashes. Depress and release FADEC/ECU maintenance button to flash the next
fault code.
A steady FADEC DEGRADED caution panel light illumination indicates that no other
faults exist for this light.
Continue to depress and release the FADEC/ECU maintenance button to step through
the FADEC FAULT and RESTART FAULT caution panel light segments as above to
reveal fault codes. If no fault code is to be displayed by the selected caution panel
light segment, the caution light (FADEC DEGRADED, FADEC FAULT and RESTART
FAULT) will be ON steady when the FADEC/ECU maintenance button is released.
Once the 3 FADEC caution lights are ON steady, the next push of the FADEC
maintenance button will cause the FADEC DEGRADED, FADEC FAULT and RESTART
FAULT caution light segments to flash 5 times and then extinguish, indicating that
the FADEC has exited the maintenance mode.
If the throttle or collective are moved during the above procedure or the FADEC
MODE switch is positioned to AUTO, the FADEC/ECU will exit the maintenance mode.
CURRENT FAULT CODES [PROCEDURE FOR VIEWING]
The procedure for checking for current faults is the same as that outlined above
except that the engine must be shutdown, FADEC MODE switch must be in MAN,
collective lower than 10% and the throttle must be in idle position.
PRODUCT DATA
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ENGINE MOUNTED ACCESSORIES
ENGINE FUEL PUMP
The dual element fuel pump is integral to the Hydromechanical Unit [HMU] and is
mounted on the right rear accessory gearbox drive pad. The drive shaft includes an
seal runner for the carbon ring seal. Pump RPM is 4205 at 100% NG. Output volume
is approximately 4.5 GPM. Fuel first reaches a centrifugal “liquid ring” boost pump.
Boost pump output, called PBF (pressure before filter), flows out to the gearbox
mounted filter. It returns to the inlet of the high pressure gear pump as PAF. Gear
pump high pressure output (PF) is directed to the metering valve, the pressure
differential (metering head) regulator valve, and the manual control components. A
relief valve limits pump output to 900 PSI above inlet pressure.
ENGINE MOUNTED FUEL FILTER
The fuel filter is combined with the oil filter into a single assembly. The assembly
is remotely mounted to the lower left rear side of the gearbox. The disposable, noncleanable fuel element is rated at 25 microns. A bypass valve across the element is
set at 3 to 3.4 psid. An impending bypass indicator will visually extend from an assembly
attached to the filter assembly housing at 2.1 to 2.9 psid.
PRODUCT DATA
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SPEED PICKUP SENSORS
The NG and NP speed pickups are the
same as existing series IV designs, except
with new connectors. Both sensors are
located on top of the gearbox housing.
PERMANENT MAGNET
ALTERNATOR
The PMA is driven by the power turbine
gear train at 4410 RPM on the C47 engine.
It is mounted on the left rear drive pad. It
provides the primary electrical power to
the ECU above 85% NP.
COMPRESSOR INLET AIR
TEMPERATURE SENSOR T1
The T1 sensor is mounted in the airframe
inlet plenum. It is mounted so as to provide
the best possible correlation to the actual
engine air temperature, while providing
for suitable protection against foreign
object damage and ice build up.
TORQUE METER OIL
PRESSURE SENSOR
The sensor signal is used for torque
limiting. It is tied in at the current TMOP
location provided for aircraft torque
sensing on the accessory gearbox.
IGNITION EXCITER
The ignition exciter is a solid-state, high
energy ignition unit. The unit is designed
to have a constant spark rate of 300 to
330 sparks per minute throughout its
operational temperature and voltage
range.
The exciter is energized when either :
•the PLA is greater than 9% and NG is
between 5% and 50%
•the anti-ice is selected
•a flameout is detected
PRODUCT DATA
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ENGINE OIL SYSTEM
The engine oil system is a dry-sump type lubrication-system. The engine oil system
includes an externally mounted oil tank, a temperature bulb, a manually operated
drain valve, and an oil cooler. The oil cooler is installed on top of the fuselage, behind
the aft firewall. The lubrication oil is supplied to and from the engine through rigid
lines and flexible hoses.
The oil tank supplies lubricating oil through the oil inlet tube to a gear-type pressure
pump installed on the aft face of the accessory gear box. Oil under pressure (pressure
oil) goes through the oil filter and a pressure regulating valve, through the oil-pressure
sensing port, to the four lubricating ports and the torquemeter pressure port. This
oil lubricates the engine. The pressure oil becomes scavenge oil and returns to the
three scavenge pumps and becomes scavenge return oil. The scavenge return oil
goes through the chip detector, through the oil return hose, through the external oil
filter, and through the oil return tube to the oil cooler. The scavenge return oil passes
through the oil cooler, through the oil-tank return tube to the oil tank. The engine-oil
pressure indication is supplied by the oil pressure transducer. The engine-torque
pressure indication is supplied by the engine torque transducer. Both of these
components are installed on the engine.
The 250-C47B lubrication system is basically a C30 system with the following additions.
A torquemeter transducer, an engine furnished and mounted scavenge oil filter,
and quick disconnect-fuzz buster type magnetic chip detectors.
The accessory case mounted transducer converts torquemeter oil pressure into and
electrical signal. This signal is required by the Electronic Control Unit (ECU) for
torque limiting.
OIL COOLER
ENGINE OIL TANK
COOLER BLOWER HOUSING
PRODUCT DATA
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ENGINE OIL SYSTEM SCHEMATIC
ENGINE OIL FILTER
The scavenge oil filter is now engine furnished and mounted on the accessory case,
together with the fuel filter. The filter element is disposable, non-cleanable, and
provides 3 mircron filtration. A bypass valve opens when the pressure differential
across the filter exceeds 10.8 to 13.2 psid. The valve reseats at 7 psid. A bypass
indicator button extends from the end of the filter bowl when the pressure differential
exceeds 8.8 to 10.8 psid at 126°F. A thermal lockout prevents button actuation
below 110° F ± 15°F. The button is reset manually.
PRODUCT DATA
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OIL TANK
The oil tank has a normal fluid capacity
of 1.5 gallons (5.7L). You can see the
oil level through the sight glass or measure
the oil level with the dipstick. The dipstick
is attached to the cap and adapter
assembly. The oil tank includes ports for
the vent tube, the oil inlet line, the oil
scavenge line, the oil temperature bulb,
and the oil tank manual drain valve.
OIL LEVEL SIGHT GLASS
OIL COOLER
The oil cooler is installed on the transition
duct. This duct is installed on the oil cooler
blower motor between the aft firewall and
the forward face of the engine oil tank.
The oil cooler is a transmission oil cooler
and an engine oil cooler. The tail rotor
driveshaft drives the fan that supplies and
moves the air through the oil cooler. The
oil cooler is a dual-element heat
exchanger. One heat exchanger cools the
engine oil while the other cools the
transmission oil.
ENGINE OIL
RESERVOIR
The oil cooler has a temperatureregulating/pressure relief valve. This
valve regulates the engine oil
temperature. The valve starts to close at
160° F (71°C) and is fully closed at 178°F
(81°C). If the system pressure goes above
2 5 pounds per square inch (PSI) the
pressure relief part of the valve opens the
valve fully.
PRODUCT DATA
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FUEL SYSTEM
Standard fuel capacity is 126 US gallons, 477 liters. A single 19 US gallon, 72liter
auxiliary fuel tank is available as an option. It installs directly behind the main tank
in the baggage compartment. The fuel system consists of two interconnected bladder
type fuel cells made of rubberized fabric. Use of two tanks reduces the center of
gravity shift resulting from fuel burn to increase stability. The forward cell is located
under and between the aft facing passenger seats. The main cell is under and behind
the rear seat bench. Both cells are made of a heavy rupture resistant material and
have passed the 50 foot free fall test. The flexible, stretchable interconnecting fuel
lines make the whole system very crash resistant. A dual, electric, submerged fuel
boost pump provides fuel under pressure to the engine pump and provides volume
and pressure flow from the forward cells to the aft cell. Fuel will be transferred even
if one of the pump elements fails. Fuel management is automatic. The single forward
tank has a capacity of approximately 35 US gallons, 133 liters, of fuel. The aft
(main) tank has a capacity of 91 US gallons, 344 liters. The Bell 407 is refueled
from one location on the right side of the helicopter just aft of the passenger door.
There is a grounding plug nearby.
FORWARD FUEL CELL
MAIN [AFT] FUEL CELL
FUEL FILLER CAP
The fuel filler cap is attached to the port by a hinge to prevent it's loss. The lift and
turn locking mechanism incorporates a raised lever which reduces the effort required
to open the cap. The Bell 407 is designed for refuelling at the maximum flow rate of
114 liters per minute (30 U.S. Gallons per minute).
HINGE
PRODUCT DATA
LEVER
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FUEL SYSTEM COMPONENTS
AIRFRAME FUEL FILTER
FUEL SHUTOFF VALVE
FUEL FILLER
VENT
LINES
FUEL QUANTITY PROBES
AUXILIARY
FUEL CELL
PROVISIONS
MAIN TANK
INTERCONNECT
LINE
FUEL BOOST PUMP
FORWARD TANK
FUEL TRANSFER PUMP
PRODUCT DATA
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TRANSFER SYSTEM
Fuel is transferred between the aft and forward cells by two means: a gravity transfer
line permits fuel to transfer from the aft to forward cell whenever the level of fuel in
the aft cell is above the height of the standpipe, and a pressure transfer line from the
forward transfer pumps to the aft (Main) tank. The pressure transfer system has two
electrical pumps mounted on a common sump plate in the bottom of the forward
cell, they each have a check valve and a pressure switch. The transfer pumps run
continuously until the forward tank is empty. The pumps will then run an additional
2½ minutes then shut off automatically.
BOOST OR TRANSFER PUMP
FUEL BOOST PUMPS
The two electrical fuel boost pumps, which
are part of the sump assembly and are
located in the bottom of the aft fuel cell.
Each pump has a pressure switch
connected to the caution panel to indicate
boost pump operation. The boost pumps
are connected to a Tee fitting that
connects the supply from each pump to a
single line. The boost pumps run
continuously, taking fuel from the aft cell
and supplying it to the engine via the check
valves, hose assemblies, shutoff valve
and filter to the engine. The pump
elements are cartridge assemblies which
may be replaced easily without defuelling
the aircraft.
DUAL
ELEMENTS
REPLACEABLE
PUMP
CARTRIDGE
FUEL SUMP DRAIN SWITCHES
FUEL TRANSFER PUMPS
An identical sump plate with two electric
pumps is located in the forward fuel tank
and transfers fuel to the aft tank.
FUEL SUMP DRAIN VALVES
The sump drain valves are located in the
fuel boost and transfer pump sumps, and
it is electrically operated by means of the
drain switches located aft of the right-side
passenger door. To draw a fuel sample,
both fuel boost and transfer pump switch
circuit breakers should be off, the battery
switch should be in the ON position, and
the fuel valve should be in the OFF
position.
PRODUCT DATA
FUEL CELL VENT, ENGINE, TRANSMISSION,
AND FUEL FILTER DRAIN
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FUEL SYSTEM SCHEMATIC
FUEL SYSTEM INDICATORS
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FUEL QUANTITY PROBES
Three fuel quantity probes are used to monitor the fuel system. The forward probe
is electrically connected to the two probes in the aft cell. Readings from all three
probes are transmitted to a fuel quantity indicator on the instrument panel. The forward
fuel quantity switch in the cockpit, when activated, will provide a momentary reading
only on the fuel remaining in the forward cell.
LOW LEVEL FUEL WARNING SWITCH
The low level fuel warning switch is mounted through the bottom of the fuel boost
pump assembly located in the aft fuel cell and is electrically connected to the caution
panel. The caution light will illuminate when approximately 110±15 pounds of fuel
remain in the aft cell.
AIRFRAME FUEL FILTER
The airframe mounted fuel filter is located on the forward firewall on the right side of
the power plant. It is equipped with built in filter bypass valve and impending bypass
switch that is electrically connected to a segment in the caution panel. Also, a red
press to test button is located on the top of the filter to check the electrical connection
to the fuel filter caution light. A manual filter drain valve is at the bottom of the
filter.
SHUTOFF VALVE AND TRANSDUCER
AIRFRAME FUEL FILTER
FUEL SHUTOFF VALVE AND FUEL PRESSURE TRANSDUCER
The fuel pressure transducer provides a
means of monitoring fuel pressure in the
system. The fuel shutoff valve is located
above the aft fuel cell servicing port and
is electrically operated by a guarded
(covered) switch on the instrument panel.
FUEL VALVE SWITCH
FUEL REQUIREMENTS
Fuel conforming to these specifications may be used at all ambient temperatures:
ASTM-D-1655, Type B; MIL-T-5624, Grade JP-4 (NATO F-40).
Refer to the Approved Rotorcraft Flight Manual for other acceptable fuel specifications.
PRODUCT DATA
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ELECTRICAL SYSTEM
The electrical system of the Bell 407 is very similar to that of the Bell 206L-4 LongRanger
IV. The changes consist of those necessary to use the LCD instruments, including
the addition of sensors and wiring for the electronic gauges which replace the "wet"
gauges (Torque & Oil Pressure). Two new additions are an avionics master switch
to speed up the start and shutdown processes and a switch to turn off the fuel
boost pumps without having to pull the circuit breaker. Controls for the 28 volt direct
current electrical system are located on the overhead console and instrument panel.
Power is supplied by a nickel-cadmium battery and a combination starter/generator.
Control relays, power relays, voltage regulators, and circuit breakers control and
regulate the voltage transfer. Malfunction monitoring circuits are located in individual
compartments.
DC POWER SYSTEM
The dc power systems include the battery, external power, starter/generator, and
starter/igniter systems.
COCKPIT OVERHEAD PANEL
BATTERY SYSTEM
BATTERY COMPARTMENT
[SHOWN WITH OPTIONAL HD 28 AMP-HR BATTERY]
The battery system includes the battery, battery relay, battery switch, and related
wiring. The Bell 407 has a standard 24 volt, 17 ampere-hour nickel cadmium battery.
The standard battery has demonstrated starts down to -29°C(-20°F) ambient
temperature. For extreme cold weather operations an optional 28 ampere-hour battery
is available, and has demonstrated starts down to -35°C(-31°F) ambient temperature.
The battery relay, located in the nose section forward of battery, is an electrically
operated switch controlling battery current to the main bus bar. It is actuated by a
battery switch, located in the overhead console, which opens and closes the circuit
to the relay energizing coil.
