Flying The Bell 206 JetRanger
A Training Manual For Pilots
John Toal
Nothing in this manual supersedes any legislation, rules, regulations or
procedures contained in any operational document issued by the Government,
the Aviation Authority of your respective country, the manufacturers of aircraft,
engines and systems or by the operators of aircraft throughout the world.
Copyright © 2012
First edition published 2012
Kindle edition published 2017
All rights reserved. No part of this book may be reproduced or transmitted in any
form by any means, electronic or mechanical including photocopying, recording
or by any information storage and retrieval system without permission from the
publisher in writing.
TABLE OF CONTENTS
PREFACE
HISTORY OF THE BELL 206
GENERAL DESCRIPTION
AIRFRAME
CREW COMPARTMENT
ELECTRICAL SYSTEM
FUEL SYSTEM
POWERPLANT
DRIVE TRAIN AND ROTORS
FLIGHT CONTROLS AND HYDRAULICS
WEIGHT AND BALANCE
PERFORMANCE
NORMAL PROCEDURES
EMERGENCY PROCEDURES
OTHER BOOKS BY JOHN TOAL
PREFACE
When teaching students how to fly the JetRanger, I was never able to get good
quality instructional material specific to this helicopter. The notes I used seem to
have been around in various formats, in different schools for the past twenty
years. These notes often contained irrelevant information and in some areas did
not contain enough information. I decided to rewrite these notes and in so doing,
I ended up with this book.
This book is not a substitution for the helicopter flight manual and should not be
treated as such. It is intended to give further information relating to the different
sections of the flight manual and to explain the helicopter systems more clearly. I
have added in some extra information where I felt it would be beneficial but
overall I have tried to keep the book as simple as possible while still covering all
the sections in the flight manual.
The helicopter flight manual or pilot’s operating handbook is by law, required to
be carried in the helicopter. It will not teach you how to fly the aircraft. It will
give you the necessary technical information to conduct your flight safely. It is
the legal document for information relating to the helicopter.
As such, it is important that you understand it and know where to find the
relevant information in it. Before flying the JetRanger for the first time, read the
flight manual thoroughly and become comfortable with finding information in
the different sections.
The book is written with the new JetRanger pilot in mind. This book assumes
you already hold your private pilot’s licence and have limited or no previous
JetRanger experience. You should be familiar with flying techniques, as this
book will not teach you how to fly.
My thanks and appreciation go to all those who helped me with this book.
Special thanks go to the engineers and instructors who made valuable input and
also to my students. My students always keep me on my toes and their mistakes
and questions gave me the motivation for writing this book.
Many thanks also to the owners of the helicopters I photographed for this
publication.
Rolls Royce was also very helpful and supplied me with high quality drawings
and diagrams that speak a thousand words.
I hope you will find this book a valuable aid while learning to fly the JetRanger.
I am sure experienced pilots will find enlightening information when reading
this book as I know when I was learning; detailed systems information was
difficult to come by.
Terminology
WARNINGS, CAUTIONS, AND NOTES: Warnings, cautions, and notes are
used in this manual to emphasize important and critical instructions as follows:
WARNING
An operating procedure, practice, etc., which, if not correctly followed, could
result in personal injury or loss of life.
CAUTION
An operating procedure, practice, etc., which if not strictly observed, could result
in damage to or destruction of equipment.
NOTE
An operating procedure, condition, etc., which is essential to highlight.
Use of Procedural Words
Concept of procedural word usage and intended meaning which has been
adhered to in preparing this manual is as follows:
SHALL has been used only when application of a procedure is mandatory.
SHOULD has been used only when application of a procedure is recommended.
MAY and NEED NOT have been used only when application of a procedure is
optional.
WILL has been used only to indicate futurity, never to indicate a mandatory
procedure.
Abbreviations and Acronyms
Abbreviations and acronyms used throughout this training manual are defined as
follows:
ADF: Automatic Direction Finder
Air Cond: Air Conditioning
A/C: Aircraft
Alt: Altitude
A/F: Airframe
ANTI COLL LT: Anti-Collision light
APU: Auxiliary Power Unit
AUX: Auxiliary
BATT: Battery
BLO: Blower
C: Celsius
CAUT: Caution
CG: Center of Gravity
CKPT COMM: Cockpit Communications
CONT: Continuous
CWP: Caution Warning Panel
DC: Direct Current
DG: Directional Gyro
ECS: Environmental Control System
ELT: Emergency Locator Transmitter
ENCDG: Encoding Altimeter
ENG: Engine
F: Fahrenheit
FS: Fuselage Station
FT: Foot, Feet
FWD: Forward
GEN: Generator
GPS: Global Positioning system
GOV: Governor
GW: Gross Weight
HD: Density Altitude
Hg: Inches of Mercury
Hp: Pressure Altitude
HYD: Hydraulic
ICS: Intercommunication System
IGE: In Ground Effect
IGNTR: Igniter
IMC: Instrument Meteorological Conditions
IN: lnch(es)
IFL: Inflate
INSTR CHK: Instrument Check
INSTR LT: Instrument Light
KCAS: Knots Calibrated Airspeed
KG: Kilogram(s)
KIAS: Knots Indicated Airspeed
KTAS: Knots True Airspeed
L: Liter(s)
LBS: Pound(s)
LDG LTS: Landing Lights
LT: Light
L/FUEL: Left Boost Pumps
MCP: Maximum Continuous Power
MM: Millimeter(s)
MPH: Miles Per Hour (statute)
NAV: Navigation
N1: Gas Producer RPM (N1)
N2: Power Turbine RPM (N2)
NR: Main Rotor RPM
OAT: Outside Air Temperature
OBS: Omni Bearing Selector
OGE: Out of Ground Effect
OVSPD: Over speed
PART SEP: Particle Separator
PAX: Passenger
POS LT: Position Light
PRESS: Pressure
PSI: Pounds per Square Inch
PWR: Power
QTY: Quantity
RECP: Receptacle
RLY: Relay
RPM: Revolutions per Minute
R/FUEL: Right Boost Pumps
SHP: Shaft horsepower
SL: Sea level
SPKR: Speaker
SYS: System
TEMP: Temperature
TOT: Turbine Outlet Temperature
T/R: Tail rotor
TRANS: Transmission
TRQ: Torque
V: Volt(s), Voltage
VFR: Visual Flight Rules
VNE: Never Exceed Velocity
VOR: VHF Omnidirectional Range
WL: Water Line
WRN: Warning
XFR: Transfer
XMSN: Transmission
XPNDR: Transponder
1
HISTORY OF THE BELL 206
Bell Helicopter History
Lawrence D. Bell founded the Bell Aircraft Corporation, on July 10, 1935 in
Buffalo, New York. Bell Aircraft Corporation created a Helicopter Division,
which moved to Fort Worth, Texas in 1951 and became the Bell Helicopter
Corporation.
In 1960, Textron bought various Bell Aircraft properties including the Helicopter
operation. Textron changed the name of the helicopter operation to the Bell
Helicopter Company, and established itself as Textron’s largest division. In 1976
the name changed to “Bell Helicopter Textron”.
When Bell Aircraft moved the helicopter operation to Fort Worth, Texas, all of
the initial operations, including flight and maintenance training in the Model 47
took place at a leased facility in Saginaw, Texas just north of Fort Worth. This
was originally a naval facility that was acquired late in the World War II era.
After the war this flying field was leased to Bell Helicopter.
The Model 206A “JetRanger” was added to the commercial production line in
1967.
206 History
To meet the US Army specification for a light observation helicopter (LOH) in
1960, Bell Helicopter’s entry aircraft for the competition was the OH-4. Four of
these aircraft were produced. SN 001-003 airframes were dismantled and SN004 was converted into a 206A which was certified with the type certificate H92. All 206, 206L and 407 aircraft have the H-92 Type Certificate.
The three finalists in the LOH contest were Bell, Hiller and Hughes. Bell
decided to use the technology developed during the LOH contest as the basis of
a new commercial helicopter. (The Model 206A). The 206A became known as
the JetRanger. It flew in December 1965 and was certified in October 1966.
When it was time for the U.S. Army to procure additional LOH helicopters the
Army reopened the LOH bidding and awarded the contract to Bell for the OH58A Kiowa, a military version of the Model 206A, for which the initial order
was 2,200 aircraft. Deliveries of the OH-58A began in 1969 and the Kiowa was
deployed to Vietnam that summer for both LOH and transport duties. Deliveries
of the JetRanger started in 1967 at Fort Worth, Texas and were transferred to
Mirabel, Canada in 1986. In 1993 the US Army ordered the TH-67 Creek
version as its new training helicopter to succeed the Bell UH-1.
The first flight for the 206L (Long Ranger) was in 1975. The aircraft was
equipped with an Allison 250C20B engine with a water-methanol injection
system to cool the exhaust gas temperature at high workloads and was certified
with a maximum all up weight of 4,050 Lbs.
The 206L-1 was introduced in 1978 and was equipped with the Allison
250C20B. This aircraft could be modified to carry 4,150 Lbs. Many of these
have been upgraded to L-3s. The 206L-3 became available in 1981 and was
equipped with the Allison 250C30P and is certified with a maximum all up
weight of 4,150 Lb. The engine output was 650HP. The 206L-4 became
available in 1992 and is equipped with the Allison 250C30P and was certified
with a maximum all up weight of 4,450 Lbs. The transmission was upgraded to
allow it to carry this weight.
From the beginning with the OH-4, to the current models 206B3, 206L4 and
407; approximately 10,000 have been manufactured to date. This total does not
include aircraft manufactured under license in Japan, Italy, and South Africa.
The 206 series is the most successful helicopter ever produced, setting the world
standard of utility, safety and reliability.
2
GENERAL DESCRIPTION
The JetRanger is a single engine, five place, utility-type helicopter designed for
takeoff and landing on any reasonably level terrain. The standard configuration
provides for one pilot and four passengers. Removable dual controls are
available for two pilot operations and training. The helicopter can cruise at
approximately 100kts and burns fuel at around 27 US gallons per hour and is
capable of climbing to anywhere between 13500 feet to 20000 feet in ISA
conditions (depending on its weight).
The JetRanger is one of the most popular light, turbine powered helicopters ever
built. It has accumulated millions of hours of flight and according to the NTSB it
is the safest single engine aircraft in the world.
Airframe Description
The fuselage consists of three main sections: (Ref Fig.1) the FORWARD section
which extends from the cabin nose to the bulkhead aft of the passenger
compartment, the INTERMEDIATE or transition section which extends from the
bulkhead aft of the passenger compartment to the start of the tail boom, and the
AFT or TAIL BOOM section.
The airframe forward section consists, primarily, of an aluminum honeycomb
structure. This construction provides the maximum strength to weight ratio and
rigidity. The windscreen is made of transparent tinted plastic as are the cabin
roof panels, and door panels.
The intermediate section is a semi-monocoque construction with a metal and
fiberglass covering. The aft section is made of an aluminum alloy monocoque
tail boom that supports the vertical fin, a fixed horizontal stabilizer, tail rotor and
tail rotor drive train.
The primary load carrying structures are two built-in cabin bulkheads, a vertical
control tunnel from the floor to the cabin roof and a pair of longitudinal beams in
the cabin roof.
The landing gear is a tubular skid type made of aluminum alloy to which
optional floats may be attached. There are various types of skids available but
the main types used are “high” skids and “low” skids. (Discussed later in the
book).
Crew Compartment (General)
The pilot’s station is situated on the right side of the crew compartment. The left
side can accommodate a passenger, or a co-pilot if the dual controls are installed.
The instrument panel is mounted on a central pedestal in front of the two crew
seats. The panel is tilted upward for maximum visibility from either seat.
The flight instruments are on the right of the panel and the systems instruments
are in two rows to the left of the flight instruments. Caution and warning lights
are mounted in a narrow band across the top of the instrument panel, under the
glare shield. This is called the Caution Warning Panel (CWP).
Electrical System (General)
The Bell 206B III helicopter is equipped with a 28 volt direct current (dc)
electrical system. Power for this system is obtained from a nickel-cadmium 24
volt, 13 ampere-hour battery and a 150 ampere, 30 volt combination startergenerator. On older helicopters this battery may have been replaced with a leadacid battery.
Fuel
The basic aircraft has a fuel capacity of 91 US gallons. Earlier models had a fuel
capacity of 75 US gallons but with the fitting of a range extender this brought
the capacity up to a maximum of 96.7 US gallons. The 206B III has a capacity of
91 US gallons. The aircraft is designed to operate on standard aviation jet fuels.
The fuel is stored in a crash proof bladder under the rear passenger seat.
Engine
The JetRanger III is powered by a Rolls Royce shaft turbine engine Model 250C20J manufactured by the Detroit Diesel Allison Division of General Motors
Corporation.
The 250-C20J is a 420 shaft horsepower engine derated to a takeoff rating of 317
shaft horsepower.
Drivetrain and Rotors
The drivetrain system provides a means of transmitting power from the engine to
the main and tail rotor assemblies. The engine powers the accessories gearbox
(discussed later) which drives the main rotor gearbox and the tail rotor gearbox.
The main rotor is a semi-rigid, seesaw, two blade design that uses a pre-coned
and under slung feathering axis to ensure smooth operation.
The tail rotor is also a semi-rigid, seesaw, two blade design.
Main Rotor Blades (General)
The main rotor blades are an all metal construction with an aluminum alloy
honeycomb core, aluminum skins, spar and trailing edge strip. All the structural
components are joined by means of metal-to-metal bonding. The blades are set
in hub grips and are secured by a single retaining blade bolt in each grip.
Main Rotor Hub
The main rotor hub consists primarily of a forged steel yoke with two spindles, a
trunnion assembly, two blade grips with pitch horns and two grip-retention strap
assemblies. Oil or grease lubricated bearings provide for smooth rotation of the
trunnion and blade grips on the yoke.
The blades are attached to the hub grips with bolts which have hollow shanks for
installation of weights for static balance of the hub and blade assembly. After
balancing, the bolts must be kept with their respective rotor hub grips. The flap
restraint assembly incorporates counterweights and springs which serve to
position limited freedom flapping stops. The stops prevent excessive flapping of
the main rotor during starting and shutdown but allow normal flapping at
operating RPM.
Main Rotor Drive System
The main rotor drive system provides a means of transmitting power from the
engine to the main rotors. The main rotor drive system consists of a
transmission, main rotor mast, main driveshaft, freewheeling assembly and oil
cooler. The freewheeling assembly is mounted on the engine accessory gear
case. The main driveshaft is located forward of the freewheeling assembly and
large “couplers” are visible at each end. These couplers have “Temp-plate” strips
stuck to the surface (two red and two yellow on each coupler). The temp-plates
will give an early indication of overheating of the couplers and during the walkround the strips should be inspected for any darkening of the white dots on their
surface.
Tail Rotor Drive System
The tail rotor is a semi-rigid, delta hinged two blade design. The delta hinge
compensates for dissymmetry of lift by allowing the blades to flap to equality.
The tail rotor gear box has an oil sight glass. There are two oil level lines drawn
on the glass. One is to indicate the oil level when the helicopter is equipped with
low skids and the other one is to be used when the helicopter is equipped with
high skids.
Flight Controls and Hydraulics (General)
The flight controls are mechanical linkages that are actuated by conventional
controls and used to control flight attitude and direction. Both the cyclic and the
collective controls incorporate hydraulic servo actuators.
The main rotor controls consist of the swashplate, drive link, and pitch links.
The hydraulic system provides pressurized fluid to operate the cyclic and
collective flight control servo actuators. Operation of the system is electrically
controlled by means of the hydraulic system switch.
There are no accumulators on the B206 to supply some power in the event of a
hydraulic failure and therefore a hydraulic failure can cause a sudden movement
of the controls.
Helicopter Components
1. Main rotor
2. Air intake grill
3. Engine compartment
4. Vertical fin
5. Tail skid
6. Tail rotor
7. Horizontal stabilizer
8. Tail boom
9. Skid gear
10. Passenger /cargo doors
11. Crew doors
12. Baggage compartment
Principal Dimensions
General Data
Engine
*Model Number:
250-C20J
Manufacturer:
Allison Div. of GMC
Max Continuous Power (SHP):
85% Torque = 270 HP
Take-Off Power (Rated):
100% Torque = 317 HP
Transient:
110% Torque for 5 SECONDS
*The 250-C20J is a 420 SHP engine de-rated to 317 SHP for installation in the
JetRanger.
Airspeed Limitations
Vne Less Than 3000 lbs:
130KTS (150 MPH)
Over 3000 lbs:
122KTS (140 MPH)
Aft Door(s) OFF:
87 KTS (100 MPH)
Front Door(s) OFF: 69 KTS (80 MPH)
85 to 100% Torque: 80 KTS (92 MPH)
Recommended Airspeeds
Engine Failure:
50-60KTS (57-69 MPH)
Maximum Autorotation:
100KTS (115 MPH)
Minimum Rate of Descent:
52 KTS (60 MPH)
Maximum Glide Distance:
69 KTS (80 MPH)
Hydraulic Failure:
61-69 KTS (70-80 MPH)
Maximum Altitude
Less Than 3000 lbs: 20,000 Ft
Over 3000 lbs:
13,500 Ft
Weights
Equipped weight with standard configuration, plus oil and lubricants (excluding
fuel) is approximately 1900 pounds (average).