PRODUCT DATA
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ELECTRICAL SYSTEM SCHEMATIC
PRODUCT DATA
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CONTROL PANELS AND CIRCUIT BREAKERS
Controls and control panels are mounted in the overhead console and instrument
panel. The circuit breakers are mounted on the overhead console. Circuits can be
separated and closed by operating these push/pull circuit breakers
OVERHEAD CIRCUIT BREAKER AND SWITCH PANEL
FADEC HORNS
LOW ROTOR
RPM HORN
GENERATOR SYSTEM
The generator system includes the generator portion of the starter/generator, generator
control unit, line control relay, generator reset switch, and generator shunt.
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The generator, located on the underside of the engine to the right of the helicopter
centerline, furnishes power at a regulated voltage for all dc electrical components
on the helicopter. The generator switches onto the main bus when the generated
voltage exceeds the voltage on the bus by 0.30 to 0.42 volts.
A generator control unit is the heart of the electrical system. It operates the line
control relay, regulates generator voltage, and provides protection against overvoltage
and reverse current. It contains an electronic voltage regulator to control the voltage
output of the dc generator and a circuit to energize the line contactor when correct
conditions exist.
The line control relay is installed on the electrical panel assembly located on the
equipment shelf above the baggage compartment. It opens or closes the power circuit
between the dc bus and generator, and it is controlled by the presence or absence
of a proper output voltage from the generator control unit.
The generator reset switch, located on the overhead console, is a double-pole,
triple-throw, spring-loaded switch with only momentary contact in the RESET position.
It completes the generator field circuit in the ON position, supplies voltage to reset
the generator field in the RESET position, and disconnects generator field circuit in
the center OFF position.
The generator shunt is installed on the electrical panel assembly on the equipment
shelf above the baggage compartment and provides a voltage drop proportional to
the current in order to operate the loadmeter.
STARTER/GENERATOR
EXTERNAL POWER DOOR
EXTERNAL POWER SYSTEM
The external power system includes the external power receptacle, external power
relay, and related wiring. (Maximum current limit of the external power source should
not exceed 500 amperes.)
The external power receptacle, located in front center at the nose section, is a
polarized receptacle used as a contact point for external power plug in.
The external power relay, located in the nose below the battery in the landing light
compartment, is an electrically operated switch between the external power receptacle
and main bus bar. It is controlled through the small positive pin from the external
power source which energizes the circuit to the activating coil of the relay.
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STARTER/IGNITER SYSTEM
The starter/igniter system includes the
starter portion of the starter/generator,
starter relay, field/igniter relay, igniter,
and starter switch.
The starter/generator, located on the
underside of the engine, is energized by
the starter relay to start the engine.
T h e s t a r t e r r e l a y, i n s t a l l e d o n t h e
electrical panel assembly located on the
electrical equipment shelf above the
baggage compartment, supplies direct
current to the starter when the starter
switch is depressed.
The field/igniter relay, installed on the
electrical panel assembly located on the
electrical shelf above the baggage
compartment, is actuated when the start
switch is depressed; and it supplies direct
current to the igniter and start field
suppress section of the generator control
unit.
The igniter , furnished with the power
turbine and located below the unused
tachometer generator pad on the lower
left section of engine, consists of a tension
capacitor discharge ignition exciter which
provides a continuous arc during engine
start cycle.
STARTER/GENERATOR
START SWITCH
The starter switch, located in the
collective stick switch box, is a single-pole,
single- throw, spring-loaded pushbutton
type switch. When the switch is pressed
to START position, a ground potential is
furnished to energize the starter and field/
igniter relays, which activates the
automatic FADEC controlled start
sequence.
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LIGHTING SYSTEMS
The lighting system includes both interior and exterior lighting. The interior lighting
system includes the cockpit light, instrument and control panel lighting, and the
cabin lighting. The exterior lighting system includes the position, anticollision, and
landing lights.
INTERIOR LIGHTING SYSTEM
The cockpit light, utilizing a narrow spotlight or wide floodlight beam, is a multipurpose
utility light designed to provide white illumination. The light is protected by a 5 ampere
COCKPIT LIGHTS circuit breaker. Controls for ON/ OFF, DIM/BRIGHT, and SPOT/
FLOOD are incorporated into the cockpit light body. Instrument and control panel 5
Vdc and 28 Vdc power is supplied through the INSTR lights circuit breaker located
on the overhead console. Light intensity is adjusted by the INST LT rheostat knob
also located on the overhead console. Rotation of this knob operates the power ON/
OFF SWITCH which provides power to both the 28 and 5 Vdc lighting systems. Two
individual white cabin lights provide passenger area illumination at pilot’s discretion.
These lights are protected by the 5 ampere COCKPIT LIGHTS circuit breaker. They
are regulated by the CABIN LT switch and by individual switches located near each
light. When the CABIN LT switch is placed in the CABIN LT position, both cabin
lights illuminate. When it is placed in PASS position, each light can be individually
controlled by the associated passenger switch.
LIGHTING
CIRCUIT BREAKERS
COCKPIT LIGHT
BRIGHT/DIM SWITCH
INST LT
RHEOSTAT
POSITION
LIGHTS
CAUTION
ANTIBRIGHT/DIM
COLLISION
LIGHT
CABIN LT
ON/OFF/PASS
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With the INST LT rheostat ON, the caution
panel lights can be dimly or brightly
illuminated by positioning the bright/dim
switch momentarily to the BRIGHT or DIM
position. The caution lights will remain at
the selected intensity until the INST LT
rheostat is turned off or another position
is selected. However, when the INST LT
switch is turned ON and the caution light
selector is positioned to DIM, the caution
lights are dimmed to a fixed intensity and
cannot be adjusted by the INST LT switch.
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EXTERIOR LIGHTING SYSTEM AND POSITION LIGHTS
The exterior lighting system includes the position, anticollision, landing lights,
and related wiring and components. The position lights include the upper horizontal
stabilizer left and right side, and lower cabin left and right side position lights, and
the tail light. The anticollision light is mounted on the top portion of the vertical fin.
The dual landing lights are located in the lower forward portion of the helicopter
nose section.
FORWARD POSITION LIGHTS
LANDING LIGHTS
AFT POSITION LIGHTS
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ANTICOLLISION LIGHT
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TAIL LIGHT
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JACKING
Three jack points are provided for the purpose of leveling the helicopter for weight
and balance measurement, and other maintenance functions. Two are located on
the underside of the forward fuselage at FS 55.14, and left and right BL's 16.82, with
the third located on the aircraft centerline at FS 204.92.
LEVELING PLATE
LEVELING
A level plate is located on the baggage bay floor at fuselage station 170.1. A slotted
plate for a plumb bob is located just above the level plate on the underside of the
engine pan.
LIFTING
Lifting the helicopter may be accomplished by attachment of a CLEVIS to the main
rotor retention nut, or by use of nylon web SLINGS through the main rotor hub.
Procedures are described in the 407 Maintenance Manual.
CLEVIS LIFTING
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SLING LIFTING
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COVERS AND TIEDOWNS
Protective covers and tiedowns are furnished as loose equipment and are used for
the parking or mooring of the helicopter. Additional equipment such as straps ,
cables, clevises, ramp tiedowns will be required during mooring.
PARKING AND MOORING
Park helicopter for normal and turbulent conditions with winds up to 50 knots in
accordance with procedures described in the 407 Maintenance Manual.
If helicopter must be parked in the open during period of high wind forecast, comply
with the precautionary measures described in the 407 Maintenance Manual.
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MODEL 407 EXTERNAL NOISE LEVEL
The Bell 407 is certified as a Stage 2 helicopter as prescribed in FAR Part 36, Subpart
H, for gross weights up to and including the certificated maximum takeoff and landing
weight of 5000 pounds (2268 Kilograms). There are no operating limitations to
meet the noise level requirements.
The following noise level complies with FAR Part 36, Appendix J, Stage 2 noise
level requirements. It was obtained by analysis of approved noise tests conducted
under the provisions of FAR 36, Amendment 36-20.
The certificated noise level for the Bell 407 is 85.1 dBA SEL.
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SECTION TWO
OPTIONAL EQUIPMENT
BHTC FACTORY INSTALLED
& VENDOR STC KITS
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INTRODUCTION
Bell maintains a wide variety of factory installed standard kits for the 407. Many of
these kits are also suitable for installation at the customer's facility. Bell also provides
factory installation of selected STC approved VENDOR kits.
LOW SKID FAIRINGS / CARGO HOOK
FLIGHT INSTRUMENT GROUP
HEAVY DUTY BATTERY
DUAL CONTROLS
HIGH SKID GEAR
EMERGENCY FLOTATION GEAR
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TYPICAL AVIONICS KITS LOCATIONS
In the illustration below, some of the factory available avionics kits are shown in the
ELT TEST
normal installed location;
/ON SWITCH
HSI [COMPASS
SYSTEM KCS-55]
NAV SOURCE
SELECT
GPS ALERT
LIGHTS
COURSE
DEVIATION
INDICATOR
RMI [ADF
INDICATOR]
GPS
KLN-89B
KMA-24H ICS
COMM/NAV #1
KCS 155 OR165
COMM#2
KY-196
TRANSPONDER
KTR-70 OR 79
ADF
KR-87
COMPASS
SLAVING SWITCH
KCS-55
AFT CABIN ICS
CONTROL SWITCH
OPTIONAL
ASHTRAY
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HIGH SKIDS
Bell Kit No. 407-706-007-xxx
The high skid landing gear is an optional kit which provides greater ground clearance
to facilitate operations in rough terrain, tall grass, and other adverse conditions. It
also provides more space to carry underslung loads. The gear is similar in design to
the standard skids with replaceable skid shoes, etc. but, it provides approximately
22.2 centimeters (8.75 inches) of additional ground clearance. The kit consists of
fore and aft cross tubes, skid tubes, and the attaching hardware.
Included with the kit are Aeronautical Accessories Flight Steps STC # 407-326001. A towing ring is fitted on the inside of each skid forward of the forward crosstube.
AAC ®FLIGHTSTEP
INCREASED GROUND CLEARANCE
The high skids use the same ground handling wheels as the standard gear.
Aerodynamic fairings are not available for this kit.
The kit fits the standard attachment points and can be installed and removed with
ease.
The high skids are qualified for the standard 2268 kilograms (5000 pounds) and
alternate 2381 kilograms (5250 pounds) maximum gross weights within the limits
outlined in the Flight Manual Supplement BHT-407-FMS-2.
The complete high skid assembly adds 14.6 kilograms (32.1 pounds) to the standard
configuration weight.
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EMERGENCY FLOATS
Bell Kit No. 407-706-008-xxx
The emergency floats are available when operations are over land and water, but
water landings are not normally anticipated. The kit consists of high skids, emergency
floats, inflation system using compressed nitrogen, electrical solenoid valves to
activate the system, attaching hardware, and a preflight test feature. The six floats
are made of nylon. The two forward floats each have a single cell. The remaining
four each have two cells.
Two cylindrical nitrogen bottles at are aligned fore and aft, under the fuselage
immediately behind the forward cross tube of the landing gear. The cross tubes fit
the standard gear attachment points enabling the kit to be installed and removed
with ease. The emergency floats are qualified for the standard 2268 kilograms
(5000 pounds) and alternate 2381 kilograms (5250 pounds) maximum gross weights
within the limits outlined in the Flight Manual Supplement BHT-407-FMS-1.
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FLOAT BOTTLE
[ONE OF TWO]
AAC ®FLOAT STEP
FLOAT BAGS
[STOWED]
SOLENOID VALVE
[ONE OF TWO]
ARM &
INFLATE
SWITCHES
The floats are capable of supporting the
maximum gross weight of the helicopter
four seconds after actuation. Inflation time
is approximately 5 seconds. When in the
water with the float system inflated, the
helicopter should generally be stable up
to a condition of sea state 5.
The flight restrictions stowed and after
inflation are listed in the Flight Manual
Supplement.
The standard 407 already has two sets
of position lights. This prevents any
change of blockage of lights when the
floats are installed and inflated. A Float
Test/Float Arm switch is on the collective
and two indicator lights are mounted on
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the instrument panel. The lights indicate
circuit continuity during a test and when
the arming circuit is closed. The Float Test
Caution Light does not dim. Once armed,
the pressure system is electrically
activated from a guarded button on the
pilot's collective stick. Operation
instructions are in the Flight Manual
Supplement.
The standard ground handling wheels are
used.
For flight with doors on and off, the
restrictions are the same for the standard
configuration.
The system adds 113.0 kilograms (249.1
pounds).
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LANDING GEAR FAIRINGS [LOW SKID GEAR ONLY]
Bell Kit No. 407-706-010-xxx
The airfoil type landing gear fairings go over the forward and aft crosstubes of the
standard low skid gear. They are made of formed thermoplastic and have aluminum
alloy supports. Additional aerodynamic covers fit over the skid attachment fittings
on the forward cross tube.
Fairings are not available for the high skid gear or emergency flotation gear.
When installed, the fairings increase airspeed approximately 3 knots (5 Kilometers
per hour).
The fairings add 5.4 kilograms (11.9 pounds).
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AUXILIARY FUEL TANK
(19 U. S. Gallons / 71.9 liters)
Bell Kit No. 407-706-011-xxy[EQUIPMENT]
Bell Kit No. 407-706-011-xxx[ PROVISIONS]
The auxiliary fuel tank consists of two kits, the provisions, and the tank. When the
kit is installed the total fuel capacity of the 407 is increased to 553 liters (146.9 U.
S. Gallons). When installed there in no requirement for additional fuel management
by the pilot. Filling of the tank is through the regular fuel fill port, and likewise the
fuel automatically flows to the aircraft fuel system. The fuel quantity indicating system
is modified to correctly show the new fuel quantity. The new full fuel indication on
the fuel gauge will be approximately 1005 pounds. The provisions kit adds 2.3
kilograms (5.0 pounds), and the tank adds 13.4 kilograms (29.6 pounds) to the
standard configuration.
AUX. FUEL
TANK MOUNT
PROVISIONS
BAGGAGE COMPARTMENT WITH OPTIONAL
AUXILIARY FUEL TANK
AUX. FUEL
LINE PROV.
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DUAL CONTROLS
Bell Kit No. 407-706-702-xxx
The dual controls kit provides dual flight control capability either for operations
requiring a two-man flight crew or for pilot training.
The kit adds cyclic and collective sticks and anti-torque pedals for the left front
seat. Similar to the main controls in appearance and relative position, they provide
all the major functions as well as identical control "feel."
The collective stick includes a fully functioning twist-trip throttle control. The directional
pedals can be adjusted for leg length. Friction controls are not duplicated.
All 407 airframes have provisions for the dual controls kit. With the dual controls
installed, there is no change in the number of usable seats.
Since the instrument panel, radio pedestal, overhead console, fire extinguisher,
cockpit light, etc., are all centrally located and within easy sight and reach of both
crew positions, there is no need to relocate anything in the cockpit.