FAA Certificated Maximum Gross Weights
Standard configuration landing gear
3200 pounds
High skid gear and emergency floats
3200 pounds
External loads (up to 1500 pounds)
3350 pounds
Standard fixed floats
3000 pounds
Fuels
Capacity: 92.06 U.S. Gallons (91 US gallons useable)
ASTM Type Jet B (JP-4) for all temperatures
ASTM Type Jet A or A-1 (JP-5) above 0 °F
The following text does not apply if the aircraft is fitted with an airframe fuel
filter:
Below 0°F use a mixture of 80/87 AVGAS and ASTM Type Jet A and A-1.
Approved Lubricants
*Engine and power train gear cases oils must conform to MIL-L-7808 and MILL-23699 or turbine oil 555.
Engine:
5.5 U.S. Quarts
Transmission:
5.0 U.S. Quarts
Tail Rotor Gearbox:
3/8 Pint
*The group number oil used in the transmission must be used in the engine. It is
possible for the seals between the engine and the transmission to fail. If oils with
a different group number were to mix; this would cause a lot of damage.
Keep the engine oil reservoir topped up. If there is a sudden drop in oil level it
could indicate a failure of the “labyrinth” seal between the engine and the
transmission.
Hydraulic System:
1 Pint – Reservoir - 1 Pint in System
General purpose grease
MIL-G-81322
Labyrinth Seal
Due to the high temperatures and speed of rotation of the bearings in the engine,
normal seals wear out too quickly and allow oil to leak out. For this reason, a
labyrinth seal was developed to keep the oil in place. This seal does not make
contact with the shaft. There is a very small gap. The oil is prevented from
leaking out by an opposing air pressure on the other side of the bearing.
Main Rotor
Number of Blades:
2
Diameter:
33’ 4”
Chord:
1’ 1”
Disc Area:
873 Sq. Ft.
Twist:
-10 Degrees
Engine to Rotor Gear Ratio:
15.228 to 1
Tip Speed”
688’/Sec.
RPM:
395 = 100%
Tail Rotor (65 Inch Dia.)
Number of blades:
2
Diameter:
5’ 5”
Chord:
5.27”
Disc Area:
23.05 Sq. Ft.
Twist:
0 Degrees
Tip Speed:
723’/Sec.
RPM
2553
Tail Rotor (62 Inch Dia.)
Number of Blades:
2
Diameter:
5’ 2”
Chord:
5.27”
Disc Area:
20.97 Sq.Ft.
Twist:
0 Degree
Tip Speed:
690’/Sec.
RPM:
2553
Airframe Dimensions
Overall Length:
39’ 1.0”
Overall Height: 9’ 7.5”
Tread:
6’ 3.5”
Cargo Area
Length (Over-all):
39.0”
Width (Floor Level):
46.75”
Height (Maximum):
34.25”
Cargo Door Opening
Height:
35.75”
Width:
35.00”
Cargo Area Volume
Main Cargo Area (Rear passenger compartment):
Baggage Compartment:
16 Cu. Ft.
40 Cu. Ft.
3
AIRFRAME
Airframe (General)
The fuselage consists of three main sections: the FORWARD SECTION which
extends from the cabin nose to the bulkhead aft of the passenger compartment,
the INTERMEDIATE SECTION which extends from the bulkhead aft of the
passenger compartment to the tail boom, and the TAILBOOM SECTION.
Forward Section
The forward section consists of an aluminum honeycomb structure and provides
the major load carrying elements for the forward cabin. It provides for the pilot
and passenger seating, the fuel cell enclosure (underneath the rear passenger
seats), flight controls and aircraft instrumentation. The cabin area accommodates
up to five persons or 40 cubic feet of cargo if the rear passenger seats are
removed. Floor loading in the cabin area is limited to 86 pounds per square foot.
The crew and passenger doors each have a wedge shaped window of blue tinted
plastic with the lower forward section as a sliding window for ventilation. This
sliding window is adjustable and the window handles function not only as a
handle to open but also as a retainer to keep the window from sliding out of the
track. The windshields are made of tinted acrylic plastic and are supported by
formed aluminum alloy sections. Take great care when cleaning these wind
screens as they are very easily scratched. Never use paper towel to clean – only
wet or damp soft chamois leather or cloth. Two blue tinted transparent lower
windows made of plastic are located in the lower cabin nose section. Two cabin
roof windows (skylights) are provided in the roof of the forward compartment.
The cabin panel window is located on the right side of the helicopter, and it is a
fixed window made of tinted plastic.
The helicopter is equipped with four entrance/exit doors for crew and
passengers, as well as a variety of hinged doors and panels, which provide
access for inspection and servicing. Crew and passenger doors are located on
both sides of the fuselage and are removable for aerial work. If the doors are
removed for flight there are Vne restrictions to be observed. Some models may
be equipped with emergency door release handles on the inside and passengers
should be briefed on their use.
A crew door is installed on each side of the forward fuselage to provide access to
the cockpit area. Each door is equipped with a latch assembly, which may be
operated from either side of the door, and a lock installed in the exterior door
handle.
The airspeed restrictions for flying with the forward doors removed are due to
possible control problems when the center of gravity is too far aft. The helicopter
flight manual refers to “cyclic fore and aft stick reversal”. This does not mean
that the cyclic operates in reverse. What it means is that the expected
acceleration or deceleration normally experienced when moving the cyclic, does
not occur due to the disturbed airflow past the helicopter (especially over the
horizontal stabilizer.) This problem was associated with the 206A but the
restriction still remains in the helicopter flight manual for all models.
A passenger door is installed on each side of the fuselage to provide access to the
cabin area. Each is equipped with a latch assembly, which may be operated from
either side of the door, and a lock installed in the exterior door handle. The doors
should not be slammed shut. Light pressure is all that is required to close the
doors. If you have trouble closing the doors, try opening one of the sliding
windows to equalize the air pressure during closing.
Door locks are installed in the crew, passenger, and baggage compartment doors
to provide security to the cockpit, cabin, and baggage areas. Locks are similar in
construction, and are keyed alike, except for the baggage compartment door lock
which has a separate key.
The battery access door, located on the nose of the helicopter, provides access to
the battery, battery relay, TOT light reset key switch, hour meter circuit breaker
and hour meter. The access door is hinged aft, and two camloc fasteners secure
the forward edge of the door to the fuselage. The battery has a quick disconnect
handle. If the helicopter is not going to be used for a period of time it is a good
idea to disconnect it.
Doors and panels are provided at various locations in the cowling and fairings
for servicing and inspection of the interior areas. The oil reservoir access door
and the oil cooler access door are located on the aft fairing. The engine cowling
has side panels which are hinged for easy access, and the air induction cowling
has doors on both sides for inspection of the transmission area. Most
miscellaneous access doors open on piano hinges and are secured with flush
latches and wing head stud fasteners.
The skid landing gear consists of two skids attached on ends of two arched cross
tubes that are secured to the fuselage by means of four strap assemblies. Each
skid tube is designed with a forward end step, a tow fitting, two saddle fittings
for cross tube attachment, three separate shoes along the bottom (for abrasion
protection), a rear cap, and two eyebolt fittings for mounting of ground handling
wheels. The aerofoil type skid landing gear fairings (optional) enclose the
forward and aft cross tubes. These fairings are constructed from white
thermoplastic with aluminum alloy stiffeners and supports. Make sure that the
skids and skid attachments are inspected thoroughly for cracks or abrasion
before flight.
When the optional high skid gear is installed (as shown above), the front
mounting hardware must include the spacer bar, eight springs, and two lower
supports per side. This additional attachment hardware will prevent medium
frequency vibrations from being transmitted from the landing gear to the
airframe.
It is recommended that no components be attached to the landing gear assembly
except as designated by the manufacturer. Unauthorized attachments may lead to
failure of the cross tube.
Intermediate Section
The intermediate section is formed from an aluminum, semi-monocoque
construction and provides a deck for the engine installation, baggage
compartment, and a compartment under the engine deck for a heater and
electrical equipment. Older models had a separate heater. The Bell 206 B3 uses
bleed air from the engine compressor for heating.
A titanium engine pan, located below the engine, acts as both a drip pan and
firewall. It provides clearance to remove accessories without having to remove
the engine. The engine pan along with the titanium fore and aft fire- walls are
fitted within the top part of this section to form the engine compartment.
Located directly below the engine area is the baggage compartment. This
compartment has a volume capacity of 16 cubic feet and can carry a maximum
of 250 pounds not to exceed 6 pounds per square foot. 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; and it is secured by means of two
pushbutton latches and a keyed lock. Within the compartment are tie-down loops
to secure cargo and equipment.
There have been instances of the baggage door not being properly latched and
opening in flight. I know of one accident that was caused by a coat being sucked
out of the baggage compartment and striking the tail rotor; causing the helicopter
to have a very hard landing. I now always fly with the door latched and locked.
Tail boom Section
The tail boom is a full monocoque structure except for the forward 10 inches
where the loads are redistributed by means of four intercostal load-carrying
members.
The tail boom is attached to the aft fuselage by four bolts. A horizontal stabilizer,
vertical fin, and fairings are installed to provide stability and to protect and cover
the tail rotor drive shaft and tail rotor gearbox. Modifications of the tail boom,
including installation of antenna, lights or equipment that change the mass or
mass distribution of the helicopter, are not permitted without the proper
approval.
The horizontal stabilizer is constructed of aluminum honeycomb. The stabilizer
is an inverted aerofoil that provides a downward force on the tail boom to
maintain the cabin in a nearly level attitude throughout the cruise airspeed range
and to aerodynamically streamline the fuselage to reduce drag. The horizontal
stabilizer has a spoiler on the leading edge. This spoiler is designed to cause
turbulence during autorotation and disturb the lift produced by the stabilizer thus
allowing the fuselage to stay in a stable attitude.
The vertical fin, constructed primarily of aluminum and honeycomb material,
provides directional (yaw) stability. The fin is canted 5.5 degrees outboard to
reduce tail rotor thrust requirements during forward flight. It contains a top
fairing to mount the anti-collision light and a rubber bumper and tail skid to
protect the tail rotor and warn the pilot of a tail-low attitude.
When the aircraft is being ground handled, it is tempting to place your hand on
the lower part of the tail boom. This may lead to indentations on the thin
aluminum covering and should be avoided.
The tail rotor gearbox fairing encloses the tail rotor gearbox and is attached to
the tail boom and vertical fin. It incorporates a white position light and screens.
The screens provide cooling airflow for the tail rotor gearbox. The oil level in
the gearbox is visible through one of the screened openings.
Cowlings
1. Battery access door
2. APU access door
3. Oil reservoir access door (right side only)
4. Oil cooler access door (left side only)
5. Engine side cowling
6. Baggage compartment (left side only)
7. Air induction cowling door
Obviously it is important to know where the access panels are. After completing
your walk round you should get into the habit of making one final inspection of
the helicopter to make sure all these panels are secure before getting into the
helicopter. These access panels can cause a lot of damage in flight if they are not
secured and may even break off and go into the tail rotor in flight if not properly
secured.
The forward, transmission, air induction, engine, and aft fairing cowlings are
constructed of aluminum alloy, fiberglass, or honeycomb materials. Access
doors are provided with flush-type latches that permit inspections and servicing
without removing the cowling.
The air induction cowling covers the aft portion of the main transmission. Inlet
ducts are provided on each side of the air induction cowling to direct the airflow
into the induction screen. Two doors are also provided for maintenance and
inspection of the induction screen and main transmission. The induction screen
is located inside and attached to the air induction cowl and the forward side of
the forward firewall. The optional particle separator (discussed later) is visible
here.
The engine cowling, which includes the engine upper cowling and engine side
cowling, is constructed of aluminum alloy. It contains two screen vents forward
of the cut-out for the exhaust stack. Hinged to the engine upper cowling are two
engine side cowlings. The side panels are constructed of aluminum alloy and
incorporate five screen vents, three flush latches, and a camloc fastener. The
right engine side cowling includes a screen vent for the starter generator cooling
duct. Each engine side cowling may be hinged open and supported by a rod for
access into the engine compartment.
The aft cowling is constructed of aluminum alloy and incorporates two screen
vents, oil cooler and oil tank access doors, and a cut-out for the oil cooler blower
exhaust.
Particle Separator
The optional particle separator may be installed in place of the standard inlet
screen. Approximately 80% of all JetRangers come fitted with a particle
separator. The particle separator provides continuous protection for the
compressor against damage from the ingestion of sand, dust, and other foreign
material that would seriously abrade the compressor blades in the engine. The
unit consists of the separator, bleed air tubing and hoses, compressor wash
fittings, and other required hardware.
The separator section has many filter elements and is positioned so that all inlet
air must pass through the filter elements before entering the engine.
Particles such as dust, dirt or sand, entering the filter elements, are spun in the
vortex generator and hurled centrifugally into the scavenge chamber. They are
ejected overboard by the venturi effect created by engine bleed air as it is
discharged through the ejector tubes. This scavenge flow, which carries the
particulate material with it, accounts for approximately 8% of the inlet air flow.
A window is installed on each side of the cowling to permit visual inspection of
the separator chamber. The ejector tubes are mounted on each side of the
cowling just below these windows.
The particle separator does not impose any aerodynamic restrictions and has no
effect on the maximum gross weight.
Flight in falling or blowing snow is not permitted as there is a high risk of the
particle separator becoming obstructed and reducing the airflow to the engine
which may cause an engine flame-out.
4
CREW COMPARTMENT
Crew Compartment
The pilot’s station is situated on the right side of the crew compartment. The left
side can accommodate a passenger, or a co-pilot if the optional dual controls are
installed.
The instrument panel is mounted on a center pedestal forward of the two crew
seats and is tilted forward at a five degree angle for maximum visibility. Flight
instruments are located on the right side of the panel and system instruments are
in two vertical rows to the left of the flight instruments. Caution and Warning
Panel lights are mounted just below the glare shield across the top of the
instrument panel.
The pedestal extends aft from the instrument panel between the seats to form a
console for the radios. The overhead console is centrally positioned aft of the
windshields on the cockpit ceiling. This console contains most of the circuit
breakers and electrical switches.
Instrument System
The instrument system is divided into four separate categories: flight, navigation,
propulsion, and miscellaneous. All indicators are installed in the hinged
instrument panel except the OAT gage, compass, and hour meter. The OAT gage
is mounted in the upper left hand corner of the pilot’s windshield. The magnetic
compass is mounted to the right side of the cabin structure slightly forward of
the instrument panel. This helps minimize any interference from the electrical
components. The hour meter is mounted in the battery compartment and is
visible behind the battery.
The flight instruments include the airspeed indicator, altimeter, turn and slip
indicator (optional) and vertical speed indicator (optional).
The primary navigation instrument on the basic 206 helicopter is the magnetic
compass. Most helicopters today are fitted with some form of GPS or radio
navigation aid but this is optional and at the discretion of the owner.
The propulsion instruments are the dual tachometer, gas producer tachometer,
engine torque meter indicator, turbine outlet temperature indicator, engine oil
pressure/temperature indicator, transmission oil pressure/temperature indicator,
fuel pressure indicator and fuel quantity indicator.
The miscellaneous instruments are the OAT gauge, clock, load meter, and the
hour meter.
Pitot Static System
The pitot tube is mounted on the most forward part of the cabin nose structure.
The tube supplies dynamic air pressure to the airspeed indicator. Static air
pressure for instrument operation is obtained from two non-heated static ports on
the left and right sides of the aircraft, aft of the cabin lower windows. Turning
the PITOT HEAT switch to the ON position activates the heater element for the
pitot tube. This tube becomes extremely hot when the pitot heat is used and care
should be used when applying the cover after flying. Internal and drain lines for
the pitot tube and static ports are made of nylon tubing.
Flight Instruments
Airspeed Indicator - The airspeed indicator is a standard pitot/static instrument.
This indicator provides an airspeed reading in knots (KTS) and miles per hour
(MPH) by measuring the difference between dynamic air pressure from the pitot
tube and the static air pressure from the static ports.
The indicator presents airspeed from 0 to 150 knots and is scaled in 5-knot
increments at 20 knots and above. A maximum autorotation speed blue line is
located at 100 knots and a maximum speed (Vne) red line is located at 130
knots.
Altimeter - The barometric pressure altimeter presents an altitude reading in feet
above mean sea level (MSL) based on the relationship between the static air
pressure and the barometric setting on the altimeter.