The quick disconnect feature means that the cyclic and collective sticks are rapidly
removable without tools. The installation of the litter kit requires the removal of the
dual controls. The kit adds 5.6 kilograms (12.3 pounds).
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AFT AUDIO ICS
Bell Kit No. 407-705-003-xxx[3-station]
Bell Kit No. 407-705-003-xxy[5-station]
The Aft Audio ICS Kits provide communication between crew and passengers. It
consists of either 3 or 5 rear cabin mounted David Clark Headsets and control cable
assemblies. The 3 station kit provides three headsets for the aft row of cabin seats,
and the 5 station kit adds two more stations for the rear facing cabin seats.
This enables the pilot to communicate through the KMA 24H Intercom Control and
permits the passenger using the headsets to converse with each other and with the
pilot(s). A control switch on the radio pedestal allows the pilot to isolate the cockpit
headsets from the passenger conversation, while allowing the passengers to continue
their conversation.
The 3 station kit adds 3.9 kilograms (8.6 pounds). The 5 station kit adds 6.1
kilograms (13.5 pounds).
5-STATION ICS
ADDS TWO REAR FACING SEATS
3-STATION ICS
AFT ROW OF SEATS
EACH STATION HAS A DAVID CLARK HEADSET & CONTROL CORD
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VHF COMMUNICATION/NAVIGATION TRANSCEIVER
King KX 155
Bell Kit No. 407-705-001-xxy[Transceiver]
Bell Kit No. 407-705-001-xxx[Provisions]
The KX 155 Communications/Navigation Transceiver provides:
•VHF two-way voice communications, 720 channels in the 118 to 135.975
megahertz range.
•Intercommunication between pilot and copilot.
•200 navigation channels in the 108 to 117.95 megahertz range.
•Channeling for distance measuring equipment.
The KX 155 has simultaneous digital display of one active and one standby frequency
(either communication or navigation or both). A push button control switches to
either one.
The Transceiver Kit includes the transceiver, two David Clark headsets, each
with an adequate length of coiled cable and built-in microphone, a vendor manual
and warranty, and necessary hardware to complete the installation.
The transceiver is part of the KING Silver Crown series. It is crystal controlled, solid
state, and panel-mounted (159X52X258mm; 6.25X2.05X10.2 inches). Internal lighting
is blue/white.
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KMA-24H ICS CONTROL PANEL
For additional Capability a second KX 155 Transceiver can be added. Although it
will cover the same frequency range, both can be operated simultaneously on different
channels.
When a second KX 155 and a second indicator are installed, a second source of
navigational information is provided.
Bell or King service center installation is required in order to obtain the King full
year warranty.
The Provisions Kit, No. 407-705-001-101, includes a King KMA 24H audio panel,
two Quam-Nichols speakers in the forward cabin, and the antenna for the KX 155.
Hardware includes all the cables, mounting brackets, jacks and circuit breakers to
complete the installation for either single or dual flight controls. The panel is backlit with blue/white lighting.
The transmit and ICS switches on the cyclic stick grip are connected. If dual controls
are installed, the transmit and ICS switches on the copilot's cyclic grip are connected
and the copilot's intercommunication foot switch is installed.
The audio panel is mounted above the transceiver, on the radio pedestal. The
audio panel's function is to manage the flow of radio communications to and from
the helicopter when there is more than one radio in the system.
The pilot can select which transmitter is active and which combination of receivers
will be directed to headsets, and to cabin speakers or passenger headset, if installed.
The King KMA 24H Audio Panel enables the audio management of up to three
transceivers and five receivers. These can be individually switches on and directed
to headphones or speakers by push buttons.
A rotary switch allows the pilot to transmit on the selected communication transceiver,
Comm 1, Comm 2, Comm 3, Comm 4, public address system, or external speakers,
and emergency.
The integral intercom with its own volume control can coordinate up to five microphone
inputs and has adequate power to handle up to six headphones connected in
parallel."Pilot priority" has been incorporated into the mike switching. If both the
pilot's and copilot's mikes are keyed simultaneously, the pilot's mike is automatically
given transmit priority. This authority can be designated, by internal adjustment, to
either the pilot or copilot position to accommodate flying training needs.
The AUTO switch automatically links the speaker or headphones to the selected
transmitter. When AUTO is being used, it can be left in the OFF position.
The complete installation adds 7.8 kilograms (17.2 pounds). Items removed to
permit installation are shipped with the helicopter.
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AUTOMATIC DIRECTION FINDER
King KR 87 and KI 227
Bell Kit No. 407-705-001-xyy
The KR 87 ADF receiver with the KI 227 indicator can provide the relative bearing
to a selected transmitting station which is within the 200 to 1699 kilohertz range.
The needle on the indicator shows the direction of the radio transmitter relative to
the aircraft's heading. The KR 87 displays the active ADF frequency on the left.
The KI 227 has a manually rotatable compass card.
KI-227 RMI
KR-87 ADF
An integral timer provides display of flight and elapsed time and can be programmed
to count up or down for timed approaches, trip leg timing or fuel management.
The KR 87/KI 227 combination is a part of the King Silver Crown Series.
The kit includes the receiver, 1 76mm (3 inch) indicator, the loop/sense antenna,
the wiring and hardware required to complete the installation, a vendor manual and
the warranty.
The receiver is crystal controlled with digital tuning for positive, easy operation. It
is solid-state, self contained, radio pedestal mounted [159X33X257mm (
6.25X1.35X10.13 inches)]. Once the volume control is set, audio reception volume
is held relatively constant over side variations of signal input. Both the indicator and
the receiver are backlit, blue/white lighted, with easy to read numbers. Both meet
FAA's Technical Standard Orders.
Bell or King Service Center installation is required in order to obtain KING's full
year warranty.
The kit adds 2.9 kilograms (6.3 pounds). Items removed to permit installation are
shipped with the helicopter.
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TRANSPONDERS
Bell Kit No. 407-705-001-yyy[King KT-76A]
Bell Kit No. 407-705-001-yyz[KT-76A Provisions]
Bell Kit No. 407-705-001-zzy[King KT-70]
Bell Kit No. 407-705-001-zzx[KT-70 Provisions]
The King KT-76A and KT-70 transponders enable the air traffic controller to identify
your helicopter clearly, quickly and accurately. The response to interrogation is
automatic. Additional identification can be requested in which case the reply is activated
by pressing the ident push-button.
The Provisions Kits, include the antenna and the wiring, bracket and hardware
necessary to complete the installation.
Items removed to permit installation are shipped with the helicopter.
The Transponder Kits, includes the transponder, a vendor manual and warranty.
Both the KT-76A and KT-70 are Class I radio transmitter/receivers operating on
radar frequencies and have adequate power for use at all operating altitudes.
The sets receive ground radar interrogations at 1030 MHz and these trigger a coded
response of radar pulses at 1090 MHz, which is transmitted back to the ground radar.
KT-76A
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KT-70
They respond in any of 4096 preselected "codes." Codes differ only in number of
pulses transmitted, not in frequency. Transmitter frequency remains always at 1090
MHz. The assigned code, or emergency codes, are easily set and appear as an
illuminated digital readout.
The transponder's coded reply pulses reinforce the normal "blip" which appears on
air traffic control radar. When the IDENT button is pressed, the blip will flash or
"bloom" so the controller can positively identify the aircraft and its location.
When connected to an OPTIONAL encoding altimeter or blind encoder, both also
transmit the aircraft altitude in flight level increments (three digits, representing
even 100 foot increments referenced to 29.92 in. of Hg.).
The primary difference between the KT-76A and the KT-70 is the code display and
setting controls. The KT-76A has a mechanical display and individual rotating knobs
which manually set each of the four numbers. The KT-70 has an LED code display
with individual rotating knobs which sets each code number in turn. Additionally,
when the main function selector is depressed and held momentarily, the ATC code
will automatically revert to 1200 [VFR].
The KT-70 also has a second digital display window which will show the altitude
reported by the OPTIONAL encoding altimeter or blind encoder.
Both transponders have blue/white control face lighting. The KT-76A has an amber
reply light. The automatic dimming feature matches its intensity to the cockpit lighting
conditions. They are panel mounted and meet FAA Technical Standard Orders.
Bell or King Service Center installation is required to obtain the King full year
warranty.
The complete KT-76A installation adds 1.5 kilograms (3.4 pounds).
The complete KT-70 installation adds 2.1 kilograms (4.5 pounds).
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GPS / KLN-89B
Bell Kit No. 407-705-001-zzz[PROVISIONS]
Bell Kit No. 407-705-001-zzx[EQUIPMENT]
The Bendix King KLN-89B is a GLOBAL POSITIONING SYSTEM based Navigation
System with a pilot updatable JEPPESEN database. The KLN-89B provides navigation
information suitable for VFR use ( NOT currently approved for IFR operations in the
407), as well as database information, including; Airport Data; Communication
Frequencies; Navaid Data; Airspace Information; ARTCC Information; and Minimum
Safe Altitudes.
The standard KLN-89 includes a MOVING MAP display mode.
A basic installation consists of the GPS unit, a database cartridge, an altitude input
[requires the addition of an encoding altimeter or blind encoder], and a KA 91 antenna.
Additional components may be added to increase the KLN 89's capabilities. Some
of the optional interfaces include: an external course deviation indicator (CDI) or
HSI with GPS switching provisions, RMI; some Shadin, Sheltech or ARNAV fuel
management systems; several external moving map displays; and certain models
of Bendix/King or Shadin air data systems.
Bell or King Service Center installation is required in order to obtain the King full
two year warranty.
The approximate increase in weight with the KLN-89 installed is 4.4 kilograms
(9.7pounds).
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COURSE DEVIATION INDICATOR
King KI 208
Bell Kit No. 407-705-001-121
The Course Deviation Indicator KI 208, when used in conjunction with a VHF NAV/
COMM Transceiver such as the KX 155, provides course deviation information from
the selected VHF Omni Range (VOR) station.
The kit consists of a 76 millimeter (3 inch) indicator, a VOR antenna, the required
installation hardware and wiring.
The indicator has a built-in VOR/LOC converter, blue/white lighting, prominent warning
flag when inoperative, graphic TO and FROM indication and large easy to read numbers.
The antennas mounts on either side of the aft intermediate fuselage section.
Bell or King Service Center installation is required to obtain King's full year warranty.
The kit adds 1.6 kilograms (3.6 pounds).
The Audio Panel King KMA 24H must be installed when the transceiver and Course
Deviation Indicator are installed.
The compass card of the KI 208 is turned, placing the desired course on the index at
the top, using the Omni bearing selector (OBS) knob. The deviation needle swings
right or left indicating deviation left or right of the selected course. The heading of
the helicopter has no effect on the TO or FROM indicators which react only to the
heading set on the indicator. The OFF (barber pole) disappears when a valid signal
is received.
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KCS 55A COMPASS SYSTEM & HSI
Bell Kit No. 407-705-002-101/103 [PROVISIONS/EQUIPMENT]
Bell Kit No. 407-705-002-105/107[GPS PROVISIONS/EQUIPMENT]
The KCS 55A system is an integrated compass slaving system and horizontal situation
indicator. The KI 525A Horizontal Situation Indicator (HSI) is the panel display for
the KCS 55A Compass System. It replaces the standard Directional Gyro. It combines
Magnetic Heading and VOR or LOC course information in a pictorial symbolic
presentation of the complete navigation situation. It simplifies VOR/LOC course
orientation, intercept and tracking and relieves the pilot of the chore of integrating
information from several instruments. In addition, the internally lighted instrument
includes dual glideslope pointers. The glideslope pointers are in view only during
an ILS approach. Enroute, the pointers are out of view leaving an uncluttered
presentation. The KCS 55A System, features automatic fast slaving. This means
precise aircraft magnetic heading is automatically displayed shortly after the system
is turned on. No manual setting of the compass card is required. A Slaving Control
and Compensator Unit (the KA 51B) provides selectable “slaved gyro" or “free gyro"
modes for the system and manual slaving capability when the system is in “free
gyro”mode.A visual meter displays the slaving error. It is simple to operate and is
internally lighted. The KI 525A Indicator displays the complete navigation situation
at a glance: Slaved Gyro magnetic heading Selected heading Selected VOR/RNAV
or LOC course RNAV/VOR/Localizer deviation TO/FROM RNAV or VOR indicator
Glideslope deviation. The GPS installation of the KCS 55A includes the input switching
capability between VHF NAV signals from a KX-155/165 and GPS signals from a
KLN-89B. The basic kit adds 5.2 kilograms (11.5 pounds). The GPS kit adds 6.0
kilograms (13.2 pounds).
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ENCODING ALTIMETER
(w/Avionics Master Switch)
Bell Kit No. 407-706-001-xxx
The Encoding Altimeter is a direct replacement for the standard 407 altimeter. The
function of the altimeter is identical to standard unit with the addition of AIRINC
pressure altitude signal output [referenced to standard day barometric pressure
29.92 inches of mercury (1013 millibars)] for use by (either or both) a Transponder
or GPS navigation system. The kit includes an Avionics Master Switch located on
the Overhead Console which controls all comm/nav equipment wired to the avionics
bus. The Encoding Altimeter (w/ Avionics Master Switch) adds 1.0 kilograms (2.3
pounds).
PRODUCT DATA
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FLIGHT INSTRUMENTS GROUP
Bell Kit No. 407-706-003-xxx
The flight instruments group kit augments the basic panel with the most popular
grouping of additional flight instruments: a directional gyro, attitude indicator,
and a turn and slip indicator. Although the selection of instruments and the layout
offered in the kit is only one of any number of possible arrangements, it has been
designed to complement the basic panel and to prepare for a further addition of
navigation and other aids. It is the first step to a well-coordinated full panel. The
flight instruments group kit reduces the pilot work load by providing better displays
that are easier to interpret. This can lead to more accurate flying techniques. That,
in turn, extends the operational area. It is invaluable for flight safety in low visibility
conditions, particularly at night when operating away from inhabited areas and at
any time that weather conditions deteriorate towards the limits of continued VFR
flight.
ATTITUDE
INDICATOR
DIRECTIONAL
GYRO [DG]
TURN
& SLIP
In addition to the instruments, the kit contains all hardware, tubing and wiring needed
for installation. The electrical instruments operate on direct current which is standard
in the helicopter. The integral blue-white lighting matches the standard basic panel
lighting. The kit adds 3.9 kilograms (8.7 pounds).
PRODUCT DATA
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EMERGENCY LOCATOR TRANSMITTER
Bell Kit No. 407-706-005-xxx
The Emergency Locator Transmitter (ELT) installation provides an automatic beacon
in the event of a crash. The built-in computer analyzes the nature of the shock and
operates the beacon in loads of 6 ± 1 G in any direction.