The barometric setting may be adjusted manually to reflect the current
barometric pressure corrected to sea level in inches of mercury or in millibars,
depending on the instrument installed.
Inclinometer - The inclinometer consists of a curved glass tube, ball, and
dampening fluid. The ball indicates when the helicopter is in a coordinated turn
or balanced straight and level flight, i.e. “in trim”, if the helicopter is in a slip or
skid, the ball will move off center.
Turn and Slip Indicator – (See also ‘Inclinometer’ photo). The optional turn
needle is powered by the 28Volt DC bus.
The turn needle indicates the rate at which the aircraft is turning about the
vertical axis degrees per second. The ball (slip) principal is the same as the
inclinometer ball.
Navigation Instruments
Magnetic Compass - The magnetic compass is a standard, non-stabilized,
magnetic type instrument mounted on a support attached to the pilot side of the
forward cabin next to the door post. A compass correction card is located below
the instrument.
All other navigation instruments are not part of the basic aircraft. See the
avionics manufacturer’s documentation for applicable information relating to
these instruments.
Propulsion Instruments
Dual Tachometer Turbine/Rotor (N2/Nr)
The Dual Tachometer is equipped with two pointers, one displaying Main Rotor
RPM (Nr) on the inner scale and one that displays Engine Power Turbine RPM
(N2) on the outer scale. All scales are presented in percent of rated RPM.
The “T “(engine speed) indication is powered by the power turbine (N2)
tachometer generator that is mounted on the left front side of the engine
accessory gearbox case.
The “R” (rotor speed) indication is powered by the rotor tachometer generator
mounted onto the transmission lower case.
Gas Producer Tachometer (N1)
The gas producer tachometer (N1) displays engine gas producer (compressor)
speed in percent of rated RPM.
The outer scale of the gauge displays in 2% increments and the smaller dial scale
displays in 1% increments. The indication is powered by the gas producer
tachometer generator that is mounted onto the right front of the engine accessory
gear case.
NOTE
The rotor (Nr), turbine (N2), and gas producer (N1) tachometer generators are
self-generating and will operate without electrical bus power.
Engine Torque Meter Indicator
The engine torque meter is a wet oil gauge.
The system utilizes engine oil that has been metered by the torque meter gear
shaft assembly to indicate power output of the engine. The gauge displays in
percent torque.
At low altitudes this instrument will normally be the limiting instrument in flight
and should be monitored carefully.
Turbine Outlet Temperature Indicator
The Turbine Outlet Temperature (TOT) indicator displays engine gas
temperature in degrees Celsius. Four probes measure gas temperature between
the gas producer turbine and the power turbine.
The TOT gage is powered by the 28 V dc bus through the 3amp circuit breaker.
An optional TOT Light Test button is located on the left side of the instrument
panel.
At high altitudes this instrument will normally be the limiting instrument in
flight and should be monitored carefully.
Engine Oil Pressure/Temperature Indicator
The Engine Oil Pressure gauge is a wet oil gauge. The engine oil
pressure/temperature indicator displays oil pressure in PSI on the left side of the
instrument and oil temperature in degrees Celsius on the right side.
The engine oil temperature input signal is provided by a thermo-bulb installed on
the engine oil tank. The temperature gauge is powered by the 28 VDC bus
through the 3 amp ENG TEMP circuit breaker in the overhead panel.
Transmission Oil Pressure/Temperature Indicator
The transmission oil pressure/temperature indicator displays oil pressure in PSI
on the left side of the instrument and oil temperature in degrees Celsius on the
right side. The Transmission Oil Pressure gauge is a wet oil gage.
The transmission oil temperature input signal is provided by a thermo-bulb
installed on the transmission oil filter assembly. The temperature gauge is
powered by the 28 Volt DC bus through the 3 amp TRANS TEMP circuit
breaker in the overhead panel.
Fuel Quantity Indicator
The fuel quantity indicator displays total fuel quantity in US Gallons. The signal
is generated by the fuel quantity transmitter that is connected to fuel probes
located in the fuel system. The fuel quantity indicator is powered by the 28 Volt
DC bus through the FUEL QTY circuit breaker on the overhead panel.
NOTE
The low fuel warning light will come on at approximately 20 gallons. If one of
the fuel pumps fails, there are approximately 10 gallons of unusable fuel in the
tank. In my experience the helicopter uses approximately 27 gallons per hour but
you should use a fuel burn rate of 30 gallons per hour for planning purposes.
Miscellaneous Instruments
Fuel Pressure/DC Load meter
The Fuel Pressure/DC Load meter displays fuel pressure in PSI on the right side
of the instrument and DC current load on the generator on the left side of the
instrument. Fuel pressure is derived from a pressure-operated potentiometer
(transducer) that measures the output of the boost pumps. The gauge is powered
by the 28 Volt dc bus through the 3 amp FUEL PRESS circuit breaker in the
overhead panel.
The load meter indicates the load (in percent) that is supplied to the 28Vdc bus
by the generator. The load meter signal is from the load meter shunt located in
the electrical equipment shelf (behind the rear passenger compartment). The
circuit is completed through two circuit breakers in close proximity to the shunt.
The load meter will show a high load (50 to 70%) on startup but when the
generator is turned on it should immediately start to reduce to around 10 to 20 %
if the battery is in good condition. If it remains high during flight then there is
possibly of a short circuit somewhere. If after a few minutes the load meter is
still showing 40% or higher, then the voltage regulator is putting out too little
voltage or the battery is weak.
Hour meter
The hour meter (or “DATCON”) is mounted on the aft bulkhead of the battery
compartment.
The hour meter is powered by the 28 Volt dc bus through a 1 Amp remote circuit
breaker that draws power from the CAUT circuit breaker. The system is
activated when the N1 is greater than 55%.
This meter does not start to read until the collective is raised. Remember that if
you raise the collective just a little while the engine is running on the ground; the
datcon will start to count.
Another type of hour meter you may come across is the HOBBS meter. This
works the same way as the DATCON meter.
Caution and Warning System
The caution and warning system includes the Caution Warning Panel (CWP),
engine rpm sensor and warning horn, low rotor rpm sensor and warning horn,
associated components, and interconnecting wiring.
The Caution Warning Panel is powered by the 28 Volt dc bus and protected by a
5 Amp circuit breaker. Each segmented indicator light is in series with its
respective system. The Warning lights are colored RED while Caution lights are
colored AMBER. Pressing the CAUTION LT TEST button will illuminate all
the lights for bulb testing.
Switches and Circuit Breakers
The center pedestal contains switches that allow the pilot to activate certain
aircraft systems. Switches on this panel include the Caution Light Bright/Dim
switch, the Engine Anti-icing switch, and the Hydraulic System switch.
When the INST LT rheostat (on the overhead console) is ON, positioning the
BRIGHT/DIM switch momentarily to the BRIGHT or DIM position can dimly
or brightly illuminate the caution panel lights. The caution lights will remain at
the selected intensity until the INST LT rheostat is turned off or another position
is selected.
The engine anti-ice solenoid valve is powered from the 28 Volt dc bus through
the 5 amp ENGINE ANTI-ICE circuit breaker and the ENGINE ANTI-ICE
switch. When the solenoid valve is de-energized (ON), hot air passes from the
compressor diffuser through the anti-ice valve to the engine inlet housing. This
hot air aids in preventing ice formation on the hollow inlet guide vanes. Power is
only provided to the solenoid when it is positioned to OFF In the event of total
electrical failure, the system is fail-safe ON. Note that this is not a de-icing
system.
The hydraulic system switch controls the operation of the hydraulic bypass
solenoid. This switch should be in the on position for normal helicopter
operations. The solenoid is normally open when de-energized. The switch
provides power to the solenoid in the OFF position only. In the event of total
electrical failure, the system is fail-safe ON.
Lighting Systems
The lighting system includes both interior and exterior lighting. The interior
lighting system includes the cockpit map reading light, instrument and control
panel lighting, and cabin lighting. The exterior lighting system includes position,
anti-collision, and landing lights.
Interior Lighting System
The cockpit map reading light features a narrow spotlight or wide floodlight
beam. It is a multipurpose utility light designed to provide blue or white
illumination and is protected by its own circuit breaker. The 28 V dc power is
supplied through the CKPT LIGHTS 5 amp circuit breaker located on the
overhead console.
Instrumentation lighting
28 V dc power is supplied through the INSTR LIGHTS circuit breaker located
on the overhead console.
The INST LT rheostat knob, also located on the overhead console, adjusts light
intensity. Rotation of this knob operates the power OFF/BRT switch that
provides power to both the 28 V and 5 V dc lighting systems. The INST LT
rheostat also enables the BRT/DIM switch (on the center pedestal) to operate.
Exterior Lighting System And Position Lights
The exterior lighting system includes position lights, anti-collision light, landing
lights, and related wiring and components. The anti-collision light is mounted on
the top portion of the vertical fin. Position lights are located on the outboard
sides of the horizontal stabilizers (left is red and right is green) and on the tail
rotor gearbox (white) fairing.
Dual fixed landing lights
Located in the lower portion of the helicopter nose section and are controlled by
the landing light switch located on the collective switch box.
Flight Controls and Switches
Cyclic Control and Cyclic Switches — The cyclic control provides a means for
the pilot to provide tilt input from the flight station to the rotor system. The
mechanical connections will be discussed in the Flight Control Section of this
book.
The cyclic control incorporates a series of switches that provides the pilot
electrical connection access to specific aircraft systems. A two position transmit
switch is mounted on the forward side the cyclic grip. Depressing the switch to
the first detent position will activate internal communication system (ICS)
between the pilot and passengers. Depressing the switch to the second detent
allows for the pilot to transmit on the radio frequency selected on the audio
panel. Additional switches are used to control additional optional equipment that
may be installed on the aircraft, (e.g. cargo hook).
Fuel Valve
A fuel valve switch is located on the lower left or right side of the instrument
panel and electrically operates a motor driven fuel valve providing a means of
shutting off fuel flow to the engine.
The FUEL VALVE switch is covered by a red locking device to prevent
inadvertent closing of the fuel valve during flight.
Collective Control and Collective Switches
The collective control provides a means for the pilot to provide angle of attack
change from the flight station to the rotor system. The mechanical connections
will be discussed in the Flight Control Section of this book. The collective
control incorporates a series of switches that provides the pilot electrical
connection access to specific aircraft systems. The collective switches are the
GOV RPM INCR/DECR switch, LDG LTS switch, and the START switch.
The GOV RPM INCR/DECR switch is a spring-loaded, momentary contact
switch that enables the pilot to change the governor rpm actuator setting. With
the switch in the INCR position, the actuator motor moves the arm and the
N2/Nr rpm increases. With the switch in the DECR position, the actuator motor
moves the arm in the opposite direction and decreases the N2/Nr rpm. With the
switch in the normal (centered) position, the circuit is de-energized. Do not keep
this switch held in the INCR or DECR position as this can jam the governor arm
linkage. The switch should be “beeped” only.
The landing light switch located on the left side of the switch box controls the
two fixed landing lights under the nose of the helicopter. The switch is a threeposition switch, OFF/FWD/BOTH. In the OFF position, the circuit from the 28
V dc bus to the lights is disconnected and the lights are extinguished. In the
FWD position, the most aft mounted forward facing light is energized while the
forward mounted downward facing light is de-energized. In the BOTH position,
both lights are energized.
The STARTER switch allows the pilot to engage the starter through the starter
relay, when power is applied to the helicopter electrical system.
Tail Rotor Pedals and Adjuster
The tail rotor pedals provides a means for the pilot to change the pitch of the tail
rotor blades.
The mechanical connections will be discussed in the Flight Control Section of
this book. Tail rotor pedals can be positioned by means of the pedal adjuster to
allow for individual comfort.
Ventilation System
Cockpit vent knobs - Air for cabin ventilation is obtained by opening the sliding
window in each of the entrance doors. Additional ventilation to the cockpit may
be obtained by pulling out the VENT control knobs and positioning the DEFOG
blower switch on the overhead console to the DEFOG (ON) position.
Ram Air - When the VENT control knobs below the instrument panel are pulled
out; ram air is forced into the air intake grills and directed through the plenum
and flapper valve. The flapper valve assembly is opened or closed with the
VENT control knob.
Defog - Two electrically driven axial flow blowers are installed in the inlet end
of the defroster nozzles. The blowers are controlled by a DEFOG switch on the
forward row 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.
Seating
Crew And Passenger Seats - All seat back cushions are constructed of soft
sponge with styrene backing. Seat cushions are constructed of soft urethane
foam. The back and seat cushions are bonded to tubular frames and upholstered
with cloth fabric or leather. A very small storage area is located behind each
crew seat. There is no provision for seat adjustment or under seat storage.
The rear seats are easily removed for cleaning or to increase storage space. They
are normally held in place by Velcro strips. The storage area is increased
significantly when the rear seats are removed.
Restraint Assemblies
Each crew and passenger seat is equipped with a restraint assembly that consists
of an inertial reel, shoulder harness, and an adjustable seat belt. The inertia reel
is provided with an anti-rebound lock feature and is capable of retracting 22
inches of web belt. The front seats have a four point harness and the rear seats
have a three point harness.
Make sure that any unused belts are buckled. If a door is closed and the buckle is
trapped outside; the resultant damage due to the buckle bouncing against the
fuselage can be considerable.
Flight Manual, Airworthiness Certificate and Aircraft
Registration
The Rotorcraft Flight Manual is used for operation of the helicopter and must be
on board the aircraft during operation. The airworthiness and registration
certificates are located in the document case on the right forward side of the
instrument panel console. In the USA these certificates are required to be carried
in the aircraft at all times.
Hand Held Fire Extinguisher
A manually operated fire extinguisher is located in the crew compartment
between the pilot and co-pilot seats. It is mounted on a quick opening bracket for
rapid removal in an emergency.
Check the Extinguisher for security and make sure the pressure gauge is reading
correctly. Some countries have an inspection date written on the extinguisher and
this date should be current.
Make sure you are familiar with the extinguisher mounting system and how to
use it in an emergency as fire extinguishers vary greatly in their design.
Rotor Brake
When the optional rotor brake is installed, application is limited to ground
operation after the engine has been shut down and Nr has decreased to 38 - 30%.
Engine starts with the rotor brake engaged are obviously prohibited.
The rotor brake is engaged by pulling the lever fully down and forward. It is
common to feel a shudder through the airframe as the blades are about to stop
and the brake pressure should be released to reduce this vibration.
Some pilots use only light pressure on the brake but this can lead to a build-up of
heat in the brake disc without causing any significant braking effect. I prefer to
use the brake lever in its fully forward position until the blades are about to stop.
5
ELECTRICAL SYSTEM
Description: The 206B III helicopter is equipped with a 28 volt direct current
electrical system. Power for this system is obtained from a nickel-cadmium 24
volt, 13 ampere-hour battery and a 150 ampere (derated to 105 ampere), 30 volt
combination starter-generator. Some helicopters may be fitted with a lead acid
battery. Major components of the DC power system include the battery, startergenerator, voltage regulator, relays, and circuit breakers. All circuits in the
electrical systems are single wire with a common earth return. The negative
terminals of the starter-generator and the battery are earthed to the helicopter
structure.
Switches: Switches used in controlling the various electrical units of the
helicopter are mounted on the overhead console panel. Circuit breakers for the
protection of electrical circuits against overloads are also located on the
overhead console panel. The panel is designed to provide easy access to these
units. The generator is provided with over voltage protection set at 31 ± 1 volt.
Battery System
The battery system includes the battery, a battery relay, and a battery switch with
related wiring. The nickel- cadmium 24 volt 17 ampere-hour battery is located in
the nose of the aircraft. As mentioned before, this battery may have been
replaced with a lead-acid battery on older helicopters.
NICAD batteries always have a metal case whereas the lead-acid batteries
normally have a plastic case.
During engine operation a fully charged battery can be determined by moving
the battery switch from BAT to OFF and observing the effect on the generator
load meter. If the change in indication is less than 1.0% on the load meter, the
battery is charged.
The battery relay, located in the nose section forward of battery, is an electrically
operated switch controlling battery current to the main bus bar. The relay is
actuated by the battery switch, located in the overhead console, which opens and
closes the circuit to the relay energizing coil.
Nickel-cadmium batteries are different from lead-acid batteries. The terminal
voltage remains constant over 90% of the total discharge time. This means that
the battery will produce a steady voltage for a longer period of time than a leadacid battery but whereas a lead-acid battery decays steadily; the nickel-cadmium
battery will lose its voltage suddenly. This would be very unfortunate during a
start sequence.
The battery is normally good for three or possibly more starts. If numerous starts
are required or the outside temperature is freezing or below, it is advisable to use
an auxiliary power unit (APU) or power pack to preserve battery life.
External Power System
The external power system includes the external power receptacle, external
power relay, and related wiring. The external power receptacle is located in the
lower front center of the nose section. It is controlled through the small positive
pin from the external power source that energizes the circuit to the activating coil
of the relay. The external power is connected to the Bus.
The 28 VDC Ground Power Unit (GPU) should be limited to 500 amperes or
less to reduce the risk of starter damage due to excessive heat. 350 to 450
amperes, 28.5 volts is recommended.
When external power is being used, the BATtery switch should be turned off.
This will prevent the battery from becoming overcharged and consequently
overheating. Make sure that the appropriate section in the start-up check list is
used during the start-up procedure.
External power may be supplied to the helicopter by means of a receptacle
located on the lower front section of the helicopter fuselage. An auxiliary power
unit capable of delivering 28 volts DC with a 400 amperes rating is required.
Generator System
The generator is used to start the engine, charge the battery and supply power for
the operation of the dc equipment.
The generator is driven from the compressor drive train through the accessory
gearbox. This means that high generator loads will show a rise in Turbine Outlet
Temperature (TOT). For this reason, the generator is normally turned off during
power checks as it requires power from the engine.
The generator is located on the lower right aft side of the accessory gearbox and
provides power at a regulated voltage for all dc electrical components on the
helicopter. The generator switches onto the main bus when the generator switch
is on and generated voltage exceeds the voltage on the bus by 0.30 to 0.42 volts.
Reverse current relay connects the generator power to the bus when the voltage
regulator senses adequate generator voltage.
The over-voltage sensing relay and the reverse current relay provide protection
against over-voltage and reverse-current conditions.
The voltage regulator compensates for voltage fluctuations caused by the
varying load conditions.
The duct leading from the generator to the fairing is actually used to draw air
into the generator for cooling.
Starter/Igniter System
The starter/igniter system includes the starter portion of the starter/generator,
starter relay, generator field control relay, igniter and starter switch.
The starter/generator is energized by depressing the starter button on the
collective to start the engine.
Direct current is supplied to the starter through the starter relay when the start
button is pressed. Power also goes to the igniter in the combustion chamber of
the engine and provides a continuous ignition arc during the engine start cycle.
The starter heats up very quickly during the startup sequence. Because of this
there is a limit on the number of starts you may have in quick succession to
avoid overheating as described below:
Using the aircraft battery you should limit the starter energizing times as
follows:
Attempt 1:
Starter ON for 40 seconds then wait a minimum of 60 seconds.
Attempt 2:
Starter ON for 40 seconds then wait a minimum of 60 seconds.
Attempt 3:
Starter ON for 40 seconds then wait a minimum of 30 minutes.
Overhead Console
The overhead console is centrally positioned aft of the windshields on the
cockpit ceiling. This console contains most of the aircraft electrical system
circuit breakers and electrical switches and is easily accessible from the pilot’s
station. The circuit breakers are installed to protect individual systems from
excessive amperage. Circuit breakers may be either the pop out or switch style.
The electric toggle switches are designed to connect/disconnect specific systems
to the 28 VDC bus.
GEN SWITCH
The GEN switch is a three-position switch that controls a normally open Line
Control Relay that connects the generator to the 28 Volt DC bus. Switch
positions are OFF, GEN (ON), and RESET. When the GEN switch is in the OFF
position, the Line Control Relay is de-energized, interrupting the circuit from the
generator to the 28 Volt DC bus. If electrical power is applied to the bus through
the battery or external power with the GEN switch in the OFF position, the GEN
FAIL caution light will be illuminated on the caution warning panel. When the
switch is selected to the GEN (ON) position with the electrical system operating
normally, the Line Control Relay is energized closed, completing the circuit and
connecting the generator to the bus and extinguishing the GEN FAIL caution
light. The RESET position is a momentary spring-loaded position that allows the
operator the attempt to reset the generator should a generator malfunction occur.
This RESET feature is used in conjunction with the 10 amp GEN FIELD and 5
amp GEN RESET circuit breakers located immediately aft of the GEN switch on
the overhead panel.
BAT SWITCH
The BAT switch is a two-position switch that controls a normally open Battery
Relay in the nose of the aircraft that connects the battery to the 28 Volt DC bus.
Switch positions are OFF, and BAT. When the switch is in the OFF position, the
Battery Relay is de-energized, interrupting the circuit from the Battery to the 28
Volt DC bus.
When the switch is selected to the BAT (ON) position, the Battery Relay is
energized closed, completing the circuit and connecting the battery to the bus.
DEFOG BLOWER SWITCH
The DEFOG BLOWER switch is a two-position, 5 amp switch that connects the
electric blower motor to the bus. When selected to the aft position the circuit is
de-energized and the blower motor is OFF. When selected forward, the circuit is
completed, connecting the blower motor to the bus, blowing ambient air on the
windscreen to assist in fog removal.
PITOT HEAT SWITCH
The PITOT HEAT switch is a two-position, 5 amp switch that connects the
electric heating element in the pitot tube to the bus. When selected to the aft
position the circuit is de-energized and the pitot heat is OFF. When selected
forward, the circuit is completed, connecting the pitot heat to the bus, heating the
element to prevent ice formation and removing moisture inside the pitot tube.
POS LT SWITCH
The POS LT switch is a two-position, 5 amp switch that connects the five
aircraft position lights to the bus. When selected to the aft position (OFF) the
circuit is de-energized and the position lights are OFF. When selected forward
(ON), the circuit is completed, connecting the position lights to the bus. Position
lights are located on the horizontal stabilizers, and on the tail rotor gearbox
fairing.
ANTI-COLL LT SWITCH
The ANTI-COLL LT switch is a two-position, 5 amp switch that connects the
anti-collision light on top of the vertical fin to the bus. When selected to the aft
position (OFF), the circuit is de-energized and the anti-collision light is OFF.
When selected forward (ON), the circuit is completed, connecting the anticollision light to the bus.
INST LT RHEOSTAT
The INST LT rheostat controls the brightness of the background lighting on the
instrument panel and overhead console. This rheostat is also used along with the
caution and warning panel BRIGHT/DIM switch on the miscellaneous control
panel (center pedestal) to control the brightness of the caution and warning
lights.
Electrical Emergencies and Malfunctions
Battery Hot Warning Light — The 17 amp/hour NICAD battery incorporates
an internal thermal switch that will illuminate the warning light if the internal
battery temperature reaches 600 C (140° F). It is possible that the battery relay is
sticking. Turn the BAT switch OFF and LAND AS SOON AS POSSIBLE. Do
not attempt a restart using this battery.
Battery Temp Warning Light— The Battery Case temperature has reached
130°F or higher. Move the battery switch to the OFF position until the battery
cools (light extinguishes) then turn the battery switch ON. If the light keeps
illuminating there could be a fault with the battery. The battery should be
examined by an engineer at the earliest opportunity.
Electrical System - Complete Electrical Failure
Electrical power for flight is supplied by the starter/generator which is utilized as
a generator after the start has been accomplished. Evidence of the main
generator failure will be provided by observing the load meter gauge. There is no
provision for standby operation in the event of generator failure. Necessary
power can be supplied by the battery for short periods of time in the event of the
generator failing.
Cycle generator switch – if power is not restored – OFF
All electrical equipment – OFF (to conserve battery power)
Required electrical equipment – ON, only as needed.
The following will continue to operate with the battery and generator switched
off (this does not apply to digital instruments):
·
Turbine Engine continues to run normally,
·
Engine Oil Pressure gauge
·
Transmission Oil Pressure gauge
·
Torque meter
·
N1 Tachometer
·
N2 Tachometer
·
Nr Tachometer
·
Altimeter
·
Airspeed Indicator
·
Vertical Speed Indicator
·
Inclinometer
·
Compass
WARNING
Reduce altitude to below 6000 feet pressure altitude if flight permits because of
possible loss of fuel boost pumps.
6
FUEL SYSTEM
System Description
Helicopters S/N 3567 and subsequent
The fuel system uses a single bladder type, crash resistant, fuel cell. The fuel cell
has a useable capacity of 91 U.S. gallons (344.44 litres useable).
Installed within the fuel cell are two electrically operated boost pumps, lower
and upper tank indicating unit and an electrically operated sump drain valve. The
boost pumps are interconnected and supply fuel through a single hose assembly
to the fuel shutoff valve and from the shutoff valve to the airframe mounted fuel
filter. Boost pumps incorporate pressure switches in discharge ports and drain
plugs in the pump drain ports. The fuel cell is filled from the right side.
The fuel cell is filled through a filler cap located on the right side of the
helicopter, aft of the rear door. An overboard vent line is located on top of the
fuel cell.
Fuel capacity is registered on the instrument panel gauge. The helicopter can use
JP-4 or JP-5 fuel. Standard precautions should be observed during all fueling or
defueling operations.
Modern Fuels
The most common fuel is a kerosene/paraffin oil-based fuel classified as JET A1, which is produced to an internationally standardised set of specifications. In
the United States only, a version of JET A-1 known as JET A is also used.
The only other jet fuel that is commonly used in civilian aviation is called JET
B. JET B is a fuel in the naptha-kerosene region that is used for its enhanced
cold-weather performance. However, JET B's lighter composition makes it more
dangerous to handle, and it is therefore restricted only to areas where its coldweather characteristics are absolutely necessary.
Both JET A and JET B can contain a number of additives:
Antioxidants to prevent gumming usually based on alkylated phenols, eg. AO30, AO-31, or AO-37;
Antistatic agents, to dissipate static electricity and prevent sparking; Stadis 450,
with dinonylnaphthylsulfonic acid (DINNSA) as the active ingredient, is an
example
Corrosion inhibitors, eg. DCI-4A used for civilian and military fuels, and DCI6A used for military fuels;
Fuel System Icing Inhibitor (FSII) agents, eg. Di-EGME; FSII is often mixed at
the point-of-sale so that users with heated fuel lines do not have to pay the extra
expense;
Military forces around the world use a different classification system of JP
numbers. Some are almost identical to the civilian named fuels and differ only
by the amounts of a few additives:
JET A-1 is similar to JP-8
JET B is similar to JP-4.
Other military fuels are highly specialized products and are developed for very
specific applications. JP-5 fuel is fairly common, and was introduced to reduce
the risk of fire on aircraft carriers. Other fuels were specific to one type of
aircraft. JP-6 was developed specifically for the XB-70 Valkyrie and JP-7 for the
SR-71 Blackbird. Both these fuels were engineered to have a high flash point to
better cope with the heat and stresses of high speed supersonic flight. One
aircraft-specific jet fuel still in use by the USAF is JPTS, which was developed
in 1956 for the Lockheed U-2 spy plane.
Jet fuels are sometimes classified as kerosene or naphtha-type. Kerosene-type
fuels include Jet A, Jet A1, JP-5 and JP-8. Naphtha-type jets fuels include Jet B
and JP-4.
Jet A
Jet A is the standard jet fuel type in the U.S. since the 1950s and is only
available there. JET A is similar to JET-A1, except for its higher freezing point
of -40 °C. JET A-1 has a fairly high flash point of 38 °C, with an auto-ignition
temperature of 410 °F (210 °C). Jet A can be identified in trucks and storage
facilities by the UN number, 1863, Hazardous Material placards. Jet A1 trucks,
storage tanks and pipes that carry Jet A1 will be marked with a black sticker
with a white "JET A1" written over it, next to another black stripe. Jet A1 will
have a clear to straw colour if it is clean and free of contamination. Water is
denser than Jet A1, and will collect on the bottom of a tank. Jet A1 storage tanks
must be sumped on a regular basis to check for water contamination. It is
possible for water particles to become suspended in Jet A1, which can be found
by performing a "Clear and Bright" test. A hazy appearance can indicate water
contamination beyond the acceptable limit of 30ppm (parts per million).
The U.S. commercial fuels are not required by law to contain antistatic additives,
and generally do not contain them.
Range Extenders
The JetRanger B3 does not need a range extender as the filler opening is
positioned slightly higher than in earlier models.
In older models of JetRanger, the fuel capacity is 76 US gallons. This can be
increased to 96.7 US gallons by fitting a range extender. If the range extender is
used to carry this extra fuel, the minimum pilot weight is 200 lbs.
The use of the range extender has never been officially approved by Bell. Filling
the fuel tank to capacity uses up the airspace normally used for expansion. The
tank has never been drop tested when filled with this extra fuel.
The thermal relief valve prevents excessive pressure building up in the fuel line
above.
The pressure switch will detect a boost pump failure.
The check valve prevents fuel being pumped through that particular pump
should it fail.
Fuel Pump and Filter Assembly
The engine fuel pump and filter assembly are integral units mounted on the aft
end of the engine on the port side. Fuel enters the engine fuel system at the inlet
port of the pump and passes through the filter before entering the gear elements
of the pump. The fuel filter is monitored by a pressure differential switch located
on the lower firewall and connected electrically to the fuel filter caution light.
This light will illuminate if the filter becomes blocked.
Fuel Boost Pumps
Two electrically operated fuel boost pumps are located in the bottom of the fuel
cell. Do not confuse these pumps with the engine driven fuel pump. The pumps
are interconnected and supply fuel through one supply line. The pumps are
equipped with a check valve and thermal relief valve, pump drain port, seal drain
port, intake screen and pump operating pressure switch located in the discharge
port of the pump. The pumps are protected by circuit breakers located in the
overhead console.
Each pump is capable of supplying the engine independently of the other. Should
one pump fail, the engine will not flame out.
The helicopter flight manual requires that the boost pumps should be on for all
normal operations and they should not normally be turned off. Many pilots pull
the fuel boost circuit breakers out after shutting down the engine but it is not
required. Remember – they are circuit breakers – not switches.
Older models of JetRanger used to be run on JP4 fuel in some countries. This
fuel had a vapor pressure of 6000 feet. It was for this reason that the 6000 feet
warning still remains in the flight manual because if the fuel boost pumps failed
above this height then there would not be enough pressure to supply the engine.
If the generator should fail in flight, you should descend below 6000 feet
pressure altitude. This is because if your battery fails and you are still above
6000 feet, you will lose the boost pumps and the engine could stop. If you are
below 6000 feet, the engine driven fuel pump will continue to supply fuel to the
engine.
Fuel Shutoff Valve
A motor operated fuel shutoff valve incorporating a thermal relief feature is
installed in the main fuel supply line and is located in the fuel compartment
above the fuel filler cap. The valve is electrically controlled by an ON-OFF
switch located on the instrument panel and protected by a circuit breaker located
in the overhead console panel. In the event of electrical failure the valve will
remain in the position selected before the failure.
Fuel Pressure Transducer
The fuel pressure transducer provides a means of monitoring the fuel pressure.
The transducer is located on a fitting on top of the aft right side of the fuel cell
and is not normally accessible.
The aircraft must never be flown with the cover panel for the fuel pressure
transducer removed.
Fuel Quantity – Indicating Units
Two float type fuel level transmitting units (tank units) are installed in the fuel
cell. The lower unit is mounted in the tank bottom and monitors the fuel level up
to the horizontal surface of the cell, under the seat; the upper unit monitors the
fuel level in the upper section of the fuel cell, behind the seat and is mounted to
the top of the fuel cell. Both of the indicating units are connected to a common
quantity indicator.
It should be noted that the fuel bladder is attached to the airframe by screw
fasteners and if any of these should become detached then the bladder may
become partially collapsed. This could decrease the volume of the tank thus
leading to erroneous fuel quantity readings.
Airframe Fuel Filter
An Airframe fuel filter is mounted to the right side of the forward firewall on
helicopters S/N 3387 and subsequent. (Do not confuse this filter with the fuel
filter at the engine driven fuel pump on the left side.) The fuel filter consists of a
replaceable filter element, drain valve, bypass valve, impending bypass switch
and a manual test button. The airframe mounted fuel filter assembly eliminates
the requirement for adding anti-icing additive to the fuel supply when the
temperature is below 40°F (4°C). Indication of an impending bypass lights the
A/F FUEL FILTER segment on the pilot caution panel.
When the BAT switch is ON, the A/F FUEL caution warning light can be tested
by pressing the red button on top of the filter.
NOTE
The filter element must be replaced before the next flight when the caution light
comes on during engine operation.
Fuel Low Light
A float type low-level detector is mounted in the fuel cell and is calibrated to
activate a FUEL LOW caution light on the caution panel when approximately 20
gallons of fuel remain in the tank. If one or both of the fuel boost pumps fail,
the last 10 gallons is unusable.
Fuel Sump Drain Valves
A fuel drain switch provides a means of draining fuel from the cell. A sump
drain valve is located in the fuel cell and is electrically operated by means of the
drain switch located aft of the right- side passenger door. The drain switch is
wired through the fuel valve switch on the instrument panel. The fuel valve
switch must be in the OFF position to operate the drain valve.
Fuel Drain Procedures
FUEL SUMP - Drain fuel sample as follows:
a) FUEL BOOST circuit breakers - OUT
b) BATtery switch — ON.
c) FUEL VALVE switch — OFF.
d) FUEL DRAIN button — Depress, drain sample, and then release.