The Pointer 4000 w/3001 whip antenna is specially designed for helicopter use. It
activates at 6 ± G loads at 360 O in azimuth as well as full vertical components.
The ELT is a complete, self-contained, battery powered unit with an integral antenna
that is always fully deployed, ready for instant signal transmission. The ELT has a
switch so that it can be turned on manually. It also has a built-in test light for checking
the power output, modulation and the state of the battery. The battery has a three
year shelf life. It provides the beacon with a range up to 480 km (300 miles) for eight
days.
It can withstand 1000 G and function in temperatures from -45 O C to 71 O C (-50 O F to
160O F).
The completely self-contained ELT can be removed and hand carried.
The ELT weighs 1.8 kilograms (3.9 pounds).
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PARTICLE SEPARATOR
Bell Kit No. 206-706-212-xxx
The particle separator provides continuous protection of the engine against damage
from the ingestion of sand, dust, and other foreign debris. With the separator installed,
operations can be carried out in severe environmental conditions, for long periods
of time. The kit consists of the separator, bleed air tubing and hoses, compressor
wash fittings, and other hardware required to complete the installation. It is installed
forward of the engine air inlet just aft of the transmission fairing, and replaces the
standard engine air inlet screens..
The separator has 281 filter elements and is positioned so that all inlet air must
pass through the filter elements before entering the engine.
Each of the 281 filter elements or tube assemblies in the separator consists of a
vortex generator bonded into an inlet tube and a small-mounted inlet tube which
forms a scavenge chamber.
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Foreign particles such as dust, dirt or sand that enter the filter elements are spun in
the vortex generator and hurled centrifugally into the scavenge chamber, thereby
allowing only the remaining cleaned air to enter the engine.
The debris which collects in the scavenge chambers can be ejected overboard by a
venturi effect created by engine bleed air.
When environmental conditions indicate the need, the pilot activates the bleed air
by turning on the particle separator purge switch.
When the need disappears, the switch can be turned off and full engine power regained.
INSPECTION
WINDOW
EJECTOR
A window is installed on each side of the cowling to permit visual inspection of the
separator's plenum chamber. Ejector tubes are mounted on each side, just below
the windows. Compressor wash fittings permit the introduction of wash water directly
into the engine's inlet.
Installation imposes no aerodynamic restrictions and has no effect on the maximum
gross weights on any of the 407 configurations. Because of the pressure drop in the
inlet, there is a slight reduction of power available. This amounts to a 1% torque
loss. When the bleed air purge system is turned ON, there is a total loss of 4%.
These are covered in the Flight Manual Supplement: BHT-407-FMS-3. When
operations do not require the added protection, the separator can be removed and
the engine air intake screens replaced.
The particle separator has a tested efficiency of 85% by weight of for AC coarse
particles (27 micron nominal). Due to its inertial (centrifugal) action, the separator
is more efficient for particles larger than 27 microns, and less efficient for particles
smaller than 27 microns. Foreign debris such as hay, leaves, etc., which are too
large to enter the filter elements, are stopped at the face of the separator. However,
the numerous filter elements are so placed that there is sufficient air flow, bypassing
the blockage, even for prolonged periods of flight in such conditions.
The particle separator is not approved for flight in falling or blowing snow without
the baffles.
When the particle separator is installed, it adds approximately 7.7 kilograms (17.0
pounds) to the empty weight of the aircraft.
PRODUCT DATA
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ENGINE FIRE DETECTION
Bell Kit No. 407-706-015-xxx
The Fire Detection System will illuminate an ENGINE FIRE Warning Light in the
caution and warning panel if excessive temperature or fire develops in the engine
compartment. The system includes sensing elements in the engine compartment,
and a separate test feature to insure continuity of the sensing system. The system
is functioning properly if the single ENGINE FIRE caution segment illuminates when
the FIRE DET TEST button is momentarily depressed. The installed weight increase
for the system is 0.9 kilograms (2.0 pounds).
ENGINE FIRE LIGHT
SENSING ELEMENT
PRODUCT DATA
SENSOR CONTROL SWITCH
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QUIET CRUISE
Bell Kit No. 407-706-016-xxx
The QUIET CRUISE Kit permits flight operations at 92% NR when above 50 KIAS
and 200 feet AGL. Flyover noise level is reduced by 3.8 dBA SEL when in Quiet
Mode. The kit consists of an electrical selector switch on the collective (Pilot position
only), an annunciator on the instrument panel and additional markings on the dual
tachometer. FADEC software 5.201 or higher is required. The increase in weight
with the kit installed is 0.7 kilograms (1.5 pounds).
QC TACHOMETER MARKINGS
PRODUCT DATA
QC CONTROL SWITCH
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5250 POUND MAX. GROSS WEIGHT
Bell Kit No. 407-706-020-xxx
The 5250 Pound Increased Internal Gross Weight Kit permits operations with an
additional 113 kilograms (250 pounds) above the Normal Max. Gross Weight of
5000 pounds. The kit consists of an Airspeed Limitation Placard which replaces the
standard placard. There is no increase in empty weight with this kit installed. When
the 5250 pound Airspeed Placard is installed the 407 is restricted to the operation
limitations [IGE Hover Ceiling, OGE Hover Ceiling, and Airspeed] found in the FAA
approved Rotorcraft Flight Manual supplement for the kit [BHT-FMS-28]. Please
refer to the current year 407 Technical Information Book, or the Rotorcraft Flight
Manual for specific performance information.
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SNOW BAFFLES
(ENGINE INTAKE BAFFLES)
Bell Kit No. 206-706-208-xxx
The snow baffles are reverse flow baffles (engine air induction system deflectors)
which are required to permit flight in falling or blowing snow, when visibility exceeds
0.8 kilometers (a half mile). Hovering up to 15 minutes in blowing snow conditions
is also permitted.
The deflectors reduce the quantity of heavy snow entering the engine air intakes by
forcing the air to reverse its flow into the inlet.
The kit consists of a matching pair of deflectors that mount, one on each side of the
transmission cowling, just forward of the engine air inlets. The standard configuration
Bell 407 includes the mounting provisions for snow deflectors.
The deflectors are compatible with either basic inlet or the particle separator.
Performance limitations are listed in the Flight Manual Supplement: BHT-407-FMS4.
The deflectors are removed in temperatures above 30o C (86 oF).
The kit adds 2.4 kilograms (5.2 pounds).
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EXTERNAL CARGO HOOK
Bell Kit No. 206-706-341-xxx[EQUIPMENT]
Bell Kit No. 206-706-341-xyy[ PROVISIONS ONLY]
Installation of a cargo hook adds the capability for handling external loads quickly
and safely. It permits rapid, economical turnaround with minimum handling. The
hook has a rated capacity of 1200 kilograms (2646 pounds) and is mounted beneath
the helicopter, near the center of gravity.
CARGO HOOK
Two kits are involved:
The Hook Provisions Kit, No. 407-706-341-xyy, consists of the supports for the
suspension assembly, and associated structural provisions, fittings for the electrical
and mechanical release mechanisms and necessary installation hardware. It is
available separately so that a fleet owner can prepare more than one 407 for
maximum fleet versatility at minimum cost. Thus, any one so modified, can be employed
in an external cargo operation without delay by merely bolting the hook assembly in
place and attaching the quick connectors for the release mechanisms.
The Hook Assembly Kit, No. 407-706-341-xxx, consists of the hook which is
suspended from a short lateral beam with the hook opening facing to the right. The
beam fits into the two structural brackets and is bolted in place.
The electrical release is a switch on the pilot's cyclic stick. In the event of an electrical
failure, a mechanical release is provided. It is operated by a pull handle centrally
located between the front seats. A circuit breaker on the overhead panel completes
the kit.
A swivel link is not supplied, but is recommended to allow the load to assume a
natural stabilized position, with minimum oscillation in flight. (Maximum gross weight
of the 407 with external load is approved to 2722 kilograms (6000 pounds). The
provisions add 1.9 kilograms (4.2 pounds) and the hook assembly adds an additional
7.6 kilograms (16.7 pounds).
Flight limitations are covered in Flight Manual Supplement: BHT-407-FMS-4.
PRODUCT DATA
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ROTOR BRAKE
Bell Kit No. 206-706-502-xxx
The rotor brake kit stops the main and tail rotors rapidly after engine shutdown.
After the dynamic system stops rotating, the kit is particularly useful in preventing
the windmilling action of the main rotor in high and gusty wind conditions.
The kit is completely self-contained hydraulic system consisting of a brake disk,
dual brake assembly, master cylinder and operating handle, hose and tube assemblies
and the required fittings and hardware to complete the installation.
ROTOR BRAKE
The brake disk is between the aft flexplate of the MDS and the forward face of the
freewheeling unit output drive adapter, where if the Rotor Brake is NOT installed it
acts as an inertia weight. The stationary face of the F/W housing assembly is modified
to accept the dual-puck assembly, which is then bolted to it. The assembly encloses
two pairs of pucks in caliper configuration, horizontally located on either side of
the disk, the two calipers are also opposite each other so when hydraulic pressure
is applied, they squeeze the disk. The operating handle and the master cylinder
unit are mounted in the ceiling of the crew compartment, immediately to the right of
the overhead panel. The handle is convenient to either pilot or copilot. Pressure is
applied to the system manually by pulling down on the operating handle when the
rotor RPM is between 40 and 25 percent. When the rotor is stopped, and positioned
for parking, the handle can be pulled fully down into the parking brake [locked] position.
Spring loaded catches hold the handle in either the full off position or the parking
position. The end of the handle is painted red to prevent it being left in the parking
mode during start-up procedures. This kit adds 2.5 kilograms (5.5 pounds).
PRODUCT DATA
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HEAVY DUTY BATTERY [28 AMP-HR]
Bell Kit No. 407-706-004-xxx
The HEAVY DUTY Battery Kit is available when operations are conducted at either
extreme hot or cold ambient conditions. It provides a greater reserve starting power
and is capable of accomplishing multiple consecutive starts with minimal recharging
from the generator.
The heavy duty battery replaces the standard 17 amp-hr battery directly, using the
standard quick disconnect power and temperature sensor plugs, and no provisions
kit is required.
When installed the weight increase is 11.3 kilograms (24.9 pounds).
QUICK DISCONNECT
POWER PLUG
CASE TEMPERATURE
SENSOR PLUG
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LITTER KIT
Bell Kit No. 407-706-631-xxx[PROVISIONS]
Bell Kit No. 407-706-631-xxy[EQUIPMENT]
Bell Kit No. 407-706-631-xxz[SKI PROVISIONS]
The Bell Litter Kit for the 407 provides the capability of transporting a single litter
patient internally [dual litters are available as a customized installation]. Two
medical attendants can be accommodated at the same time and both have access
to the patient. The cabin of the standard 407 is already fitted with a double door on
the left side which gives a 1.6 meter (5 feet) wide unobstructed access to the aft
cabin. The kit includes modifications to the left front seat back and bulkhead installation,
a folding litter, and the necessary supporting hardware. The litter is aluminium and
is fitted with a pair of patient restraining straps. "Ski Provisions" is available to
accommodate litter occupants with leg injuries in bulky "air splints" by repositioning
the upper bulkhead cross member in an elevated position. Once the kit is installed
the litter can be removed or added in a few seconds, thus providing complete versatility
from ambulance to cargo or to full passenger configuration. The simple spring-loaded
restraints and the wide access make loading easy for the untrained or inexperienced
helpers.
"SKI PROVISIONS"
BULKHEAD
CROSS MEMBER
MODIFICATION
The litter can be folded and carried in the baggage compartment, taking up little
space and weight, yet with instant availability during virtually any other mission.
The copilot's collective and cyclic sticks must be stowed when the kit is installed,
however, the modification does not interfere with any other configuration requirement.
The kit adds 12.6 kilograms (27.7 pounds) if the litter and all cushions are carried
(in place or stowed). The litter weighs 10.0 kilograms (22.0 pounds). The provisions
alone weigh 2.6 kilograms (5.7 pounds). The Ski Provisions weigh 0.7 kilograms
(1.5 pounds). The Litter Kit is covered in the Flight Manual Supplement: BHT-407FMS-7.
PRODUCT DATA
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The loading sequence is as follows:
1.
Stow copilot and cyclic & collective sticks.
2.
Stow the left front seat and seat back cushions. Swing the metal back support
for the front seat forward.
3.
Open left passenger door. Open center narrow door. Swing both forward.
Secure them in the forward position with the red tape (located in the passenger door
pocket). The ring on the tape goes over the hook on the helicopter nose.
4.
Stow the left aft facing and the left and center forward facing seats and their
seat back cushions.
5.
Raise the support assembly on the center seat to a vertical position. Pull the
center seat-back support down so that it rests on the vertical support.
6.
Position the upper bulkhead cross member in the lower (standard) or upper
"air splint" position.
7.
Position litter against the forward support and the stop. Secure aft end of
litter and lock in position by pulling down and outboard on the spring loaded strap.
8.
Unhook and stow and red tape from the nose of the helicopter. Close the
doors.
PRODUCT DATA
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MAIN ROTOR BLADE FOLDING PROVISIONS
NOTE: This kit is now part of the Paravion Blade Fold Kit
The main rotor blade folding provisions kit provides the capability to reduce space
required for storage or hangaring of the helicopter. The kit consists of two expandable
bolts, which replace two of the standard blade bolts. The kit permits the use of the
Vendor STC'd kit by Paravion to fold two of the main rotor blades [one fwd. and one
aft] to reduce the required hangar size of the 407.
MINIMUM HANGAR SIZE*
ROTOR NOT FOLDED
25.1 FT X 36.6 FT
[ 7.7 M X 11.2 M ]
MINIMUM HANGAR SIZE*
ROTOR FOLDED
7.5 FT X 41.8 FT
[ 2.3 M X 12.7 M ]
TAIL
BLADE
TIPS
36.6 FEET
*ALLOWANCE SHOULD BE MADE FOR HIGH SKID GEAR, GROUND
WHEELS, EMPTY FUEL CONDITION, AND DOOR LIP WHEN
CONSIDERING HANGAR DOOR WIDTH AND HEIGHT
STING
11.2 METERS
STANDARD
BOLT
EXPANDABLE
BOLT
The main rotor blade folding provisions
[two expandable bolts] weigh 0.8
kilograms (1.8 pounds).