A/F Fuel filter - Drain and check before first flight of the day as follows:
a) FUEL VALVE switch — ON.
b) FUEL BOOST circuit breakers — IN.
c) CAUTION LIGHTS circuit breaker — IN.
d) Fuel filter drain valve — Open, drain sample, then close.
NOTE
Filter test button is located on top of fuel filter.
Fuel filter test button — Depress and check FUEL FILTER caution light is on.
Release button and check light out.
FUEL VALVE switch — OFF.
BATtery switch — OFF
Emergency Operation
There are two caution lights associated with the fuel system:
Fuel Pump — When the fuel pressure falls below 35 PSI, the fuel pump caution
light will illuminate. When a fuel boost pump is inoperative there are
approximately 10 US gallons of unusable fuel available.
In the case of a single boost pump failure, pressurized fuel continues to flow to
the engine from the remaining boost pump. The pilot should descend below
6000’ Hp and land as soon as practical as the engine will not receive pressurized
fuel if the other boost pump fails while above 6000’ Hp and has a high risk of
flame out.
NOTE
If the generator fails in flight you should also descend below 6000 feet pressure
altitude as battery depletion will cause the fuel boost pumps to stop working.
Fuel Filter – If installed, this caution light will illuminate if the engine fuel filter
(on the port side) becomes blocked. Land as soon as practical. The filter must be
cleaned or replaced before further flight.
7
POWERPLANT
Powerplant (General)
The working cycle of the turbine engine is similar to that of the piston engine.
There is induction, compression, combustion and exhaust. The major difference
is that unlike the piston engine, the turbine engines cycles are continuous. Unlike
airplane jet engines where the exhaust gases provide the thrust, the exhaust gases
for the helicopter turbo shaft engine are intercepted by a turbine that transfers the
energy from these gases to a gearbox that provides power for the helicopter.
It is vital that you learn not to over-torque or over-temp the engine as this can
have disastrous consequences for someone flying it weeks or even months later.
Misusing the engine will not necessarily cause a problem right away.
Starting the engine causes more stress to the components than just about any
other operation due to thermal shock and wear. For this reason, the start cycles
are counted and recorded in the technical log after each flight. As helicopters
tend to fly for short periods compared to airplanes, it is possible that the engine
could reach its start count limit before it reaches its useful life hour limit. Startups are expensive and so it is better to leave the engine running for ten minutes
while waiting for a passenger rather than shutting the engine down and then
starting up again.
Before starting the JetRanger, always check your battery condition (a volt meter
is very handy if it is installed). The engine relies totally on the battery for
startup. The engine has to be turning fast enough to keep enough cold air
flowing through the turbine otherwise it will become very hot and damage the
internal components. A weak battery could run out of power before the
combustion process becomes self-sustaining and this would be disastrous.
During startup it is vital that you keep your finger on the startup button (on the
collective) to keep the air flowing through the engine until it reaches the correct
speed. Even if something were to go wrong with the startup, you must keep the
engine turning to keep it cool by keeping your finger on the starter button.
A “Hung Start” can happen if the engine fails to accelerate and the RPM stays
constantly low. This uses up your battery power and if the battery runs out of
power then the engine slows down, less and less cold air is drawn through the
engine and subsequently the temperature in the combustion chamber becomes
really hot within a few seconds causing a lot of damage.
A “Wet Start” is the equivalent of a flooded engine and the igniter had failed to
light the fuel. Wait for at least 5 minutes and then vent the engine. Because the
battery has already been partially drained by the failed start and the venting, it is
probably a good idea to get an external start to make sure you do not run out of
power on your next start attempt.
If you follow the checklist exactly then you should have no problem starting the
engine every time. The most common reason I have seen for hot starts is low
time pilots starting the engine with the throttle already partially (or fully) open.
Therefore it is vital for you to double check that the throttle is fully closed before
pressing the starter button.
Should you have to abort the startup for any reason you must follow the
following simple steps:
Close the throttle to the fully closed position (you must press the idle release
button to do this).
Keep pressing the starter button.
If you have time, pull the igniter circuit breaker (above your head)
If you are in any doubt about the amount of power in your battery, you should
use an external start. It is better to be safe than sorry. Try to observe how quickly
the RPM rises and how quickly the igniter clicks. You will become familiar with
what to expect and so if some day your startup does not sound right or if the
RPM is slow to rise, you will have time to abort the startup procedure.
Monitor the fuel pressure as you press the startup button because if it drops to
near zero then your battery is weak.
Be particularly careful of starting aircraft that have not flown in a while.
Turbine engines take time to “Spool up” or “Spool down” when you make power
changes. For this reason you should operate the collective or throttle very
smoothly to prevent engine “surging”. Surging can happen when the airflow
over the compressor blades becomes disturbed causing them to stall. This leads
to loud banging noises (similar to a shotgun going off) and a very noticeable
vibration from the engine. Shut down immediately if this happens during startup.
During flight there are some gauges you should monitor to make sure you do not
exceed the engine limits.
Below 6000 feet AMSL the engine will normally be limited by the amount of
torque.
Between 6000 feet and 10000 feet, the Turbine Outlet Temperature (TOT) will
become the limiting factor.
Above 10000 feet, engine speed (N1) normally becomes the limiting factor.
If any of these limits are reached you must stop raising the collective.
The Air Inlet
The air inlet is designed to stabilize the air before it enters the compressor. A
particle separator may be fitted at this point to remove any foreign matter such as
dust or sand which could cause erosion of the compressor blades.
The Compressor
The compressor is made up of a series of blades and an impellor (or centrifugal)
compressor. It is designed to take large quantities of air and compress it before
directing it to the combustion section.
The engine in the JetRanger has a six stage axial flow compressor and a single
impellor compressor.
In the axial compressor, each stage is separated by stator vanes to make sure the
air hits the following blades at the correct angle. As the air flows past these
blades it becomes compressed more and more. The pressure rises and its velocity
decreases.
The centrifugal compressor diverts the air outwards into channels that lead to the
combustion section while compressing the air further.
During startup there is a bleed valve at the fifth stage that opens and allows some
of the air to escape from the compressor. This makes it easier to get the engine
up to speed and takes less power from the battery. When the engine reaches 87%
N1, the valve closes automatically. During your preflight check you should be
able to see that this valve is in the open position.
Because of the high temperatures of the compressed air in the compressor
section, this air is used to heat the cabin and for anti-icing. Anti-icing uses air
from the rear of the compressor and directs it through the compressor casing and
the inlet guide vanes to prevent ice forming there. When anti-icing is used there
will be a small rise in TOT.
The Combustion Section
In this section the fuel is mixed with the air and ignited. The air ducts are shaped
in such a way that the flame never comes into contact with the metal casing but
instead is contained within a shroud of cooler air. A large proportion of the air is
used for cooling.
Once the fuel is lit and the engine is up to running speed, the combustion is self
sustaining.
The engine is turned initially by a starter. When there is enough airflow through
the engine to keep everything cool, the fuel is ignited by the igniter plug. The
fuel enters the combustion chamber through a fuel nozzle that atomizes the fuel.
On ignition the gases expand and flow to the turbine section at an increased
velocity. Approximately 60 to 80% of the air entering the combustion chamber is
used to keep the liners cool.
The fuel nozzle is highly polished and engineers have to handle them very
carefully as the tiniest scratch will disturb the spray pattern and cause hot spots
which will eventually damage the turbine blades. It has two sets of holes for
delivering fuel. There is a small, primary hole in the center that provides fuel for
starting and secondary, larger holes for delivering fuel when the fuel pressure
reaches >150psi.
The Turbine Section
This section creates the power. The turbine inlet is the hottest part of the engine
and it is too hot for temperature sensors to survive here. Temperature is therefore
measured between the turbines by four thermocouples (the readings are averaged
and displayed in the cockpit instrument display) and is called the Turbine Outlet
Temperature (TOT). The heat is kept to a manageable level by cold air extracted
from the compressor which is driven through a connection by the turbine.
The gases pass in a forward direction through the compressor turbine blades
(N1) thus ensuring that the compressor is continually powered. From there the
gases pass through a two stage “Free Turbine” (N2). As the free turbine is not
directly connected to the compressor, the engine is easier to turn during startup.
The free turbine is connected to the Accessory gearbox which reduces the high
speed of the turbine to a more manageable level.
When more power is required, the compressor speed (N1) increases to supply
more air. At the same time more fuel enters the combustion chamber and
therefore N2 is maintained at a constant speed.
The turbine blades are operating in a very hostile environment. As the
temperatures are so high and the blades are spinning so very fast, centrifugal
force causes the blades to stretch (blade creep). This is normal however if the
engine has had a hot start, the blade creep becomes much larger than normal and
becomes permanent. The blades can make contact with the sides and expensive
repairs will be required. Higher TOT temperatures are permitted during startup
as the turbine is spinning relatively slowly.
The N1 turbine is doing more work than the N2 turbine. It is also exposed to
hotter gases. For these reasons the N1 turbine has only half of the service life of
the N2 turbine.
The Accessory Gearbox
The accessory gearbox converts the high speed of the free turbine (N2) to a more
manageable level. It has a forward driveshaft (short shaft) powering the main
rotor gearbox, a rear driveshaft powering the tail rotor, a freewheel unit and
attachment points for all the accessories such as fuel pump, tachometers,
generator etc.
The Allison 250-C20J Engine
The Rolls Royce 250/C2OJ engine is a reverse flow internal combustion
turboshaft engine featuring a free turbine. The compressor is composed of a six
stage axial flow and one stage centrifugal flow compressor and is directly
coupled to a two-stage gas producer turbine (N1). The power turbine (N2) is a
two-stage free turbine which is gas coupled to the gas producer (Compressor)
turbine (N1). When the N1 is 100% the RPM is 50,940. There is an integral
reduction gearbox which has front and rear drive splines to mate with the aircraft
drives. The engine has a single combustion chamber with single ignition. The
output shaft centerline is located below the centerline of the engine rotor and the
dual exhaust outlets are directed upward.
The engine lubrication system is a circulating dry sump type with an external
reservoir and oil cooler. Pressure and scavenge pumps are mounted within the
engine and are driven by the accessory gearbox of the engine. All engine oil
system lines and connections are internal with the exception of the pressure and
scavenge lines to the front and rear bearings. Screened openings are provided on
both sides of the engine for cooling. Below the engine is a titanium floor which
acts as a drip pan and also gives insulation from heat. This pan is curved to
provide sufficient clearance to allow the removal of accessories from the bottom
of the engine without removing the engine.
The Model 250/C2OJ is a thermodynamically rated 420 shaft horsepower (SHP)
engine. This means that Rolls Royce demonstrated the ability of this engine to
produce this power for a short duration in a test cell. For installation in a Bell
206 the maximum fuel flow the engine will receive has been limited. This fuel
flow will produce approximately 317 SHP at 100% torque at sea level on a
standard day. This flat rating is set lower than the thermodynamic limit to protect
the drive train from excessive torque. Torque gauge markings create a de-rated
limit. At takeoff the pilot may use 317 SHP or 100% torque for takeoff, (5
minute limit), and 270 SHP or 85% torque for continuous operation. This
combination of flat rating and derated will enable the aircraft to have consistent
lifting performance from sea level to moderate altitudes (see performance charts
for detailed information).
Air is supplied to the engine through air scoops on either side of the fuselage
located immediately aft of the transmission cowling. The air is directed through
a protective screen, the air inlet, and then into the compressor.
The compressor compacts and increases the velocity of this incoming air, then
discharges it into two external ducts that direct the air to the combustion section
of the engine. An automatic bleed air system is incorporated to allow a portion
of the compressed air to vent overboard during the start cycle. This will allow
the engine speed to accelerate at a desired rate and minimize the opportunity for
compressor stall to occur. As the compressor speed increases, the pneumatically
controlled bleed valve gradually closes.
The combustion section burns the fuel at peak efficiency. After the igniter is
energized and when fuel is introduced into the chamber at a precisely controlled
rate, the mixture ignites. Combustion is even and continuous as long as the
proper air/fuel mixture is maintained.
The turbine section consists of a two-stage gas producer turbine, gas coupled to
the two-stage power turbine. The turbine section is designed to utilize the energy
from combustion to drive gear trains. These gear trains drive components that
will sustain the engine operation and provide drive to the helicopter drive train.
The engine is horizontally mounted aft of the transmission and above the
fuselage to simplify the drive system, improve the inlet/exhaust arrangement,
and to reduce cabin noise. The engine is supported by bipod mounts on the
service deck to provide structural integrity. The transmission is coupled to the
engine by means of the main drive shaft and freewheeling unit.
The major engine components are a compressor, combustion section, turbine
section, and power and accessory gearbox.
Compressor
The compressor assembly consists of a compressor front support, shroud
housing, diffuser, rear support assembly, six stage axial and one stage centrifugal
impeller, scroll assembly, mount assembly and bearings. The front support
encloses the front bearing and supports it through seven hollow inlet struts. The
struts are hollow to provide anti-icing using compressor discharge air. Oil
passages are incorporated to provide lubrication of the
compressor front bearing. The compressor rear bearing is mounted in the rear
support assembly and is lubricated from the gearbox.
The compressor impeller is machined from a single piece of forged titanium. The
vanes transition from axial to centrifugal, eliminating the need for stators. At
100% N1 the compressor is rotating at 50,940 RPM. The compression ratio is
approximately 8.5:1 (8.5 Bars or 123 PSI). This rapid compression will increase
the temperature of the compressed air to approximately 290°C (555°F)
The compressor impeller requires a considerable amount of shaft horsepower to
pump air to provide the required airflow and pressure. The shaft horsepower
required by the compressor rotor varies directly with air density and N1 rpm;
and the gas producer turbine rotor must develop the horsepower required. As a
general rule a compressor will consume approximately two horsepower for each
horsepower at the output shaft.
Compressor Bleed System
This is an entirely automatic system which bleeds air from the fifth stage of the
compressor during start-up, acceleration and at low compressor pressure ratio
operations. It consists of a bleed-air control valve attached to the compressor
case which bleeds air off the fifth stage compressor wheel at low pressure ratios
to prevent compressor stall and surge.
The compressor blades are aerofoils. These aerofoils can stall if the angle of
attack becomes too great or the air velocity becomes too slow. This would result
in a loss of efficiency. Fuel consumption is reduced and rapid acceleration is
ensured by operating close to the stall angle of these compressor blades.
At low RPM the compressor does not have the same ability to force the air as it
does at high RPM. To keep the angle of attack and the air velocity within the
desired limits, it is necessary to reduce the load on the compressor during startup and low power operations.
Two features are used to control pressures in the compressor. The compressor
liner has a narrow slot machined into the perimeter near the front. This slot is
known as the inducer bleed port. During low RPM operation, air exits through
the port, reducing drag. This reduction of drag allows rapid acceleration. At
higher RPM when maximum engine efficiency is required, air enters the
compressor through this slot, increasing the volume of air entering the engine.
This inducer bleed port works in conjunction with a pressure activated, diffuser
scroll mounted bleed valve, to complete the system design.
The bleed valve is open during starting and idle operation and remains open until
a predetermined pressure ratio is obtained. At this pressure ratio, the valve
begins to modulate from open to the closed position. It reaches the full closed
position during acceleration to full power. When power is reduced the bleed
valve will open. Opening and closing of the bleed valve is controlled by
compressor discharge pressure in the diffuser.
Compressor Stall
Compressor stall can occur on any turbine engine if the conditions present
themselves.
In order to meet the design requirements, the engine must have a relatively high
power output, good fuel consumption and fast acceleration characteristics. For
these reasons it is beneficial to operate as closely as possible to the stall angle of
the compressor blades. Operating close to the stall angle has the following
benefits:
The volume of air passing through the engine is increased.
The pressure ratio of the engine is increased thus increasing the power output.
Turbine temperatures can be increased because of the greater airflow.
The efficiency of the compressor and turbine sections are increased.
To reduce the risk of the compressor stalling during startup or acceleration, the
fuel flow is carefully regulated.
So what is compressor stall?
Many pages could be devoted to explaining this but the following explanation
should help answer the question.
Compressor blades and vanes are aerofoils. The airflow over an aerofoil will
separate and become turbulent if either of the following occurs:
The velocity of the air passing over the aerofoil is too low.
The angle of attack is too high.
If the airflow over an aerofoil separates then the aerofoil stalls.
Approximately 85% of the air entering the engine is used for cooling. This
means that much more air has to enter the engine than what is needed for
cooling. The cooling air is used to control the length of the flame in the
combustion chamber and prevent it from touching the sides of the container. The
hot combustion gases are cooled by the cooling air and the cooling air is also
heated by the combustion gases. This keeps the gases at an acceptable
temperature as they mix and enter the turbine section.
If too much fuel is supplied to the burner, there will be more than enough air to
allow proper combustion. However as extra air is used during this combustion,
there will be less air available for cooling and therefore the temperature inside
the combustion chamber will rise. As the temperature rises, there will be more
gases to be exhausted. It is possible that the volume of gases to be exhausted
may exceed the capacity of the turbine and the turbine will “choke”.
When this happens, the pressure inside the combustion chamber will rise rapidly
and may equal or exceed the pressure that the compressor is producing.
If the pressure in the combustion chamber is equal to the pressure of the
compressor discharge air, then the compressor will stall.
If the pressure in the combustion chamber exceeds the pressure of the
compressor discharge air, then not only will the compressor stall but also the hot
gases will flow from the combustion chamber into the compressor section.
Both of these conditions will result in a loss of air into the combustion chamber.
The flame will not have enough oxygen and will die, resulting in a rapid drop in
temperature. As the temperature drops, the expansion is stopped (or greatly
reduced). The turbine is no longer choked and the combustion chamber pressure
drops to a very low value.
The low pressure in the combustion chamber means that air can flow in the
proper direction again. The compressor is no longer stalled and a “Surge” of air
flows back into the combustion chamber. This extremely fast movement of air
elongates the flame downstream and through the turbine causing another rapid
expansion of the gases. The cycle repeats itself at approximately 120 times per
second.
Compressor stalls may or may not have an audible sound but there will often be
a vibration. If the stall is severe a flame may emanate from the exhaust or a very
loud backfire may be heard. Smoke may also be seen.
If the proper corrective action is made immediately then it is unlikely that any
damage will occur. The actions to be taken are:
Reduce the throttle to flight idle.
If the stall remains then close the throttle completely and shut down the engine.
Rapid throttle movements may induce stalling therefore it is prudent to make
smooth, slow throttle movements.
Anti-Icing System
Operation of the engine during icing conditions could result in ice formations on
the compressor front support. If ice were allowed to build up, airflow to the
engine could be affected and performance decreased. The engine anti-icing
system is designed to prevent ice formation on the compressor front support. The
pilot activates the system by placing the engine anti-ice switch to the ENGINE
ANTI ICE (on) position.
When the system is in operation, compressor discharge air, which has been
heated due to compression, flows through the anti-icing valve and tube to the
compressor front support passages. Hot air flows between the double wall outer
shell and into the seven hollow radial struts. The hot air flowing through the
radial struts exhausts either from small slots in the trailing edge of the struts or
from the double wall bullet nose hub of the compressor front support. The
compressor inlet guide vanes and front bearing support hub are the only engine
components with anti-icing provisions.
Using the engine anti-ice will reduce the power available from the engine. Check
the flight manual for power available when using anti-ice.
Combustion Section
The combustion section consists of an outer combustion chamber and a
combustion liner. The combustion liner is located inside the outer combustion
chamber.
The fuel nozzle and igniter plug are mounted in the aft end of the outer
combustion case. Compressor discharge air (P3 air) enters the combustion case
via two air tubes connected to the diffuser scroll. Air then enters the combustion
liner through vanes and holes in the liner dome and skin. The air is mixed with
fuel sprayed from the fuel injector and combustion takes place. Air enters the
combustion liner in such a manner that the flame pattern is prevented from
touching the wall of the combustion liner, thus insulating and protecting the liner
from damage and deformation caused by excessive heat. Combustion is even and
continuous as long as the proper air/fuel mixture is maintained.
Approximately 20% to 25% of the air delivered to this section is required to burn
the fuel. The other 75% to 80% is used to expand the gases and cool the internal
parts. Approximately 2% of the air is used to seal oil passages.
Turbine Section
As the gas stream leaves the combustion chamber, it passes to the turbine
section. This high-energy gas stream powers the two turbine sections to sustain
the airflow through the engine and provide output power. The turbine section
consists of a gas producer turbine support, power turbine support, a two-stage
gas producer turbine rotor, a two-stage power turbine rotor, and an exhaust
collector support.
The gas producer turbine drives the compressor and certain engine accessories
through the N1 drive train. At 100% N1 the turbine rotates at approximately
50,940 rpm.
The power turbine drives the power output shaft, and certain engine accessories.
At 100% N2 rpm, the power turbine rotates at a constant 33,290 rpm.
The turbines take advantage of the impact and reaction of the gases passing
through the gas producer and power turbines. The turbines may be broadly
classified as an impact, reaction type. Varying the fuel flow changes the
temperature of the gases passing through the turbine section and the amount of
energy in the gas stream. This variation of gas energy will result in a change in
the expansion rate of the gases as well as a change in gas velocity through the
turbines. Consequently, any increase in gas temperature will result in an increase
in the torque developed by the turbines.
Power and Accessory Gearbox
The main power and accessory drive gear trains are enclosed in a single gear
case (Accessory gearbox). The gear case serves as one of the structural supports
for the engine. All engine components are attached to the case. There are two
independent drive trains in the gearbox, gas producer (N1), and power turbine
(N2).
The gas producer (N1) gear train drives: N1 tachometer generator, fuel pump,
fuel control, starter generator, and oil pump.
The power turbine (N2) gear train drives: N2 tachometer generator, power
turbine governor, torque meter, and freewheeling unit.
A two-stage helical and spur gear set is used to reduce N2 rotational speed from
33,290 rpm at the power turbine to 6016 rpm at the output drive spline. The
output drive shaft is connected to the main rotor gearbox.
The accessory gearbox also incorporates a torque meter to measure the engine
output torque. The power turbine tachometer generator and power turbine
governor are driven by the power turbine gear train. The gas producer gear train
provides drive for the oil pumps, fuel pump, gas producer fuel control, gas
producer tachometer generator and starter generator.
Torque Meter
The torque meter in the 250 series engine gearbox is a hydraulic type that uses
the engine lubrication system as its oil (hydraulic) pressure source. In order to
minimize friction effects and provide accurate measurement of torque, the axial
gear thrust on the helical torque meter gear shaft is high. System pressure must
always be greater than the torque-meter oil pressure. Therefore, it is necessary to
regulate the system oil pressure to the relatively high value of 115-130 psi. If the
torque meter fluctuates erratically then there is quite likely an oil pressure
problem and an engineer should be informed.
The power turbine gear train has two stages of helical gearing. The helical power
train drive gear driving the larger diameter gear on the helical torque meter gear
shaft accomplishes the first stage reduction. The smaller diameter gear on the
helical torque meter gear shaft driving the helical power takeoff gear-shaft
accomplishes the second stage reduction. Helix angles are such that both stages
of reduction produce a forward axial thrust on the helical torque meter gear
shaft. If friction is neglected, this axial thrust is directly proportional to the
torque transmitted through the gears.
A ball bearing transmits helical gear thrust from the rotating helical torque meter
gear shaft to the non-rotating torque meter piston. This piston has an antirotation pin which contacts a second anti-rotation pin on the flange of the
support shaft. The piston is free to slide axially on the non-rotating and axially
retained support shaft. This sliding piston serves as a variable position valve
which admits regulated engine pressure oil from the support shaft to the oil
chamber formed by the piston and the flange on the support shaft. The flange
also has an external groove which houses a piston ring and expander.
The support shaft and piston are arranged in such a manner that an increase in
gear thrust increases the piston inlet port opening. Since the outlet openings of
the piston remain constant, the pressure inside the oil chamber is directly
proportional to torque. Pressure in the oil chamber is directed to the torque meter
pressure sensing port on the front side of the accessory gearbox. The chamber oil
is transferred through the support shaft anti-rotation pin and filter screen that is
located in this passage.
During a stabilized power condition, the axial thrust force acting on the piston is
counter balanced by the torque meter pressure in the oil chamber. Let us assume
that the torque output of the engine is increased. This results in increased axial
thrust acting on the piston and an unbalanced condition on the piston. The piston
moves forward slightly to increase the piston inlet port opening. With a larger
opening, there is less restriction to the flow of oil into the oil chamber. As a
result, the pressure in the oil chamber increases. The pressure in the oil chamber
continues to increase until its force acting on the piston counterbalances the axial
thrust acting on the piston. When the two forces are equal, the piston stops
moving and the pressure in the oil chamber will be higher than it was prior to the
torque increase. The torque meter indicating system, sensing a higher pressure,
registers the increased torque.
Engine Accessories
The gas producer tachometer/generator generates an electrical signal that is a
function of N1 gas producer turbine rotor rpm. The output of this tachometer
generator is delivered to the gas producer tachometer and displays gas producer
turbine (N1) RPM.
The power turbine tachometer/generator generates an electrical signal that is a
function of N2 power turbine rotor rpm. The output of this tachometer generator
is delivered to the power turbine tachometer indicator and displays power
turbine (N2) RPM.
The ignition exciter converts 28V DC energy into high temperature/high
amperage arcs at the spark igniter gap and is required during the starting cycle.
The igniter is threaded into the combustion outer case. It extends into the
combustion liner providing ignition sparks that ignite the fuel/air mixture during
start. Once the start is complete, the combustion is continuous and the ignition
source is no longer required. Electrical power is supplied from the electrical bus
and is protected by a 5 amp C/B marked IGNTR.
The oil pump supplies a pressurized volume of oil for proper lubrication and
cooling of the bearings and gears. A scavenge pump collects oil from the various
sump cavities and returns it to the supply tank.
The fuel nozzle atomizes and injects fuel into the combustion liner at the proper
spray angle and pattern.
The burner drain valves drain any unburned fuel from the combustion section
following an engine shutdown. During start, the drain valves close when the air
pressure within the combustion section exceeds the air pressure on the outside of
the combustion section by a predetermined value. The valves open on shutdown
by means of spring action.
The starter/generator is used as a DC motor to crank the engine during the
starting cycle. Once the start is completed the DC generator supplies all the
electrical needs of the helicopter and charges the battery.
The engine-driven fuel pump assembly is a single element gear type pump that
produces approximately 600 psi. This high pressure is needed to operate the fuel
control system. The pump receives filtered fuel from the airframe mounted fuel
filter, increases the pressure, and delivers that high pressure fuel into the fuel
control unit.
The gas producer fuel control and power turbine governor serve as the fuel
control system to provide speed governing of the power turbine rotor and over
speed protection for the gas producer turbine rotor. The fuel control system
senses N1 and N2 rpm, compressor discharge pressure, fuel control lever
position, and collective position to regulate and maintain fuel flow between
established limiting values. The system regulates the engine functions during
starting, acceleration, governing, deceleration, and shutdown.
Oil System
The engine incorporates a dry sump oil system with an externally mounted
supply tank and an oil cooler located on the top aft section of the fuselage. Oil is
supplied from the tank to the pressure and scavenge pumps mounted within the
engine accessory drive gearbox. The oil filter assembly, consisting of an oil
filter, filter bypass valve, and pressure regulating valve, is located externally in
the upper left-hand side of the gearbox. Magnetic chip detectors are installed at
the bottom of the gearbox and at the engine oil outlet connection. All engine oil
system lines and connections are internal except the pressure and scavenge lines
to the front compressor bearing and to the bearings in the gas producer and
power turbine supports.
The system is designed to provide adequate lubrication, scavenging, and cooling
as needed to the bearings, splines, and gears, regardless of the helicopter attitude
or altitude. Jet lubrication is provided to all compressor, gas producer turbine,
and power turbine rotor bearings, and to the bearings and gear meshes of the
power turbine gear train, with the exception of the power output shaft bearings.
The power output shaft bearings and all other gears and bearings are lubricated
by oil mist.
Oil from the tank is delivered to the pressure pump. System oil pressure is
regulated by the pressure regulating valve. Pressurized oil passes through the
filter and then to various points of lubrication. The scavenge pump returns the
heated oil to the engine oil outlet port where it is routed to the oil cooler.
The oil cooler and blower assembly are mounted aft of the engine. The cooler
blower is a part of the tail rotor drive system and forces air upward through the
cooler core. Return oil from the engine flows through the oil cooler bypass
valve. The valve bypasses the cooler if the oil temperature is low or routes the
flow through the cooler when the oil temperature is high. The valve is designed
to regulate oil temperature between 71 and 81 degrees centigrade.
The normal capacity of the engine oil tank is 5.5 U.S. quarts. The oil level is
checked by a dipstick mounted on the cap and adapter assembly. The oil tank
provides port openings for the supply tube, vent tube, scavenge tube,
temperature bulb, and drain valve.
Oil pressure is monitored by routing regulated oil pressure through an external
line to the engine oil pressure gauge in the cockpit. Oil temperature is monitored
by the use of a temperature bulb mounted in the oil line at the base of the oil
tank. That signal is delivered to the temperature side of the engine oil gauge
through the aircraft 28 volt electrical system and is protected by a 5 amp circuit
breaker labeled ENG XMSN TEMP IND.
If the oil filter becomes blocked, a bypass valve opens to ensure oil still flows. If
the bypass valve is opened a red pop-out button is visible and an engineer should
be informed. The bypass valve can be inspected on the port side cowling.
An erratic torque meter will normally be an indication of an impending engine
oil pressure problem and an engineer should be informed.
If the engine oil pressure drops to zero, you have only 30 seconds to get the
engine shut down before severe damage is done to the engine. The two minute
shutdown period is not required if this happens as there is no oil to carry heat
away. Shut the engine down as quickly as possible (either in flight or on the
ground).
Only certain oils conforming to following specifications are approved for use in
the engine:
Specification OAT Range
MlL-L-7808 (NATO 0-148) for any OAT
MIL-L-23699 (NATO 0-156) OAT above -40°C (-40°F)
DOD-L-85734 OAT above -40°C (-40°F)
NOTE
Because of availability, reduced coking and better lubricating qualities at higher
temperatures, qualified MIL-L-23699 oils are preferred by the engine
manufacturer.
NOTE
Long-term use of DOD-L-85734 oil may increase the probability of seal leakage
in the accessory gearbox.
Refer to the flight manual for engine oil limitations.
Engine oil Capacity: 5.5 U.S. quarts (5.2 liters).
The engine oil tank is located under the aft fairing and access doors are provided
for filling and draining the oil tank. A dip stick attached to the filler cap is
provided to determine the quantity of oil in tank.
It is beneficial to try and keep this tank full.
It is possible for the labyrinth seal to fail between the freewheel clutch and the
accessory gearbox. If this happens, high pressure oil from the main gear box will
flow into the engine. If you notice the engine oil reservoir fuller than expected or
overflowing when the dip stick is removed; this is a good indication that the seal
has failed and an engineer should be informed. The oil level in the tank should
not change significantly from flight to flight.
Do not use your finger to “feel” for the oil level. There are many hazardous
chemicals in this oil and they can be easily absorbed through the pores in your
skin. If your helicopter is on high skids you would be well advised to invest in a
small step to enable you to reach and see the oil level.
Engine Fuel System
The principal components of the engine fuel system are the engine driven fuel
pump, gas producer fuel control, power turbine governor, and a fuel nozzle.
The engine driven fuel pump assembly incorporates a single gear type pumping
element and a bypass pressure regulating valve. The pump produces highpressure fuel (600 psi) based upon N1 speed and sends it to the fuel control unit
for metering. The aft face of the pump provides a mounting pad for the gas
producer fuel control.
The gas producer fuel control and the power turbine governor provide for a fuel
metering system. This system senses gas producer (N1) rpm, power turbine (N2)
rpm, compressor discharge pressure (P3), and throttle position to regulate and
maintain fuel flow.
The gas producer fuel control unit is the device that will meter fuel into the
engine to determine idle speed, fuel flow for the initial combustion process, and
start acceleration.
When the throttle is in the full open position, the fuel control unit will control the
fuel flow into the engine based upon the pneumatic signal (Bendix), or hydromechanical signal (Ceco) on older models received by the power turbine
governor.
An accumulator is located in the pneumatic line between the power turbine
governor and the gas producer fuel control. The accumulator dampens pressure
signals from the power turbine governor to the fuel control due to torsional
vibrations encountered from the rotor system and maintains a constant N1 rpm.
The fuel nozzle has a single entry and a dual outlet orifice. This nozzle provides
a finely atomized spray of fuel at all flow conditions that are required by the
engine. It is designed to provide an optimum spray angle for starting the engine
and an even distribution of fuel into the combustion liner. The nozzle is equipped
with a filter to minimize the possibility of contamination.
The fuel control system maintains engine power output by regulating the gas
producer speed. Gas producer speed levels are established by the action of the
power turbine fuel governor, which senses power turbine speed. Power turbine
speed (N2) is selected by the pilot utilizing the “beep” switch on the collective.
The power required to sustain the selected rpm is maintained by the power
turbine governor action on metered fuel flow. The power turbine governor sends
pneumatic signals or hydro mechanical signals (depending in the type of fuel
control unit installed) to the fuel control to increase or decrease fuel flow to the
combustion chamber based on N2 rpm changes and movement of the collective.
The engine controls consist of both the N1 gas producer controls and the N2
droop compensator controls. The gas producer controls are operated by a twist
grip on the collective stick, and the droop compensator controls are operated
from a bell crank in the collective system.