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CORPORATE INTERIOR
Bell Kit No. 407-705-602-xxx
CORPORATE SEATS
Bell Kit No. 407-705-702-xxx
CORPORATE SEATING & INTERIOR TRIM - The corporate cabin seating consists
of five ergonomically designed fabric covered seats with individual seat belts and
single strap shoulder harness, arranged with two extra wide forward facing outboard
seats and a middle seat for occasional use across the rear of the cabin (with a fold
down arm rest between the outboard seats) and two individual rearward facing seats
aft of the cockpit. Available with Beige, Burgundy, Blue, Teal, Tan, Charcoal, or
Light Gray upholstery fabric with color coordinated seat belts. All leather or leather
surround/fabric insert upholstery is available as an option. The corporate interior
trim consists of color coordinated Kydex ® covering for the cabin sidewalls
and doors, with lower door panel fabric / leather accent, izid leather covered arm
rests, and izid leather covered Kydex ® headliner. Selection of the corporate
interior trim also specifies installation of Corporate Soundproofing, which provides
a reduced interior noise level. The floor is covered in plush deep pile wool carpet.
The corporate seating and interior trim (and corporate soundproofing) increase the
basic aircraft empty weight 10.5 kilograms (23.1 pounds).
PRODUCT DATA
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CORPORATE SOUNDPROOFING
Bell Kit No. 407-706-603-xxx
The kit consists of a Custom Soundproofing Blanket replacing the standard blanket,
located in the aft passenger cabin roof and hatbox area to reduce the noise coming
into the cabin from the transmission and engine compartment.
The heavy duty blanket consists of two layers of fiberglass with a lead vinyl center
and a quilted cover. The standard plastic headliner covers it. The Corporate
Soundproofing kit adds 3.7 kilograms (8.2 pounds) to the standard configuration.
SLIDING CABIN WINDOW
Bell Kit No. 407-706-301-xxx
The sliding window kit is a "NO COST" option which adds two ventilation windows to
the standard passenger cabin doors. The kit adds 1.1 kilograms (2.5 pounds) to
the standard configuration weight.
PRODUCT DATA
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HIGH VISIBILITY MAIN ROTOR BLADES
High Visibility [Sales Order Addition]
The High Visibility Main Rotor Blade Paint Option provides increased safety when
operating in congested airspace. The upper surface of the main rotor blades is painted
in alternating bands of White and International Orange. There is no increase in empty
weight when this option is chosen. The standard main rotor blades required to undergo
a repaint process to incorporate this feature.
PRODUCT DATA
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SLIDING DOOR
Vendor Kit No. 407600-201[LEFT HAND]
Vendor Kit No. 407600-202 [RIGHT HAND]
The Aeronautical Accessories Sliding Passenger Door Kits may be installed either
singly (either side) or in pairs. The Sliding Doors provide increased flexibility in
loading large and bulky cargo, and are especially suited for EMS or Law Enforcement
applications. The Sliding Doors may be opened during low speed flight and hovering,
which makes them ideal for Rescue Hoist operations. The sliding Doors are compatible
with all landing gear configurations and do not interfere with AAC Step Handles or
Folding Maintenance Step [not compatible with the AAC Fixed maintenance Step].
Each Sliding door increases the empty weight 4.8 kilograms (10.6 pounds).
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MAIN ROTOR BLADE FOLDING
407BF-100-2
The Paravion Main Rotor Blade Folding Kit for the Bell 407 permits more efficient
Hangar and Ramp storage of the helicopter. The Paravion Kit consists of two removable
folding supports which attach to the nose and horizontal stabilized of the 407, a
durable Nylon bag for storage, and all the attaching hardware necessary for their
installation. The Main Rotor Blade Folding Provisions [Two Expanding Bolts] are
also included and weigh 0.8 kilograms (1.8 pounds). . Once the attaching hardware
is installed one person is capable of installing and folding, or unfolding and removing
the supports in less than 10 minutes. The installed weight of the support attachment
fittings is 0.5 kilograms (1.0 pounds).
M/R SYSTEM FOLDED
M/R HUB FOLD DETAIL
FORWARD SUPPORT AND LOCK BAR
REAR SUPPORT AND ATTACHMENT
FORWARD SUPPORT ATTACHMENT
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FORCED AIR VENTILATION
Vendor Kit No. 407V-100-2
The Air Comm Forced Air Ventilation kit provides increased cabin air exchange
for operations in temperate climates where an air conditioner is not needed. Three
light weight axial blowers and a distribution system comprise the kit. Two blowers
are in the cockpit; one on each side of the instrument panel, with individual outlets.
Fresh air is supplied from the standard fresh air inlets on the nose. The aft blower is
located in the equipment compartment behind the hat box, and distributes fresh air
through four overhead outlets in the cabin headliner. Fresh air for the passenger
cabin is supplied from an inlet on the left side of the fuselage. To assist in air exchange,
there is a cabin air exhaust opening in the hatbox area which vents to the outside
below the baggage compartment.
The kit adds 12.9 kilograms (28.40pounds).
PRODUCT DATA
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BLEED-AIR HEATER
Vendor Kit No. 407H-200-2 [STANDARD]
Vendor Kit No. 407H-202-2 [ARCTIC ]
The Air Comm Bleed-Air Heater provides heated air to the cockpit and passenger
compartments and to the windshields. The kit is a "hard-line" type heater with
individual flow control outlets located under the pilot/forward passenger seat and
in the aft passenger cabin on either side of the rear facing seats. When in operation,
engine bleed air is mixed directly with cabin air at the controllable outlet. The heated
air is directed by a powerful venturi effect eliminating the need for electric blowers.
An electric ON/OFF switch activates a solenoid valve which controls the source of
bleed air for the heater. It is mounted on the overhead control panel. The main flow
valve is conveniently mounted on aft portion of the radio pedestal. It is turned
counterclockwise for more heat [more flow]. The adjustable distribution outlets provide
individual control for different seat locations. A diverter valve located on the forward
pilots seat controls distribution to the defog outlets at the base of each windshield.
Two additional defroster slots are also provided on the inboard side of the chin
windows. A temperature sensor for the heating system is connected to a caution
light and the solenoid valve in the bleed air supply line. If the sensor detects a
temperature in excess of 104±4°C (220±8°F) the caution light illuminates. This in
turn closes the solenoid which shuts off the bleed air. Two kits are available; a standard
heater kit, and a heavy-duty ARCTIC kit with a higher air flow capacity for severe
conditions. The standard heater adds 10.3 kilograms (22.6 pounds), while the
ARCTIC heater adds 10.7 kilograms (23.6 pounds).
COCKPIT OUTLETS
[TWO PER SIDE]
"EYEBROW"
DEFOG OUTLET
REAR CABIN OUTLET
[ONE PER SIDE]
AIR RECIRCULATE OPENING
PRODUCT DATA
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AIR CONDITIONER
Vendor Kit No. 407EC-201-1 [ SINGLE FWD. EVAPORATOR]
Vendor Kit No. 407EC-201-2 [DUAL FWD. EVAPORATOR]
The Air Comm Air Conditioner Kit is a light weight driveshaft / belt powered [nonfreon R-134A refrigerant] vapor cycle cooling system. It provides cooled and
dehumidified air to the cockpit and passenger compartments. The kit consists of an
engine compartment mounted compressor [driven by a belt /pulley from the first
tail rotor drive shaft section], a belly mounted condenser unit, and evaporator units
mounted in the equipment compartment and forward of the instrument panel, and
light weight aluminum tubular plumbing with o-ring fittings and the wiring needed to
complete the installation. The system includes a refrigerant by-pass valve to eliminate
compressor cycling. The condenser has a retractable scoop which directs air through
the heat exchanger during operation. The evaporators each have electric blowers
and ducting to distribute conditioned air to the aircraft interior. When the kit is specified
an air distribution modified headliner is also specified. The headlined has four large
flow control outlets conveniently located for easy adjustment by the passengers.
The cockpit outlets are located on either side of the instrument panel and are adjustable
in direction and airflow. The system is controlled by three switches mounted in the
overhead control panel. The main switch permits operation of the blowers only o r
blowers and air conditioner. The forward and aft blower speed switches have
two positions; low or high. For operations in severe heat conditions a second forward
evaporator is available for optimum cockpit cooling. The single forward evaporator
system adds 43.3 kilograms (95.4 pounds). The dual forward evaporator system
adds 47.9 kilograms (105.6 pounds).
CONDENSER SCOOP & OUTLET
COCKPIT OUTLETS
COMPRESSOR
PRODUCT DATA
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NOTE: EITHER AIR CONDITIONER AND EITHER BLEED AIR HEATRER
MAY BE INSTALLED ALONE OR IN COMBINATIONTO P ROV IDE
COMPLETE ENVIRONMENTAL CABIN/COCKPIT CONTROL.
AIR CONDITIONER / HEATER SWITCHES
REAR CABIN CONDITIONED AIR OUTLETS [TWO PER SIDE]
REAR CABIN HEATER OUTLETS [ONE PER SIDE]
PRODUCT DATA
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WIRE STRIKE PROTECTION SYSTEM
Vendor Kit 965-37401-001[Low Skid Gear]
Vendor Kit 965-37401-002[High Skid Gear]
The wire strike protection system provides a measure of protection against inadvertent
flight into horizontally strung wires. It is designed to:
•protect the helicopter against most of the wires normally encountered in and around
metropolitan areas (15000 pound tension strength wire). •help prevent entry of wire
into the cockpit through the windshield. •reduce the possibility of flight control damage
after wire strike. •decrease the change of wire becoming entangled with the landing
gear.
The kit consists of a roof-mounted deflector/cutter, a windshield-mounted protector/
cutter, a belly-mounted deflector/cutter and all necessary hardware to complete
the installation. There are two kits; One for the standard LOW SKID GEAR, and
another for the HIGH SKID or EMERGENCY FLOAT GEAR, which has a larger deflector/
cutter due to increased skid height. The sawtoothed deflector guides the cabin into
the cutting wedge while inflicting damage to the cables. The wedge provides the
required mechanical advantage to cut the cables without any moving parts. There is
no change in flight characteristics or handling qualities of the 407 with the kit installed.
The system has been successfully tested at speeds as low as 8 kilometers an hour
(5 mph). It is equally effective at higher speeds. The LOW SKID kit adds 5.5 kilograms
(12.1 pounds). The HIGH SKID kit adds 5.7 kilograms (12.8 pounds) .
UPPER
CUTTER
WINDSHIELD
PROTECTOR
LOWER
CUTTER
PRODUCT DATA
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SEARCHLIGHT (NIGHTSUN®)
SPECTROLAB INC. SX-16
Vendor Kit SX16/016551
The Spectrolab Nightsun® series Xenon searchlight (SX-16) is a high intensity,
airborne unit, which is directionally controllable and beam width adjustable. Ground
coverage varies depending on altitude and beam width, which ranges from 4 degrees
for search to 20 degrees for flood.
The SX-16 provides a peak beam intensity of 30 million candlepower. Optional
Infrared filters are available, and may be either clip on, or electrically controllable
for visible / IR change in flight.
Installation of the SX-16 requires the addition of the High Skid landing gear. The
approximate increase in weight with the SX-16 installed is 24.9 kilograms (55.0
pounds).
NIGHTSUN® INSTALLATION
[SHOWN IS THE SX-5 AND FLIR 2000]
PRODUCT DATA
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RESCUE HOIST - 300 POUND
Vendor Kit To Be Determined
The Aeronautical Accessories, Inc. Rescue Hoist Kit for the 407 is designed for
maximum safety, utility, and ease of installation and operation. The Rescue Hoist
Kit is designed to be quickly installed or removed once the provisions are initially
fitted to the aircraft. Attachment to the external left hand cabin roof is by means of a
one piece mounting plate which incorporates a "dovetail" joint for maximum strength.
A "PIT" pin locks the assembly in place. The 300 lb. capacity Breeze-Eastern hoist
for the Bell 407 can be easily adjusted for optimum cable/skid tube clearance for
low skid gear, high skid gear, and emergency float configurations. Various cable
guard assemblies are available for the different skid configurations and when the
left hand sliding door is installed. The standard cable length is 30 meters (100 feet)
usable length. The kits come complete with 300lb. electric hoist, aerodynamic cover,
hoist provisions and all required hardware, electrical materials and complete installation
instructions. FAA/STC approved for Bell 407 helicopters. Net Weight-26 kilograms
(57 pounds).
AERODYNAMIC COVER
ELECTRICAL CONNECTION
LOCKING PIN
HOIST MOUNTING PLATE
CABLE GUARD FOR SLIDING DOOR
PRODUCT DATA
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SECTION THREE
DIMENSIONS,
TABULAR SUMMARY,
& KIT WEIGHTS
PRODUCT DATA
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PRODUCT DATA
200
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EXTERNAL DIMENSIONS
feet (meters)
DIMENSIONS ARE APPROXIMATE
[DI MENSIONS ARE ROUNDED TO THE NEXT HIGHEST TENTH FOOT(METER)]
BLADE TIPS
ENGLISH
MET RIC
10.2 FT
10.9 FT
3.1 M
3.3 M
16.0 IN
24.5 IN
0.42 M
0.62 M
34.7 FT
41.8 FT
36.6 M
10.6 M
12.7 M
11.2 M
7.5 FT
35.0 FT
25.1 FT
2.3 M
10.7 M
7.7 M
MINIMUM HANGAR SIZE*
ROTOR NOT FOLDED
25.1 FT X 36.6 FT
[ 7.7 M X 11.2 M ]
OVERALL HEIGHT
LOW SKIDS
HIGH SKIDS
GROUND CLEARANCE
LOW SKIDS
HIGH SKIDS
OVERALL LENGTH
FUSELAGE
ROTOR TURNING
ROTOR IN "X"
MAXIMUM WIDTH
FUSELAGE
ROTOR TURNING
ROTOR IN "X"
MINIMUM HANGAR SIZE*
ROTOR FOLDED
7.5 FT X 41.8 FT
[ 2.3 M X 12.7 M ]
MAIN ROTOR DIAMETER
MAIN ROTOR CHORD
NUMBER OF BLADES
35.0 FT
10.75 IN
4
10.7 M
0.273 M
4
TAIL ROTOR DIAMETER
TAIL ROTOR CHORD
NUMBER OF BLADES
5.4 FT
6.4 IN
2
1.7 M
0.16 M
2
36.6 feet
11.2 meters
TAIL
STING
*ALLOWANCE SHOULD BE MADE FOR HIGH SKID GEAR,
GROUND WHEELS, EMPTY FUEL CONDITION, AND DOOR LIP
WHEN CONSIDERING HANGAR DOOR WIDTH AND HEIGHT
PRODUCT DATA
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INTERNAL DIMENSIONS
[APPROXIMATE]
Approximate cargo space:
Aft Cabin-2.4 cubic meters (85 cubic feet)
Left front-0.6 cubic meters (20 cubic feet)
Baggage compartment-0.45 cubic meters
(16 cubic feet)
Floor loading:
Cabin-3.7 kg/sq. meter (75 lb/ sq. foot)
Baggage-4.2 kg/sq. meter (86 lb/sq. foot)
Max Baggage weight-113 kg (250 pounds)
PRODUCT DATA
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INTERNAL DIMENSIONS
PRODUCT DATA
203
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SPECIFICATION SUMMARY (ENGLISH)
(Serial No. 53467 and Subsequent)
WEIGHTS
LBS
Standard Configuration Weight (Note 1)
2640
Internal Gross Weight [Normal / Optional](*Note 2)
5000 / 5250*
External Load Gross Weight
6000
Useful Load [Gross Wt - Standard Configuration Wt] {Normal / Optional}(*Note 2)
2360/ 2610*
Maximum External Load [Cargo Hook Limit]
2646
Note 1: The Standard Configuration Weight includes 7-place upholstered interior with individual seat
belts, carpeting, and soundproofing material. Ballast is not included since it is a function of installed
optional equipment. 13 pounds of oil is included.