The twist grip throttle on the collective controls the gas producer fuel control
and consists of a flexible control cable which extends from the throttle arm on
the rear of the collective stick to a bell crank assembly mounted on the engine
deck. A control tube is connected between the bell crank and a lever mounted on
the fuel control shaft.
N2 Droop Compensator System
The droop compensator control system consists of a mechanical linkage between
an idler in the collective system and a lever mounted on the power turbine
governor shaft. Movement of the collective stick results in a repositioning of the
governor shaft. This action provides droop compensation to prevent rpm
variations as power changes are made. The system incorporates a linear actuator
that is controlled electrically by a GOVernor RPM INCrease/DECRease switch
mounted on the collective stick.
The droop compensator maintains N2 engine rpm as power demand is increased.
It is a direct mechanical linkage between the collective stick and the speed
selector lever on the N2 governor and will maintain N2 rpm when properly
rigged.
Droop is defined as the speed change in N2 engine rpm as power is increased
from a no load condition. It is a characteristic designed into the governor system
to prevent instability from developing as engine output is increased. Without this
characteristic, N1 speed would overshoot or hunt the value necessary to satisfy
the new power condition. If N2 is allowed to droop other than momentarily, the
reduction in rotor speed could become critical.
Shutting down
The engine requires a two-minute shut down period. It is a misconception that
this time is for only cooling the engine down. While the engine is at idle, oil is
being purged from the injectors in the turbine section. Without the two minute
shut down period these injectors can become clogged up with oil and lead to a
premature failure of the turbine bearings.
Some pilots have a habit of cranking the engine over after shutdown to force the
temperature down. This is acceptable if another start is required and the TOT is
above 150°C. It is not acceptable to do this at any other time. Forcing cold air
through the still hot engine causes thermal shock and will lead to cracking of the
turbine blades.
8
DRIVE TRAIN AND ROTORS
Drive Train and Rotors (General)
The drivetrain system provides a means of transmitting power from the engine to
the main and tail rotor assemblies. The drivetrain includes the freewheeling unit
assembly, main drive shaft, transmission, mast, tail rotor drive shaft, and the tail
rotor gearbox. The rotor systems include the swashplate assembly the main rotor
hub and blades and the tail rotor hub and blades, and the tail rotor pitch change
mechanism.
Freewheeling Unit Assembly
The freewheeling unit is mounted on the accessory gearbox and driven under
power from the power takeoff gear shaft. Engine power is transmitted to the
outer shaft of the freewheeling unit, through the engaged sprag clutch and then
to the freewheeling inner shaft. The freewheeling inner shaft is connected to the
main driveshaft on the forward adapter flange. The tail rotor drive system is
driven through a flexible coupling and a splined adapter mounted on the aft end
of the freewheeling inner shaft.
During autorotation, the sprag clutch disengages and the rotational forces of the
main rotor are utilized to drive the transmission accessories and tail rotor drive
system.
Main Driveshaft
The main driveshaft is designed to transfer power from the engine to the
transmission during normal operations and to transfer from the transmission to
the tail rotor system during autorotation. The main driveshaft rotates at 6016 rpm
at 100% N2.
The main driveshaft is located forward of the freewheeling assembly and large
“couplers” are visible at each end. These couplers have “Temp-plate” strips
stuck to the surface (two red and two yellow on each coupler). The temp-plates
give an early indication of overheating of the couplers and during the walkround; the strips should be inspected for any darkening of the white dots on their
surface.
During your walk-round, use your fingers to feel where the shaft enters the
forward coupler. There is a rubber boot here that seals grease inside the coupler.
If this boot should deteriorate or become damaged, grease will start to leak out
and you will feel it with your fingers. Another indication of a failing coupler is
visible streaks of grease on the floor below the coupler. If the coupler has had a
very serious over-temp then it will be discolored and the metal will have a
yellow colour.
Transmission
The main transmission is mounted forward of the engine and is coupled to the
engine by means of a short driveshaft. The transmission is a reduction gearbox
which transmits engine power at reduced RPM to the main rotor by means of a
spiral bevel gear and a planetary gear stage.
The transmission reduces engine rpm, changes the angle of drive from the
engine to the rotor, and serves as the structural member that links the mast to the
airframe. It provides a two-stage reduction of 15.23 to 1.0 (6016 to 395 rpm).
The transmission incorporates a self-contained lubrication system with an oil
cooler mounted on the transmission. Accessory mounting pads and drives are
provided on the transmission for the hydraulic pump and rotor tachometer
During autorotation the main rotor drives the power input shaft. Under this
condition, the freewheeling unit provides a disconnect from the engine and
allows autorotational forces of the main rotor to drive the transmission, tail rotor
and all transmission accessories.
The transmission housing is secured by an isolation mount which minimizes
airframe vibration and permits higher flight speeds without compromising the
fatigue life of the airframe.
The transmission is supported by a pylon support link on each side. This frame
allows limited sideways movement and zero fore and aft movement. A
cylindrical drag pin extends down from the forward end of a drag link and fits
loosely in a hole in the pylon stop mounted on the airframe, providing a positive
limit of travel of the pylon. A heavy landing will cause this pin to contact the
airframe causing an indentation or scratch on the pin. Unless you want to get
blamed for the heavy landing, you should point this out to the engineer before
flight.
During preflight inspections, the main rotor gearbox should be inspected for
correct oil level in the sight glass. Look for oil leaks. Inspect the gearbox
mountings for security. Inspect the drag pin (this should be centered in the
opening on the floor). Inspect the isolation mount for gouging. Check the
witness marks on all visible nuts and bolts. Check that the button on the oil
bypass valve has not popped out.
Transmission Oil system
Lubrication is provided by means of an oil pump, a pressure regulating valve, an
oil cooler, a filter element, and two spray jets. The oil pump is a constant volume
type driven by internal oil pump, which delivers pressurized oil externally to the
filter and housing assembly, cooler, and then returns it to the main transmission
where spray jets lubricate the internal parts. A sight glass is located on the right
side of the main transmission lower case where the oil level can be easily
checked.
A non-vented filter cap is located on the transmission top case.
The transmission oil system also provides lubrication for the freewheeling unit
mounted in the engine accessory gear case. Oil pressure for the lubrication of the
freewheeling assembly and to the transmission oil pressure gauge on the
instrument panel is taken from a tee fitting installed in the pressure line.
Oil temperature indications are provided by an oil temperature bulb located in
the outlet side of the oil filter housing and a high oil temperature switch that is
connected to the TRANS OIL TEMP caution segment in the caution warning
panel (CWP). The TRANS OIL PRESS caution segment light is connected to the
transmission oil pressure switch located in the oil pressure line below and
forward of the instrument panel.
The pump scavenges oil from the lower case sump through a wire screen and the
lower chip detector. Oil is then directed to the transmission oil filter assembly.
The transmission oil pressure regulating valve, which is located on the left rear
corner of the lower case, is used to adjust the oil pressure to normal operating
limits and to relieve excess oil pressure back into the transmission case.
The transmission oil filter and housing assembly, which is mounted on the port
side of the transmission, has a filter element, bypass indicator, bypass valve,
thermostat, temperature bulb, and a temperature switch.
The transmission oil temperature switch is located on the oil filter housing and is
electrically connected to the TRANS OIL TEMP caution light segment. When
the oil temperature reaches 110°C, the temperature switch will illuminate the
caution light segment.
The oil pump inlet screen assembly, which is located in a boss adjacent to the
transmission oil pump, filters oil prior to its entering the oil pump.
Two oil jets are incorporated in the transmission. The No. 2 oil jet directs a
lubricating oil spray to the transmission bevel gears, and the No. 1 oil jet
lubricates the planetary pinions and mast bearing.
The four electric chip detectors (three on the transmission and one on the
freewheeling unit) consist of a self-locking, bayonet type probe (TRANS CHIP)
with a permanent ceramic magnet at the end. In the event that metal particles
should become free in the oil, the magnet will attract them. When sufficient
metal is attracted to complete a circuit between the pole and earth, the applicable
segment on the caution panel will illuminate.
The oil cooler contains a single core and is mounted to the top aft side of the
transmission. The thermal bypass valve controlling the flow of oil through the
oil cooler core is located on the outlet side of the oil cooler and allows oil to
bypass the oil cooler when oil temperature is low.
Transmission Mounts
The transmission, mast, main rotor hub, and the blade assembly are supported by
and isolated from the fuselage by means of two pylon support links and an
isolation mount.
The pylon support links permit the transmission system limited movement
laterally and prevent longitudinal movement. This can be checked by holding the
main rotor shaft above and moving it from side to side.
The isolation mount is attached to the rear of the gearbox and is also attached to
the airframe. It has an elastomeric material that dampens vibrations from the
gearbox and prevents these vibrations from being transmitted through the rest of
the airframe. This isolation mount should be checked for security and also to
make sure that the outer coupling (situated above it) has not made contact with it
in the past as this may indicate a previous hard landing.
The drag pin is situated directly underneath the transmission gearbox. This pin
extends through a hole in the deck and is there solely as an indicator for
excessive fore or aft movement of the transmission system. If the aircraft has had
a hard landing then this pin may shear off. It should be centered over the hole on
the deck.
Freewheel Unit
The freewheel unit allows the main rotor gearbox to “freewheel” in the event of
an engine failure. You can autorotate normally and keep control of the helicopter.
It is located in the forward section of the accessory gearbox.
During shutdown the Nr and N2 needles should split, indicating that the
freewheel unit is functioning normally.
If the transmission chip light comes on during flight you should land as soon as
possible. This light means that there are metallic particles in the transmission oil
which also lubricates the freewheel unit. It could indicate an impending failure
of the freewheel unit.
If the sprag clutch inside the freewheel unit fails, it can have a similar effect as
an input drive shaft (the shaft coming from the accessory gearbox to the main
rotor gearbox) failure.
If the input driveshaft fails there will be a loud bang to warn you and the tail
rotor driveshaft will be powered by the engine. If the sprag clutch fails there will
be no loud bang and the tail rotor driveshaft will be powered by the
transmission. This will cause both the main rotor and tail rotor RPM to decay
and you have to enter autorotation.
In the event of an input driveshaft failure, the tail rotor RPM will increase
rapidly to 120% of its design speed causing a very sudden yaw to the left and
subsequent roll to the left. The N2 will overspeed significantly. You should
immediately enter autorotation and close the throttle to the idle detent. This will
bring the N2 back to 100% and subsequently the tail rotor RPM back to 100% of
its design speed also. The tail rotor should therefore perform as normal during
the landing phase of the autorotation.
Main Rotor System
The main rotor system is a semi rigid, underslung hub, feathering axis rotor with
two metal blades. The hub assembly is attached to the mast by means of a
splined trunnion that also functions as the rotor flapping axis. The flapping
bearings are grease lubricated roller bearings that assemble into the pillow block
housing that is attached to the yoke and also serves to provide hub centering
provisions together with the trunnion.
The yoke serves as the pitch change axis, and the grips are attached to the yoke
by means of the tension-torsion strap assemblies which transfer centrifugal
loading from the blades to the yoke and also assist to counteract aerodynamic
forces. The pitch change bearings are housed in the grips and are grease
lubricated. The change in blade pitch angle is accomplished by rotating the grips
on the yoke with an input at the pitch horns. The blades are attached to the grips
with bolts which have hollow shanks that are used to install weight for static and
dynamic balance of the hub and blade assembly. After balancing, the bolts must
be kept with their respective rotor hub grips. Blade alignment is accomplished
by adjustment of the blade latches which engage the root end of the blade.
The blade grips are retained on the hub yoke by means of tension-torsion strap
assemblies. Changes in blade pitch angle are made by rotating the grips on the
yoke journals.
The trunnion mates the rotor system to the mast, transmits drive to the yoke
assembly, and serves as a pivot point for the rotor flapping axis.
The yoke supports the main rotor system. It has two hollow and pre-coned
journals to relieve stresses. Each journal has the hollow bore especially
machined at the outboard end to fit a sealed assembly, and the inboard end is
machined to fit the tension-torsion (T-T) strap fitting which opposes centrifugal
force.
The T-T strap, which consists of stainless steel wire wrapped around two steel
ends, provides a connection between the grips and yoke. They anchor the blades
to the rotor head (yoke). During rotor feathering, the flexibility of the strap
allows the grip bearings to rotate on the yoke journals. As this item cannot be
checked for condition, it is time limited and must be replaced at certain calendar
dates even if the helicopter never flies.
The pitch link connects the pitch horn on the blade grip to the swashplate outer
ring for control input from the cyclic and collective controls. A pitch link is
required for each main rotor blade. The pitch links should be inspected regularly
for corrosion or surface damage.
The grip assembly is the structural member between the yoke and the main rotor
blades. The outboard end of each grip has a machined hole for the blade
retention bolt, and the grip body has drilled tangs to fit the blade latch bolt that
retains the outboard end of the T-T strap.
Main Rotor Blades
The rotor blades are all metal construction consisting of an aluminum alloy spar,
spar spacer, aluminum alloy trailing edge, aluminum honeycomb filler and
aluminum skin. All structural components are joined by metal to metal bonding.
The blades are set in hub grips at a preconed angle and are secured by a single
retaining blade bolt in each grip. The blades have swept tips and are individually
interchangeable.
An inboard trim tab and an outboard trim tab are provided on the trailing edge.
The blades are manufactured to very high tolerances but they are all slightly
different. These trim tabs enable the differences to be corrected. They are not
designed to make the blade fly higher or lower by changing the camber of the
blade. Instead they produce a twisting moment to neutralize the twisting moment
of the blade. Bending these tabs down will make the blade track down as the
upward force generated will push the leading edge down.
The blades rotate at 688 feet per second (470mph/407kts) therefore at 100kts
IAS the retreating blade is experiencing airflow reversal along approximately
25% of its length.
Diameter
33’ 4”
Chord
1’ 1”
Twist
-10°
Tip Speed
688’/second
RPM
395 RPM @ 100%
The flap restraint is installed on the main rotor hub trunnion and incorporates
counterweights and springs which serve to limit flapping during starting and
shutdown but permit normal flapping at operating rpm. The flap restraint arms
operate in the range of 25 to 31% as RPM increases or decreases.
Blade Tape
The leading edge of the blades may be covered with blade erosion protection
tape (Blade Tape). This is a polyurethane tape and is recommended for
protecting the blade leading edge against sand and dust erosion. It also gives
partial protection against rain. The tape may be removed and replaced with new
tape as often as necessary. If the tape ends start to peel or appear loose, the tape
should be removed before the next flight. Blade tape should be inspected daily
for bubbles, cracks and signs of loss of adhesion.
If the tape ends peel or appear loose, remove the tape before the next flight.
Tail Rotor Driveshaft
The tail rotor drive shaft is made up of the forward short shaft, the oil cooler
blower shaft, the aft short shaft, and the tail rotor drive shaft segments. Steel
laminated flexible couplings (Thomas couplings) are used to connect the shaft
sections and the tail rotor gearbox.
The forward and aft short shafts are located on either side of the oil cooler
blower assembly. The forward short shaft is constructed of steel and is connected
to the aft end of the freewheeling assembly and the forward end of the fan shaft
by means of splined adapters. The forward shaft should have movement (fore
and aft).The aft short shaft is constructed of aluminum alloy and is connected to
the aft end of the fan shaft by means of a splined adapter and the first tail rotor
drive shaft segment.
The segmented tail rotor drive shaft consists of five segments that extend along
the top of the tail boom. Each segment of the drive system is identical and
interchangeable with the others.
The oil cooler blower assembly is mounted on the upper structure, aft of the aft
firewall and is driven by the tail rotor drive shaft. The squirrel cage type impeller
fan is mounted on a flanged shaft that is mounted in bearing hangers. The oil
cooler shaft connects to the forward and aft short tail rotor shafts and is part of
the tail rotor drive system.
The oil cooler blower provides cooling air for the engine oil system,
transmission oil system and the hydraulic system. The engine oil cooler mounts
above the blower housing while a flexible duct conveys cooling air forward to
the transmission oil cooler and the hydraulic reservoir.
Hanger Bearings
Hanger bearing assemblies support the drive shafts and flexible, steel, disc
couplings. Thomas couplings are used to both connect the shaft sections and
allow continued alignment with the tail boom.
Tail Rotor Gearbox
The tail rotor gearbox, located on the aft end of the tail boom, drives the lefthand side mounted tail rotor at 2553 RPM. It contains two spiral bevel gears
positioned at ninety degree angles to the other. The direction of drive is changed
ninety degrees and there is a speed reduction of 2.35 to 1.0 at the gearbox.
The magnesium housing is attached to the tail boom by means of four studs. The
assembly includes a breather type filler cap, oil level sight gauge, and a
combination electrical chip detector and self-closing valve.