PERFORMANCE SUMMARY (International Standard Day Except as Noted)
• • • REFER TO DEMONSTRATED TAKEOFF & LANDING AND MAXIMUM OPERATING ALTITUDE NOTES ON THE PERFORMANCE CHARTS • • •
TAKEOFF, GROSS WEIGHT
IGE Hovering Ceiling
ISA
(4.5 Foot Skid Height)
ISA+20C
ISA+30C
OGE Hovering Ceiling
ISA
ISA+20C
ISA+30C
Service Ceiling @ MCP
ISA
(100 ft/min)
ISA+20C
Lbs
ft
ft
ft
ft
ft
ft
ft
ft
4000
19,200
15,600
13,250
17,500
13,800
11,300
20,000+
20,000+
Cruise at Long Range Cruise Speed (LRC)
Range [standard fuel]
SL, ISA
LRC Speed
Range [standard fuel]
4000 ft, ISA
LRC Speed
Maximum Cruise Speed
SL, ISA
@ Takeoff Gross Weight
SL, ISA+20C
4000 ft, ISA
4000 ft, ISA+20C
Maximum Endurance
SL
@ Loiter Speed, ISA
4000 ft
nm
kn
nm
kn
kn
kn
kn
kn
hr
hr
342
120
383
118
136
130
140
133
3.8
4.3
4500
15,600
11,600
8850
13,800
9800
6700
20,000+
17,750
336
120
375
120
135
129
138
130
3.7
4.2
5000
12,200
7900
4550
10,400
6000
1950
17,600
14,300
330
121
364
120
133
126
135
126
3.7
4.1
Note 2: Operation at Internal Gross Weight above 5000 pounds / 2381 Kilograms requires the Optional
Increased Internal Gross Weight Kit (Performance data shown on Page 17 ).
ENGINE RATING:
Allison 250-C47B with Full Authority Digital Electronic Control (FADEC):
Takeoff Horsepower(Uninstalled Thermodynamic Capability)
SHP
(Mechanical Limit)
SHP
Maximum Continuous(Uninstalled Thermodynamic Capability)
SHP
(Mechanical Limit)
SHP
813
674
701
630
TRANSMISSION RATING (Engine Output):
Takeoff Horsepower (5 minutes)
Maximum Continuous
674
630
SHP
SHP
FUEL CAPACITY (USABLE):
Standard
Auxiliary(Optional)
127.8 US Gallons
19 US Gallons
THE DATA SET FORTH ON THIS DOCUMENT ARE GENERAL IN NATURE AND MAY VARY WITH CONDITIONS. FOR PERFORMANCE DATA AND OPERATING LIMITATIONS FOR
ANY SPECIFIC FLIGHT MISSION, REFERENCE MUST BE MADE TO THE APPROVED FLIGHT MANUAL.
PRODUCT DATA
204
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SPECIFICATION SUMMARY (METRIC)
WEIGHTS
Kg
(Serial No. 53467 and Subsequent)
Standard Configuration Weight (Note 1)
1198
Internal Gross Weight [Normal / Optional](*Note 2)
2268 / 2381*
External Load Gross Weight
2722
Useful Load [Gross Wt - Standard Configuration Wt] {Normal / Optional}(*Note 2)
1070/ 1183*
Maximum External Load [Cargo Hook Limit]
1200
Note 1: The Standard Configuration Weight includes 7-place upholstered interior with individual seat
belts, carpeting, and soundproofing material. Ballast is not included since it is a function of installed
optional equipment. 6 kilograms of oil is included.
PERFORMANCE SUMMARY (International Standard Day Except as Noted)
• • • REFER TO DEMONSTRATED TAKEOFF & LANDING AND MAXIMUM OPERATING ALTITUDE NOTES ON THE PERFORMANCE CHARTS • • •
TAKEOFF, GROSS WEIGHT
IGE Hovering Ceiling
(1.4 Meter Skid Height)
OGE Hovering Ceiling
Service Ceiling @ MCP
(0.5 m/s)
ISA
ISA+20C
ISA+30C
ISA
ISA+20C
ISA+30C
ISA
ISA+20C
Cruise at Long Range Cruise Speed (LRC)
Range [standard fuel]
SL, ISA
LRC Speed
Range [standard fuel]
1200m, ISA
LRC Speed
Maximum Cruise Speed
SL, ISA
@ Takeoff Gross Weight
SL, ISA+20C
1200m, ISA
1200m, ISA+20C
Maximum Endurance
SL
@ Loiter Speed, ISA
1200m
Kg
m
m
m
m
m
m
m
m
1814
5852
4755
4039
5334
4206
3444
6096+
6096+
km
km/h
km
km/h
km/h
km/h
km/h
km/h
hr
hr
634
222
710
219
250
241
259
246
3.8
4.3
2041
4755
3536
2697
4206
2987
2042
6096+
5410
2268
3718
2408
1387
3170
1829
594
5364
4359
623
222
695
222
250
239
256
241
3.7
4.2
612
224
675
222
246
234
250
234
3.7
4.1
Note 2: Operation at Internal Gross Weight above 5000 pounds / 2381 Kilograms requires the Optional
Increased Internal Gross Weight Kit (Performance data shown on Page 17 ).
ENGINE RATING:
Allison 250-C47B with Full Authority Digital Electronic Control (FADEC):
Takeoff Horsepower(Uninstalled Thermodynamic Capability)
kW
(Mechanical Limit)
kW
Maximum Continuous(Uninstalled Thermodynamic Capability)
kW
(Mechanical Limit)
kW
606
503
523
470
TRANSMISSION RATING (Engine Output):
Takeoff Horsepower (5 minutes)
Maximum Continuous
503
470
kW
kW
FUEL CAPACITY (USABLE):
Standard
Auxiliary(Optional)
484 Liters
72 Liters
THE DATA SET FORTH ON THIS DOCUMENT ARE GENERAL IN NATURE AND MAY VARY WITH CONDITIONS. FOR PERFORMANCE DATA AND OPERATING LIMITATIONS FOR
ANY SPECIFIC FLIGHT MISSION, REFERENCE MUST BE MADE TO THE APPROVED FLIGHT MANUAL.
PRODUCT DATA
205
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Optional Accessories
[ BHTI Manufactured Kits]
• • • • • REFER TO NOTES FOR KIT COMPATIBILITY • • • • •
PartNumber
AIRFRAME
407-706-007-101
407-706-008-101
407-706-010-101
407-706-011-101
407-706-011-113
407-706-702-105
AUDIO
407-705-003-101
407-705-003-103
AVIONICS
407-705-001-133
407-705-001-103
407-705-001-105
407-706-012-101
407-705-001-107
407-706-009-101
407-705-001-109
407-705-001-111
407-705-001-113
407-705-001-115
407-705-001-117
407-705-001-119
407-705-001-121
407-705-001-125
407-705-001-129
407-705-002-101
407-705-002-103
407-705-002-115
407-705-002-117
407-706-001-101
407-706-003-101
407-706-005-101
407-706-021-101
ENGINE
206-706-212-119
407-706-015-101
407-706-016-109
407-706-020-103
Nomenclature
HIGH SKID GEAR with FLITESTEPS
EMERGENCY POPOUT FLOATS with FLOATSTEPS
SKID GEAR FAIRINGS (W/ LOW SKID ONLY)
AUX. FUEL TANK PROVISIONS (19 GAL)
AUX. FUEL TANK EQUIPMENT (19 GAL)
DUAL CONTROLS
Weight Change
lbs
Kg
32.2
14.6
110.2
243.0
11.8
5.4
5.3
2.4
29.6
13.4
12.3
5.6
AFT AUDIO ICS - 3 STATION - with 3 HEADSETS
AFT AUDIO ICS - 5 STATION - with 5 HEADSETS
8.6
13.5
3.9
6.1
VHF/ADF PROVISIONS w/ AUDIO PANEL ( KMA24H-71 )
VHF EQUIPMENT (KX-155) FOR OMNI w/CDI KI208
VHF EQUIPMENT (KX-165) FOR KCS-55A w/ HIS KI-525A
VHF ANTENNA RELOCATION
ADF EQUIPMENT (KR-87)
ADF ANTENNA RELOCATION
TRANSPONDER PROVISIONS (KT-76A) MODE C
TRANSPONDER EQUIPMENT (KT-76A) MODE C
TRANSPONDER PROVISIONS (DIGITAL) (KT-70)
TRANSPONDER EQUIPMENT (DIGITAL) (KT-70)
GPS PROVISIONS (KLN-89B)
GPS EQUIPMENT (KLN-89B, w/KA33 BLOWER)
OMNI w/ CDI KI208
COMM#2 KY196A EQUIPMENT
COMM#2 KY196A PROVISIONS
COMPASS SYSTEM KCS-55A PROVISIONS w/KI-525A
COMPASS SYSTEM KCS-55A EQUIPMENT w/ KI-525A
COMPASS SYSTEM KCS-55A PROV w/ KI-525A (GPS Comp.)
COMPASS SYSTEM KCS-55A EQUIP w/ KI-525A (GPS Comp.)
ENCODING ALTIMETER (w/ Avionics Master Switch)
FLIGHT INSTRUMENTS (for les s DG see credits)
ELT POINTER 4000 w/3001 WHIP ANTENNA
C.I.S. CERT.
9.1
8.1
8.8
2.8
6.3
-2.0
0.9
2.5
0.7
3.8
4.3
5.3
3.6
3.4
1.8
3.9
7.6
4.7
8.4
2.4
8.9
4.0
1.2
4.1
3.7
4.0
1.3
2.9
-0.9
0.4
1.1
0.3
1.7
2.0
2.4
1.6
1.5
0.8
1.8
3.4
2.1
3.8
1.1
4.0
1.8
0.5
PARTICLE SEPARATOR w/ 407H-500-1
ENGINE FIRE DETECTION
QUIET CRUISE MODE
5250 LB. MAX. GROSS WEIGHT
17.0
1.4
1.5
0.0
Notes
(1,6
(1,4
(5
(13
(18
(16
(7
(7
(7
(7
(13
(12
(19
(17
7.7 (1,15
0.6
0.7
0.0
Specifications subject to change without notice.
PRODUCT DATA
206
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Optional Accessories (continued)
[ BHTI Manufactured Kits]
• • • • • REFER TO NOTES FOR KIT COMPATIBILITY • • • • •
PartNumber
EQUIPMENT
206-706-208-103
206-706-341-109
206-706-341-123
206-706-502-105
407-705-201-103
407-706-004-101
407-706-013-101
407-706-631-101
407-706-631-105
407-706-631-107
TBD
INTERIOR
407-705-602-103
407-705-603-103
407-705-604-105
407-705-605-103
407-705-605-105
407-705-702-107
407-706-301-103
PAINT
EXTERIOR
Weight Change
Kg
lbs
SNOW BAFFLE
5.2
2.4
CARGO HOOK EQUIPMENT
16.7
7.6
CARGO HOOK PROVISIONS
4.2
1.9
ROTOR BRAKE
5.5
2.5
CARGO RESTRAINT INTERNAL PROVISIONS
1.9
0.9
28 AMPS BATTERY
24.8
11.2
COCKPIT STORAGE KIT
1.4
0.6
LITTER EQUIPMENT (ONE)
22.0
10.0
SKI PROVISIONS (for LITTER PROVISION KIT)
1.5
0.7
LITTER PROVISIONS
5.6
2.5
BLADE FOLD PROV. FOR PARAVION STC KIT [BOLTS ONLY]
1.5
0.7
Nomenclature
INTERIOR, CORPORATE
SOUND-PROOFING, CORPORATE
HEADLINER STANDARD w/ AIR CONDITIONER
ARM REST, CORPORATE
ARM REST, SLIDING DOOR
SEATS, CORPORATE
WINDOW SLIDING PASSENGER
6.9
8.2
4.0
0.3
-3.9
7.7
2.5
3.1
3.7
1.8
0.1
-1.8
3.5
1.1
Notes
(1
(2
(8
(2
(1,3,11
(1,11
(1
(1
(1
(1
(1
ANY THREE COLORS {DUPONT POLYURETHANE ONLY] APPLIED IN STD. SCHEMES
SHOWN ON PAINT SELECTOR PAGES INCLUDED IN BASIC A/C PRICE AND WEIGHT
COLOR STRIPE
ADDITIONAL COLOR STRIPE [ADDED TO STD. 3 COLORS]
TBD
TBD
407-015-001-117
MARKINGS for HIGH VISIBILITY M/R BLADES
0.0
0.0 (9
All Equipment Kits require Provision Kits prior to installation
Credit Items:
AVIONICS
206-070-263-101
INDICATOR - ALT
-1.4
-0.6
INSTRUMENT
407-899-042-101
DIRECTIONAL GYRO
-3.7
-1.7 (19
EQUIPMENT
206-706-146-105
GROUND HANDLING WHEELS
95.0
43.1
INTERIOR
407-705-601-103
INTERIOR STANDARD
-58.6
-26.6
407-705-603-101
SOUNDPROOFING STANDARD
-5.8
-2.6
407-705-604-101
HEADLINER STANDARD w/o AIR CONDITIONER
-8.0
-3.6
407-705-605-101
ARMREST STANDARD
-6.9
-3.1
407-705-700-107
CREW SEATS
-56.5
-25.6
407-705-701-107
SEATS STANDARD
-82.8
-37.6
PAINT
NO EXTERIOR
NO EXTERIOR PAINT
-16.7
-7.6
WHITE
WHITE PAINT ONLY
0.0
0.0
Specifications subject to change without notice.
PRODUCT DATA
207
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
Optional Accessories [VENDOR KITS] (continued)
• • • • • REFER TO NOTES FOR KIT COMPATIBILITY • • • • •
PartNumber
STC Items:
AIRFRAME
407-511-103
407-512-103
407-513-002
407-531-001
407-536-101
407-536-102
407-539-101
407-539-102
407-600-101
407-600-202
DL-407BF-100
ENVIRONMENT
407EC-100-1
407EC-100-3
407EC-201-1
407EC-201-2
407H-200-2
407H-202-2
407H-500-1
407V-100-2
EQUIPMENT
965-37401-001
965-37401-001
INTERIOR
407-450-001
407-453-101
407-460-001
Nomenclature
Weight Change Notes
lbs
Kg
DOOR OPENERS CREW
DOOR OPENERS PASSENGER
DOOR OPENERS BAGGAGE
STEP / HANDLE (One step)
MAINTENANCE STEP L/H
MAINTENANCE STEP R/H
MAINTENANCE STEP FIXED L/H
MAINTENANCE STEP FIXED R/H
SLIDING DOOR L/H
SLIDING DOOR R/H
PARAVION M/R BLADE FOLDING KIT
ECU w/ SINGLE EVAPORATOR, BLEED AIR HEATER
ECU w/ DUAL EVAPORATORS, BLEED AIR HEATER
AIR CONDITIONER (SINGLE FORWARD EVAPORATORS)
AIR CONDITIONER (DUAL FORWARD EVAPORATORS)
BLEED AIR HEATER w/ CHIN BUBBLE DEFROSTER
BLEED AIR HEATER ARCTIC w/ CHIN BUBBLE DEFROST
ENGINE BLEED AIR NETWORK
FORCED AIR VENTILATION
WIRE STRIKE - LOW SKID GEAR
WIRE STRIKE - HIGH SKID GEAR
FLOOR PROTECTOR (Crew and Pass enger)
FLOOR PROTECTOR (Baggage)
SPACEMAKER
2.8
2.8
0.9
1.0
2.0
2.0
2.0
2.0
10.6
10.6
TBD
1.3
1.3
0.4
0.5
0.9
0.9
0.9
0.9
4.8 (1
4.8 (1
TBD (10
125.1
135.1
99.0
110.0
20.1
22.1
2.1
26.0
56.7
61.3
44.9
49.9
9.1
10.0
1.0
11.8
12.1
12.6
5.5
5.7
8.3
5.3
6.0
3.8
2.4
2.7
(14,15,16,18
(14,15,16,18
(14,16,18
(14,16,18
(15
(15
(14
Specifications subject to change without notice.
PRODUCT DATA
208
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
Optional Accessories [EXPLANATORY NOTES] (continued)
• • • • • REFER TO NOTES FOR KIT COMPATIBILITY • • • • •
Notes: For commonality, notes shown below are identical in Technical Information and Price List.
(1) Price and / or Weight includes credit for basic ship hardware removed.
(2) A sec ond litter kit requires customizing.
(3) Must be installed prior to or concurrent with 407-705-605-103 Corporate Arm Rest.
(4) Kit includes Floatsteps (STC # 407-303-003) [weight; 29 pounds(13 kilograms)].
(5) Only compatible with low skid gear.
(6) Kit includes Flitesteps (STC # 407-326-003) [weight; 28 pounds(12.7 kilograms)].
(7) Encoding Altimeter or Blind Encoder is required to enable Mode C or Mode S altitude reporting.
Customer is responsible for obtaining Aircraft ID code for Mode S.
(8) Requires installation of -107 Litter provisions.
(9) Rework of standard blades is required to incorporate this feature.
(10) Requires ins tallation of Blade Fold Provisions (Bolts Only).
(11) Corporate interior requires c orporate soundproofing. Please s pecify corp. sndprf. kit 407-705-603-103.
(12) Will not function without the GPS provisions and equipment kit (407-705-001-117, -119)
(13) Omni kit is only compatible with KX-155, and is not compatible with any KCS-55A compass system.
(14) Includes Headliner, standard, with airconditioning(407-705-604-105), weight 4.0 lb(1.8 kg).
(15) Includes installation of 407H-500-1 Engine Bleed Air Network. Only one Bleed Air Network is required.
If both Particle Separator and Bleed Air Heater or ECU are installed subtract 2.1 lb(1 kg).
(16) ADF Antenna Relocation is required if VHF Relocation is required for Air Conditioner.
(17) Requires installation of VHF/ADF Prov. and VHF/eqpt.; Flight Inst.; ELT; and Engine Fire Det.
(18) VHF Antenna Relocation is required if VHF Prov/Equip. is in conjunction with Air Conditioner.
(19) To remove DG specify on sales order (407-899-042-101) [-3.7 lbs(-1.7 Kg)].
P. O. R. - Priced On Reques t
Specifications subject to change without notice.
PRODUCT DATA
209
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
PRODUCT DATA
210
JANUARY 2003
INDEX
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
A
AAC ® FLIGHT STEPS 155
AAC ® FLOAT STEP 157
ACCELERATION BLEED VALVE
110
ACCESSORY GEARBOX 112
ADVISORY POINTS 51
AFT AUDIO ICS 161
AFT COWLING 76
AFT FAIRING 74
AFT FIREWALL 67
AFT POSITION LIGHTS 147
AFT ROLLOVERBULKHEAD 64
AIR COOLING 117
AIR CONDITIONER 192
AIR DAM 68
AIR INTAKES 108
AIR SPLINTS 182
AIRCRAFT IDENTIFICATION 80
AIRFLOW SCHEMATIC 106
AIRFRAME 61
AIRFRAME FUEL FILTER
137, 140
AIRSPEED INDICATOR 45
ALTIMETER 45
ANTI-ICING AIR SYSTEM 110
ANTICOLLISION LIGHT 72, 147
ARINC-429 DATA BUS 127
ARM & INFLATE SWITCHES 157
ATTITUDE INDICATOR 171
AUDIO PANEL 163
AUTO RELIGHT-[WHITE] 54
AUTOMATIC DIRECTION FINDER
164
AUTOMATIC START 125
AUTORELIGHT 124
AUTOROTATION 104
AUX. FUEL LINE PROV. 69, 159
AUX. FUEL TANK
MOUNTPROVISIONS
69, 159
AUXILIARY FINS 70, 71
AUXILIARY FUEL CELL
PROVISIONS 137
AUXILIARY FUEL TANK 159
AVIONICS MASTER SWITCH
141
AXIALBLOWER 60
B
BAGGAGE COMPARTMENT 69
BAGGAGE DOOR 22
BAGGAGE DOOR-[AMBER] 54
BASIS OF CERTIFICATION 9
BATTERY ACCESS 21
PRODUCT DATA
BATTERY AND BAGGAGE
DOORS 19
BATTERY COMPARTMENT 65
BATTERY HOT-[RED] 54
BATTERY RELAY 141
BATTERY RELAY-[AMBER] 54
BATTERY SWITCH 141
BATTERY SYSTEM 141
BELL HELICOPTER CANADA 9
BLADE GRIP 81
BLEED AIR SOURCES 110
BLEED SYSTEM 110
BLEED AIR VENT EXHAUST 76
BLEED-AIR HEATER 191
BRAKE DISK 180
BRIGHT/DIM SWITCH 146
BULGED SIDE PANELS 62
BYPASS RELIEF VALVE 34
C
CABIN LIGHTS 16
CABIN STRUCTURE 64
CABIN WINDOW 24
CARBON FIBER SKIN 66
CARGO 17
CARGO HOOK 179
CARGO LOADING
INFORMATION 22
CASE TEMPERATURE SENSOR
PLUG 181
CAUTION AND WARNING
SYSTEM 52
CAUTION LIGHT PANEL 52
CAUTION PANEL SEGMENTS
53
CHECK INSTR LIGHT 51
CHECK INSTRUMENT-[AMBER]
54
CHIP / DRAIN DOOR 73
CIRCUIT BREAKERS &
SWITCHES 57
CLEVIS LIFTING 148
CLOCK / O.A.T. / VOLTMETER 48
COCKPIT INSTRUMENTS 39
COCKPIT LIGHT 146
COCKPIT OVERHEAD PANEL
141
COLLECTIVE CONTROL 95
COLLECTIVE PITCH CONTROLS
29
COLLECTIVE SLEEVE 95
COLLECTOR GEAR 93
COMBUSTION SECTION 116
COMBUSTOR MODULE 104
COMPOSITE PANEL 66
COMPRESSOR MODULE 104
211
COMPRESSOR INLET AIR TEMP
SENSOR T1 132
COMPRESSOR SECTION 109
COMPRESSOR WASH 108
CONDENSER UNIT 192
CONSTRUCTION 62
CONTINUOUS FUNCTIONAL
CHECK 126
CONTROL PANELS AND
CIRCUIT BREAKERS 143
COPPER MESH 63
CORPORATE INTERIOR 185
CORPORATE SEATS 185
CORPORATE SOUNDPROOFING
15, 185, 186
COURSE DEVIATION INDICATOR
168
COVERS AND TIEDOWNS 149
COWLINGS AND FAIRINGS 74
CREW DOORS 19
CROSSTUBES 77
CYCLIC CENTERING-[AMBER]
55
CYCLIC CONTROL 95
CYCLIC CONTROLS 28
D
DAMPER 81
DATA DOWNLOAD
CONNECTORS 40
DATA PLATE 80
DAY OR NIGHT VFR 11
DAY/NIGHT BACKLIGHT 39
DC POWER SYSTEM 141
DEBRIS COLLECTOR 93
DEFOG AIROUTLET 60
DEFOG OUTLETS 191
DEFOG SYSTEM 60
DEFROSTER SLOTS 191
DIFFERENTIAL PRESSURE
INDICATOR 34
DIRECTIONAL [TAIL ROTOR]
FLIGHT CONTROL SYSTEM
30
DIRECTIONALGYRO 171
DIVERTER VALVE 191
DOCUMENT CASE 58
DOCUMENT STOWAGE 12
DOCUMENTCASE 80
DOOR LOCKS 22
DOOR SEALS 21
DOOR SILLS 19
DOORS 18
DRAIN HOLES 25
DRIVESHAFT COVERS 70, 72
DRIVESHAFT SEGMENTS 88
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
DUAL CONTROLS 27, 160
DUAL LITTERS 182
DUAL TACHOMETER (NP/NR) 46
DUAL-POINT LATCHING 18
E
EJECTOR TUBES 174
ELASTIC DEFLECTION 77
ELASTOMERIC BEARINGS 82
ELASTOMERIC MOUNTS 97
ELECTRICAL RELEASE 179
ELECTRICAL SYSTEM 141
ELECTRICAL SYSTEM
SCHEMATIC 142
ELECTRICALLY HEATED 44
ELECTRONIC CONTROL UNIT
121
EMERGENCY FLOATS 156
EMERGENCY LOCATOR
TRANSMITTER 172
ENCODING ALTIMETER 170
END PLATES 70
ENERGY ATTENUATION 77
ENGINE AIR INLET COWLING
74
ENGINE AIR INLET COWLING
ASSEMBLY 75
ENGINE AIRFLOW 105
ENGINE ANTI-ICE-[AMBER] 55
ENGINE BELLMOUTH 107
ENGINE BLEED AIR
DISCHARGE TUBE 119
ENGINE CHIP DETECTOR[AMBER] 55
ENGINE COMPARTMENT
INSTALLATION 68
ENGINE CONDITION
MONITORING 126
ENGINE CONTROL 120
ENGINE COOLING 117
ENGINE COWLING 74, 76
ENGINE DECK 107
ENGINE FUEL PUMP 131
ENGINE INLET SCREEN 108
ENGINE MOUNTED FUEL
FILTER 131
ENGINE MOUNTS 107
ENGINE OIL FILL DOOR 76
ENGINE OIL FILTER 134
ENGINE OIL SYSTEM 133
ENGINE OIL SYSTEM
SCHEMATIC 134
ENGINE OIL TEMPERATURE/
PRESSURE INDICATOR 47
ENGINE OPERATION PRINCIPAL
106
PRODUCT DATA
ENGINE OUT-[RED] 52
ENGINE TORQUEMETER 46
ENVIRONMENTAL CONTROL
UNIT 193
EQUIPMENT COMPARTMENT
68
EVAPORATOR UNITS 192
EVENT AND EXCEEDANCE
RECORDING 39
EXCEEDANCE LIMITING 124
EXCEEDANCES 51
EXHAUST DUCT 119
EXHAUST SYSTEM 119
EXPANDABLE BOLT 84
EXTERIOR HANDLE 18
EXTERIOR LIGHTING SYSTEM
AND POSITION LIGHTS 147
EXTERIOR PAINT SCHEMES 10
EXTERNAL CARGO HOOK 179
EXTERNAL NOISE LEVEL 150
EXTERNAL POWER RELAY 144
EXTERNAL POWER SYSTEM
144
EYEBOLTS 77, 79
F
FADEC 103, 120
FADEC DEGRADED 129
FADEC DEGRADED-[AMBER] 55
FADEC FAIL 129
FADEC FAIL -[RED] 54
FADEC FAIL HORN 129
FADEC FAULT 129
FADEC FAULT-[AMBER] 55
FADEC MANUAL-[AMBER] 55
FADEC SYSTEM OPERATIONAL
FEATURES 123
FADEC SYSTEM SCHEMATIC
122
FADEC DATA PORT 127
FADECMAINTENANCESWITCH
127
FAULT CODES 130
FEATHERBEARINGS 86
FIBERGLASS SPAR 83
FIELD/IGNITER RELAY 145
FILTER ELEMENTS 173
FIRE EXTINGUISHER 58
FIREWALLS 107
FIREWALLS & ENGINE PAN 67
FIRST REDUCTION STAGE
[COLLECTOR GEAR] 93
FLEXIBLE COUPLING 87
FLEXIBLE PLATES 91
FLIGHT CONTROLS 26
FLIGHT INSTRUMENT SYSTEM
212
43
FLIGHT INSTRUMENTS 45
FLIGHT INSTRUMENTS GROUP
171
FLOAT BAGS 157
FLOAT BOTTLE 157
FOLD DOWNARM REST 16
FORCED AIR VENTILATION 190
FORWARD COWLING 74
FORWARD FAIRING ASSEMBLY
75
FORWARD FIRESEAL COVER
67
FORWARD POSITION LIGHTS
147
FORWARD
ROLLOVERBULKHEAD 64
FORWARD SECTION 63
FORWARD TANK 137
FOUR MODULES 104
FRAHM DAMPER 84
FREE WHEELING UNIT 90
FREEWHEELING ASSEMBLY
104
FUEL BOOST PUMP 137, 138
FUEL BOOST PUMPS-[AMBER]
55
FUEL FILLER 137
FUEL FILLER CAP 136
FUEL FILTER (AIRFRAME)[AMBER] 55
FUEL LOW-[AMBER] 55
FUEL NOZZLE GROUND 80
FUEL PRESSURE INDICATOR
AND DC AMPMETER 48
FUEL QUANTITY INDICATOR 48
FUEL QUANTITY PROBES 137
FUEL QUANTITY PROBES 140
FUEL REQUIREMENTS 140
FUEL SHUTOFF VALVE 137
FUEL SHUTOFF VALVE AND
FUEL PRESSURE
TRANSDUCER 140
FUEL SUMP DRAIN VALVES 138
FUEL SYSTEM 136
FUEL SYSTEM COMPONENTS
137
FUEL SYSTEM INDICATORS 139
FUEL SYSTEM SCHEMATIC 139
FUEL TRANSFER PUMP 137
FUEL TRANSFER PUMPS 138
FUEL TRANSFER PUMPS[AMBER] 56
FUEL VALVE SWITCH 140
FUEL VALVE-[AMBER] 56
FUELFILLERCAP 80
FULL AUTHORITY DIGITAL
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
ENGINE CONTROL 103
FUSELAGE 63
FUSELAGE FEATURES 65
FWD. FIREWALL 67
G
GAS GENERATOR (NG)
GOVERNOR 123
GAS PRODUCER GEARING 113
GAS PRODUCER TACHOMETER
(NG) 47
GEAR SHAFT CASE 93
GEAR TRAIN SCHEMATIC 96
GEARBOX MODULE 104
GEARBOX COVER 70
GENERATOR 144
GENERATOR CONTROL UNIT
144
GENERATOR FAILURE-[AMBER]
56
GENERATOR RESET SWITCH
144
GENERATOR SHUNT 144
GENERATOR SYSTEM 143
GLOBAL POSITIONING SYSTEM
167
GPS / KLN-89B 167
GROUND HANDLING 79
GROUNDING PLUG 136
H
HANDGRIPS 16
HANGER BEARING 88
HEAVY DUTY BATTERY 181
HIGH ROTOR RPM-[RED] 52
HIGH SKID LANDING GEAR 155
HINGE SUPPORT 76
HISTORICAL FAULTS 130
HOOK ASSEMBLY KIT 179
HOOK PROVISIONS KIT 179
HORIZONTAL SITUATION
INDICATOR 169
HORIZONTAL STABILIZER 70,
71
HOURMETER 49
HYDRAULIC FILTERS 36
HYDRAULIC PUMP 36
HYDRAULIC RESERVOIR 36
HYDRAULIC SWITCH 37
HYDRAULIC SYSTEM 34
HYDRAULIC SYSTEM
SCHEMATIC 35
HYDRAULIC SYSTEM-[AMBER]
56
HYDROMECHANICAL CONTROL
UNIT 121
PRODUCT DATA
I
IGNITER 145
IGNITION EXCITER 132
IMPENDING EXCEEDANCE 40
INCLINOMETER 45
INDUCERBLEED PORT 110
INPUT GEAR SHAFT AND INPUT
SPIRAL BEVEL GEAR 93
INSTRUMENT PANEL 42
INTERCONNECTLINE 137
INTERIOR HANDLE 18
INTERIOR LIGHTING 146
INTERMEDIATE SECTION 66
INTERNATIONAL
CERTIFICATION 9
J
JACKING 148
K
KAFLEX 91
KCS 55A COMPASS SYSTEM
169
KEVLAR 85
KEVLAR SKINS 85
KING KMA 24H AUDIO PANEL
163
KING KX 155 162
KMA-24H ICS CONTROL PANEL
163
KR 87 ADF RECEIVER 164
KT 70 166
KT 76A 166
L
LANDING GEAR 77
LANDING GEAR FAIRINGS 158
LANDING GEAR FAIRINGS 158
LANDING GEAR MICROSWITCH
48, 49, 78
LANDING LIGHTS 147
LCD EXCEEDANCE
MONITORING 51
LCD INSTRUMENTS 41
LCD TORQUE / WITH TEST
SWITCH 40
LEADING EDGE SLAT 71
LEVELING 148
LEVELING PLATE 148
LIFTING 148
LIGHTING PROTECTION 63
LIGHTING SYSTEMS 146
LINE CONTROL RELAY 144
LIQUID CRYSTAL DISPLAYS 39
LITTER DOOR 20
213
LITTER DOOR INSTRUCTIONS
20
LITTER DOOR-[AMBER] 56
LITTER KIT 182
LITTER PANEL WINDOW 24
LITTON SPECIAL DEVICES 39
LOW LEVEL FUEL WARNING
SWITCH 140
LOW ROTOR RPM-[RED] 52
LOW SKID FAIRINGS 78
LOWER MAST BEARING
SUPPORT 93
LOWER WINDOWS 25
LOWERCUTTER 194
M
MAIN CASE 93
MAIN DRIVESHAFT 90
MAIN INPUT QUILL 93
MAIN ROTOR 81
MAIN ROTOR BLADE FOLDING
184
MAIN ROTOR BLADES 83
MAIN ROTOR DRIVE SYSTEM
90
MAIN ROTOR HUB 81
MAIN ROTOR MAST 94
MAIN ROTOR SERVOS 35
MAIN SWITCH PANEL 57
MAIN TANK 137
MAINTENANCE BUTTON 130
MANUAL MODE 127
MANUAL THROTTLE 29
MANUFACTURING AUTHORITY
9
MAST 90
MAST ASSEMBLY 94
MAST NUT 81
MEASURED GAS
TEMPERATURE (MGT)
INDICATOR 47
MISCELLANEOUS
INSTRUMENTS 48
MODULAR DESIGN 104
N
NAVIGATION TRANSCEIVER
162
NG OVERSPEED 126
NICKEL CADMIUM BATTERY
141
NIGHTSUN® 195
NOMEX HONEYCOMB 83
NON VENTED FILLER CAP 98
NON-VOLATILE MEMORY (NVM)
49
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
NON-ICING CONDITIONS 11
NONSLIP COATING 77
NP OVERSPEED SYSTEM 125
O
OCCUPANT PROTECTION 64
OCCUPANT RESTRAINT 12
OIL BREATHER AND LINE 99
OIL COOLER 103, 135
OIL COOLER FAN 89
OIL COOLEREXHAUST 76
OIL DRAIN PLUG 99
OIL FILLER DOOR/SCREEN 73
OIL FILTER ELEMENT AND
HOUSING 100
OIL LEVEL SIGHT GAGE 98
OIL PUMP 99
OIL TANK 135
OPTIONAL CABIN SLIDING
WINDOW 59
OPTIONAL M/R BLADE FOLDING
84
OPTIONAL SLIDING WINDOW
19
OVERHEAD CIRCUIT BREAKER
AND SWITCH PANEL 143
OVERHEAD CONSOLE 57
OVERSPEED FUNCTIONAL
TEST 126
P
PARKING AND MOORING 149
PARTICLE SEPARATOR 173
PEDAL ADJUSTER 33
PERFORMANCE MONITORING
SYSTEM 40
PERMANENT MAGNET
ALTERNATOR 132
PERMANENT MEMORY 128
PILOT AND FORWARD
PASSENGER SEATING 12
PILOTS' CYCLIC 28
PILOTS DOOR WINDOW 24
PITCH HORN 82
PITOT TUBE 44
PITOT/STATIC SYSTEM 44
PLANETARY ASSEMBLY 93
POLYURETHANE PAINT 10
POSITION LIGHTS 147
POWER PLANT 103
POWER TURBINE (NP)
GOVERNOR 123
POWER TURBINE GEARING 114
POWER UP FUNCTION CHECK
126
PROPULSION INSTRUMENTS
PRODUCT DATA
46
PYLON ASSEMBLY 90, 97
Q
QUICK DISCONNECT 27
QUICK DISCONNECT 34, 160
QUICK DISCONNECT FITTINGS
37
QUICK DISCONNECTPOWER
PLUG 181
R
RADIO PEDESTAL 58
RAM AIR INLET 60
RAM AIR SYSTEM 59
RECESSED STEP 65
RELIEF VALVE 37
RESTART FAULT 129
RESTART FAULT-[WHITE] 56
RESTRAINT 97
RING GEAR 93
ROLLOVER PROTECTION 64
ROTOR BRAKE 180
ROTOR BRAKE 180
ROTOR TRACK & BALANCE
PROVISIONS 89
RPM WARNING HORN MUTING
SYSTEM 53
S
SAVITAD 97
SEARCHLIGHT 195
SERVO ACTUATOR
DESCRIPTION 38
SERVO ACTUATOR SUPPORT
36
SEVEN PLACE SEATING 11
SHEAR BEARINGS 82
SHOULDER HARNESS 12
SIGHTGLASS 97
SINGLE LITTER 182
SINGLE PILOT OPERATION 11
SKID SHOES 78
SKYLIGHT 25
SLAM CLOSING 18
SLATS 71
SLIDING CABIN WINDOW 186
SLING LIFTING 148
SNOW BAFFLES 178
SOLENOID SWITCH 34
SOLENOID VALVE 37, 157
SOLID STATE DESIGN 39
SOUNDPROOFING 17
SPACER 90
SPEED PICKUP SENSORS 132
214
SPRAG CLUTCH 104
STABILIZER ASSEMBLY 71
STAINLESS STEEL 83
STAINLESS STEEL ABRASION
STRIP 83, 85
STANDARD BOLT 84
STANDARD SEATING &
INTERIOR TRIM 13
STANDBY MAGNETIC COMPASS
46
START-[WHITE] 56
STARTER GENERATOR 111,
144,145
STARTER RELAY 145
STARTER SWITCH 145
STARTER/IGNITER SYSTEM 145
STATIC PORTS 44
STOWABLE HEADREST 13
STRUCTURAL INTEGRITY 64
SUN GEAR 93
SUPPORT BEAM 77, 78
SURGE DETECTION AND
RECOVERY 124
SWASHPLATE 95
SWASHPLATE 28, 29
SYSTEM BUILT IN TEST 125
T
TAIL LIGHT 73, 147
TAIL ROTOR 85
TAIL ROTOR BLADE 85
TAIL ROTOR DRIVE SYSTEM 87
TAIL ROTOR DRIVESHAFT
FIRESEAL COVER 67
TAIL ROTOR GEARBOX 86
TAIL ROTOR GEARBOX CHIP
DETECTOR-[AMBER] 56
TAIL ROTOR GEARBOX FAIRING
73
TAIL ROTOR HUB 86
TAIL ROTOR PEDALS 33
TAIL ROTOR PITCH CHANGE 87
TAIL ROTOR SERVO 35
TAIL SKID 72, 78
TAILBOOM SECTION 70
TEMPERATURE SENSOR 191
THE PILOT'S STATION 26
THREE JACK POINTS 148
THRUST BALANCE 117
TIE DOWN LOOPS 69
TITANIUM ENGINE PAN 67
TMOP (TORQUE METER OIL
PRESSURE) 124
TOP CASE 93
TORQUE METER OIL
PRESSURE SENSOR 132
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
TORQUEMETER 115
TOW RING 79
TOWING RING 155
TRANSFER SYSTEM 138
TRANSMISSION 90
TRANSMISSION &
FREEWHEELING CHIP
DETECTORS 102
TRANSMISSION ASSEMBLY 92
TRANSMISSION CHIP
DETECTOR -[AMBER] 56
TRANSMISSION COWLING 74
TRANSMISSION COWLING
ASSEMBLY 75
TRANSMISSION GEAR RATIO
96
TRANSMISSION HIGH
PRESSURE OIL FILTER
BYPASS VALVE 101
TRANSMISSION MOUNTING 97
TRANSMISSION NR MONOPOLE
PICKUP [RPM SENSOR] 95
TRANSMISSION OIL FILLER
CAP 99
TRANSMISSION OIL
FILTERBYPASS VALVE &
INDICATOR 101
TRANSMISSION OIL JETS 100
TRANSMISSION OIL LEVEL
SIGHT GAGE 100
TRANSMISSION OIL MANIFOLD
100
TRANSMISSION OIL PRESSURE
REGULATOR VALVE 101
TRANSMISSION OIL PRESSURE
SWITCH 102
TRANSMISSION OIL PRESSURE
TRANSDUCER 102
TRANSMISSION OIL SYSTEM
98
TRANSMISSION OIL SYSTEM
HOSES & TUBE
ASSEMBLIES 102
TRANSMISSION OIL TEMP BULB
102
TRANSMISSION OIL
TEMPERATURE/PRESSURE
INDICATOR 47
TRANSMISSION OIL
THERMOSTATIC VALVE 101
TRANSMISSION OIL
THERMOSWITCH 102
TRANSPONDERS 165
TREND ARC 49
TURBINE MODULE 104
TURBINE SECTION 118
PRODUCT DATA
TURBINE SECTION SCHEMATIC
118
TURN& SLIP 171
TWO RUGGED BULKHEADS 64
U
UNION FILTER 100
UPPER CUTTER 194
V
VENTCONTROLKNOB 60
VENTILATION SYSTEM 59
VENTLINES 137
VERTICAL FIN 70, 72
VERTICALCONTROLTUNNEL 64
VHF COMMUNICATION 162
VORTEX GENERATOR 174
W
WATER DRAIN 60
WEIGHT POCKET 84
WINDOWS 23
WINDSHIELDPROTECTOR 194
WINDSHIELDS 23
WIRE STRIKE PROTECTION
SYSTEM 194
Y
YOKE 82
215
JANUARY 2003
THIS IS AN UNCONTROLLED DOCUMENT.
THIS ELECTRONIC DOCUMENT MAY HAVE BEEN SUPERCEDED.
THE MOST CURRENT DOCUMENT INFORMATION IS AVAILABLE
FROM YOUR BELL HELICOPTER TEXTRON REPRESENTATIVE.
P.O. Box 482, Fort Worth, Texas 76101,
Phone: (817) 280-2800, Fax: (817) 278-2800
www.bellhelicopter.textron.com
BELL HELICOPTER TEXTRON
DIVISION OF TEXTRON CANADA LTD.
12,800 rue de l' Avenir Mirabel, Quebec, Canada J7J1R4
Phone: (450) 437-2729 Fax: (450) 437-2066
The data set forth in this brochure are g eneral in nature and may vary with conditions. For performance data and operating
limitations for any specific flight mission reference must be made to the approved flight manual.
Written, Edited & Layout b y:
Dave Wyatt
817-280-6918
E-Mail Address: dwyatt@bellhelicopter.textron.com
PRODUCT DATA
© 2003 Bell Helicopter Textron Inc.
All Rights Reserved.
Printed in USA
[Rev. 2 - 031015] Januar y 2003
216
JANUARY 2003