The chip detector consists of a self-locking bayonet probe with a permanent
magnet at the end. Free ferrous metal particles in the oil T/R CHIP will be
attracted to the magnet; and when sufficient metal is attracted to complete the
circuit between pole and earth, the T/R CHIP detector segment on the caution
panel will illuminate. The valve automatically closes and prevents loss of oil
when the electric chip detector is removed for inspection. It also serves as a
drain plug.
Tail Rotor Assembly
During flight, the tail rotor counters the torque of the main rotor to permit
forward progress without the helicopter fuselage spinning in the opposite
direction. The pitch of the tail rotor blades is varied by means of the tail rotor
control system in order to make right and left pedal turns.
The tail rotor assembly is a delta hinge type that consists of a hub and two
interchangeable blades. The yoke of the hub is made with a 4° twist for each
blade. The aluminum alloy forged yoke is attached to the tail rotor gearbox shaft
by a splined trunnion inside the yoke to provide a flapping axis for the assembly.
At the time of assembly, spanwise balance is accomplished by using washers on
the blade bolts at the yoke, and chordwise balance is accomplished by using
weights and washers on the trunnion, bearing housing restraining bolts.
The tail rotor blades are all metal assemblies consisting of a stainless steel shell
reinforced by a honeycomb filler and stainless steel leading edge abrasive strip.
Ballast stations, located at the inboard trailing edge and the tip of the blades are
provided for mass balance of the blades. Weights used in these locations are
determined when the blade is manufactured.
The blades are attached to the yoke by means of spherical bearings (provide for
pitch change of the blades) that are mounted in the grip plates of the blades on
the pitch change axis. The blades should be inspected regularly for cracks, dents,
corrosion and de-bonding.
Connecting linkage consists of push-pull tubes, bellcranks, levers, and supports
that connect the pilot’s tail rotor control pedals to the tail rotor pitch change
mechanism.
Tail rotor pitch control is achieved by means of bell crank, rod, and lever
assembly mounted on the tail rotor gearbox which actuates a control tube
through the hollow rotor drive shaft to the crosshead and pitch links.
9
FLIGHT CONTROLS AND HYDRAULICS
Flight Controls
The flight controls are mechanical linkages that are actuated by conventional
controls and used to control flight attitude and direction. Both the cyclic (fore,
aft, and lateral) and the collective controls incorporate hydraulic servo actuators.
This makes the controls feel very light. The helicopter can still be flown in the
event of a hydraulics failure but the control forces are much greater.
The flight controls are routed beneath the pilot seat, aft to the center of the
helicopter and the vertical control column then up to the cabin roof. The vertical
control column also serves as a primary cabin structure. Access doors on the aft
side of the control column and removable seats are provided for inspection of the
control components and maintenance accessibility.
Cyclic and collective controls are routed to the main rotor blades through the
swashplate. The directional controls are routed through the tail boom to the tail
rotor. Fixed length control tubes and a minimum of adjustable tubes simplify the
rigging.
Dual controls are installed as an option to provide dual flight control capability
either for operations requiring a pilot and a co-pilot or for pilot training
operations. Installation of dual controls provides a collective stick, cyclic stick,
and a tail rotor control pedal assembly for the co-pilot. The pilot and co-pilot
controls are similar in appearance and relative position and the control input to
the rotor system is the same. The control feel for the co-pilot’s controls is the
same as that for the pilot’s controls. The co-pilot’s controls are connected to the
pilot’s by means of a jackshaft, control tubes, and electrical wiring. Quick
disconnects are provided for the co-pilot’s collective and cyclic.
Aluminum alloy control tubes are used throughout the collective, cyclic, and tail
rotor control system. Some control tubes are fixed in length with bonded end
fittings, while others may have adjustable fittings. Bellcranks, levers, and
supports are used throughout the collective, cyclic, and tail rotor control systems.
These parts transmit or control change movements in the particular system in
which they are installed.
Cyclic
The cyclic controls consist of a control stick, torque tube, yoke, two hydraulic
servo actuators, control tubes, and bellcranks. Movement of the cyclic is
transmitted through linkages 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
and provides a mounting point for the elevator controls. The cyclic system yoke
extends aft from the cyclic stick support and jackshaft. Movement of the cyclic
stick is transmitted to the mixing lever by means of the yoke.
This mixing lever transmits cyclic movement to the swashplate through
mechanical linkage and servo actuators.
A balance spring is used in the system to minimize the cyclic stick mass
imbalance forces and elevator induced forces in the longitudinal control system.
For helicopters with dual controls, the co-pilot’s cyclic stick is installed in front
of the co-pilot seat with all VFR control functions. A quick disconnect feature
permits rapid removal of the stick. If the electrical connector is disconnected, an
adapter must be installed to provide circuit continuity. A spring pin assembly is
provided to ensure positive engagement of the stick.
A cyclic friction control knob is fitted at the base of the cyclic on the pilot’s side.
The cyclic grip is equipped with switches for transmitting on the radio and using
the intercom system.
The co-pilot’s cyclic is removable. It is important to take care when installing
this cyclic. To do this, you must align the tabs on the cyclic with the notches on
the lever assembly at the base of the seat. Slide the collective into the hole and
then screw on the turn nut to engage the threads on the lower assembly. Tighten
the nut using finger pressure only. Insert the straight end of the “C” spring into
the hole just aft of the threads. Connect the electrical connector.
Collective
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 servo actuator has an
irreversible valve to reduce feedback and to provide for use of controls in the
event of hydraulic boost failure.
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. An adjustable friction bearing mounted
on the jackshaft allows the pilot to adjust the friction to his own requirements. A
minimum friction adjustment clamp located at the left end of the jackshaft
ensures that the collective stick will always have a preset minimum friction to
eliminate collective bounce.
The collective has a twist grip throttle and switches for the landing light and
governor RPM control. The starter button is also situated here. Just behind the
starter button is the idle release button. This button prevents inadvertent shutting
down of the engine and must be depressed to allow the throttle to be fully closed.
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 co-pilot collective stick is installed at the
left of the co-pilot’s seat with a fully functioning throttle control. A quick
disconnect feature permits rapid removal of the co-pilot collective stick. A spring
pin assembly is provided to ensure positive engagement of the stick. There are
no switches on this collective lever.
Tail Rotor Controls
The tail rotor controls include the control pedals, pedal adjuster, control tubes,
bellcranks and a pitch control mechanism mounted through the tail rotor gearbox
shaft. Moving the pedals causes pitch change in the tail rotor blades to offset the
main rotor torque and to control the directional heading of the helicopter. The
tail rotor control pedals mounted on the pilot’s compartment deck are connected
under the center console to a bellcrank pedal adjuster, which provides for manual
adjustment of pedal position according to the pilot’s needs.
For helicopters with dual controls, the co-pilot’s fully functional tail rotor control
pedal assembly is installed on the floor in front of the co-pilot’s seat to provide a
means for the co-pilot to control the tail rotor assembly. The control pedals are
linked to the pilot’s by means of control tubes and a bellcrank. Also for those
helicopters equipped with dual controls, the co-pilot’s pedals can be positioned,
as desired, by means of the pedal adjuster. Connecting linkage consists of push
pull tubes, bellcranks, levers, and supports that connect the pilot’s tail rotor
control pedals to the tail rotor pitch change mechanism.
Tail rotor pitch control is accomplished by means of bellcrank, rod, and lever
assembly mounted on the tail rotor gearbox which actuates a control tube
through the hollow rotor drive shaft to the crosshead and pitch links.
If an autopilot is fitted, the pedals may be hydraulically assisted.
Using the pedals changes the load on the engine as the thrust is changed in the
tail rotor. Pressing the left pedal will require more power from the engine and the
torque will increase. Pressing the left pedal will cause a decrease in torque. You
should be very careful when operating at all up weight and using high torque
values as an inappropriate left pedal input could cause an over torque.
Control Column
The flight controls are routed beneath the pilot’s and passenger seats aft to the
vertical control column then up to the cabin roof. This control column also
serves as a primary cabin support structure. Access panels on the aft side of the
column, the bottom of the aircraft, and seat panels are provided for inspection of
control components and maintenance accessibility. The cyclic controls are mixed
with collective control through the mixing bellcrank lever located at the base of
the control column.
Swashplate and Collective Lever
The main rotor controls consist of the swashplate and support assembly, sleeve
drive link, and pitch links. The swashplate transfers cyclic control motions from
the non-rotating to the rotating control system. The swashplate and support
encircle the mast directly above the transmission. The swashplate mounts on a
universal support (pivot sleeve or uniball) that permits it tilt in any direction.
Movement of the cyclic results in a corresponding tilt of the swashplate and the
main rotor.
The collective lever and link assembly is mounted to the swashplate support
assembly and transfers collective inputs to the lower swashplate. Movement of
the collective pitch lever actuates the sleeve assembly that in turn raises or
lowers the swashplate and transmits collective input to the main rotor. The
swashplate drive assembly consists of a collar set, idler link, and idler levers.
The collar set is attached to the mast and the idler links are attached to the outer
ring of the swashplate. This connects the upper swashplate to the mast, causing it
to rotate with the mast.
The swashplate and support assembly should be inspected regularly for evidence
of excessive grease leakage from the seals. Also check for excessive movement
or binding of the bearings or any evidence of damage.
The pitch link assembly connects the pitch horn on the blade grips to the
swashplate that transmits control input from both the collective and cyclic
controls.
Hydraulic System
The hydraulic system provides pressurized fluid to operate the cyclic and
collective flight control servo actuators.
NOTE
Helicopters S/N 4 to 497 have a hydraulic powered tail rotor control systems
which are also powered by the hydraulic system.
Helicopters S/N 498 and subsequent do not have hydraulic powered tail rotor
systems, unless an autopilot system is installed. The hydraulic system consists of
the pump/reservoir regulator assembly, filter, solenoid valve, tube and hose
assemblies. The pump, regulator and reservoir are mounted on the forward side
of the transmission oil pump as an assembly. The pump and regulator assembly
includes a mounting pad for the rotor tachometer generator. Operation of the
hydraulic system is electrically controlled by an ON/OFF switch mounted on the
console for pilot control of the solenoid valve. When the solenoid is energized
(ON/OFF switch “OFF”), pressurized hydraulic fluid flows to the reservoir,
bypassing the servo actuators. The tail rotor servo is a customer option and may
be removed.
Operation of the system is electrically controlled by means of the hydraulic
system switch on the miscellaneous control panel (in front of the collective).
When the hydraulic system solenoid valve is de-energized (switch ON), the
pressurized hydraulic fluid flows to the three servo actuators. When it is
energized (switch OFF), the pressurized hydraulic fluid is directed back to the
reservoir (see previous photo) and bypasses the three servo actuators. In case of
a total electrical failure, the system is fail-safe ON.
The cyclic and collective servo actuator support is installed on the cabin roof
forward of the transmission. It serves as a mount for the three servo actuators
and associated bell-cranks. The collective servo actuator is mounted in the center
position and the two cyclic servo actuators are mounted on the outboard
positions.
The primary unit of the hydraulic system is a combined pump and reservoir
assembly located on the left forward side of the transmission. The reservoir filler
opening is located on the top of the reservoir. Fluid level may be observed
through a transparent plastic sight glass in the side of the reservoir and is viewed
from the starboard side. A valve assembly is located on the right hand side of the
system, just above the work deck and incorporates a solenoid by-pass valve and
filter element assembly which incorporates a filter condition red button. If there
is restricted flow, this red button will pop up. This will occur when the pressure
is 70±10 PSI. The hydraulic system is serviced with oil specification MIL-H5606. The reservoir capacity is one pint.
Hydraulic System Schematic
The main hydraulic system contains one power pack and three servo actuators
with irreversible valves. The system operates at 600 PSI and adjustment is
provided at the relief valve and pressure regulator.
The power pack is a self-contained unit located on the forward face of the
transmission and driven by the transmission oil pump shaft. The transmission
driven hydraulic pump provides for fully boosted flight controls being available
during autorotative flight.
The irreversible valves isolate the servo system from the flight controls in the
case of hydraulic malfunction and permit extended periods of flying without
excessive feedback from the controls.
Rotor Brake System
The rotor brake provides a way of stopping the main and tail rotor blades from
turning. It is a completely self-contained hydraulic system.
It is operated by a handle located on the right side of the overhead console. This
applies hydraulic pressure to a disc brake located forward of the accessory
gearbox.
To apply the brake, the handle is pulled down and forward. When shutting down
the helicopter, wait until the rotor RPM is between 38 to 30% before applying
the brake. On icy or frozen ground make sure the brake is applied slowly to
prevent the helicopter slipping on the ground.
10
WEIGHT AND BALANCE
General Weight and Balance
The pilot must find the correct data for weight and balance calculations from the
helicopter flight manual. It is the pilot’s responsibility to ensure that the
maximum gross weight and center of gravity limitations are observed throughout
each planned flight. Operations outside of the limitations are prohibited and may
result in a reduction of aircraft performance, handling qualities, stability, and
structural integrity.
Changes in aircraft configuration (example: doors on or off), loading, seating of
passengers, positioning of cargo, and fuel consumption are all factors that must
be considered in weight and balance calculations.
These calculations must be made for every flight.
Terms
The following terms are used in the calculation of weight and balance
information:
Weight: The actual aircraft weight, weight of crewmembers and passengers, fuel,
etc. is required to accurately compute weight and balance. The flight manual
provides data based on either U.S. or metric measurements, expressed in pounds
or kilograms.
Arm: When calculating weight and balance information, this term refers to the
distance from a known point to where new weight is added. For a given quantity
of weight added, the greater the arm, the greater the affect upon the balance
point. In the 206 B3 flight manual, this data is referred to as fuselage stations or
buttock lines. It is provided in both U.S. and metric formats (inches or
millimeters) and is provided for both longitudinal and lateral axes of balance.
Moment: A mathematical expression of the affect that weight (at a certain
position) will have on the balance point. This number is obtained by multiplying
the weight and the arm. The result is referred to as the moment, and can be
expressed as inch-pounds (U.S.) or kilogram-millimeters (metric).
Center of Gravity: Used to identify the position at which the aircraft is in
balance. Maintenance personnel compute the Empty Weight Center of Gravity.
Pilots use the empty weight CG to determine the Gross Weight Center of
Gravity. Total Moment divided by total weight equals the center or gravity. The
result is expressed in either inches (U.S.) or millimeters (metric).
Datum: An arbitrary point from which to measure the fuselage station (arm) for
longitudinal CG calculations. Fuselage Stations aft of the datum line are
expressed as positive numbers, while points forward would be expressed as
negative numbers. The 206’s Datum Zero is located 55.16 inches forward of the
forward jacking point centerline which equates to approximately 1 inch forward
of the nose. Placement of Datum Zero forward of the nose results in all
longitudinal calculations in positive numbers.
Center line: A reference point for measuring distances from the lateral axis of
the aircraft. Buttock lines are measured from the center line, either to the left or
to the right. Positions to the right of the center line are considered positive
numbers, while positions to the left are considered negative numbers.
Empty Weight and Center of Gravity: Computed by maintenance personnel. The
empty weight configuration consists of the basic helicopter with required
equipment, operational and special equipment, fixed ballast, hydraulic fluid,
transmission and gearbox oil, unusable fuel, and undrainable engine oil. Weight
empty and center of gravity is recorded on the Actual Weight Record, a copy of
which should be carried in the helicopter for use in weight and balance
calculations. Each time equipment is installed or removed, the actual weight
record should be updated.
Gross Weight Center of Gravity: Computed by the pilot. Gross weight includes
the total weight of the helicopter, with contents. Contents include crew,
passengers, engine oil, fuel, baggage, and cargo.
Zero fuel weight (ZFW): A calculated value used as a starting point to calculate
changes in CG caused by fuel consumption during flight. ZFW includes
helicopter empty weight, crew, passengers, engine oil, baggage and cargo.
Effects of Fuel Consumption on Center of Gravity
Under normal circumstances, a center of gravity calculation completed prior to
takeoff will be affected only by fuel consumption during the flight. Other
loading considerations (crew, passengers, and cargo) will not change. If the
aircraft begins a flight with full fuel and lands with minimum fuel, the center of
gravity will move aft as fuel is consumed. It is the pilot’s responsibility to ensure
that the helicopter remains within CG as fuel is consumed.
Baggage Compartment Loading
The baggage compartment is accessible from the left side of the fuselage and
contains approximately 16 cubic feet of space. It has a load limit of 250 pounds
(113.4 kg), not to exceed 86 pounds per square foot (4.2 kilograms/100 square
centimeters). These are structural limitations only, and do not infer that the CG
will remain within approved limits.
When weight is loaded into the baggage compartment, the pilot must calculate
the gross weight and CG to assure loading within approved limits. The load shall
be secured to tie-down fittings if shifting of the load in flight could result in
structural damage to the baggage compartment or in gross weight center of
gravity limits being exceeded. Tables and examples assume items in the baggage
compartment have a longitudinal CG at the midpoint of the door opening.
Weight and Balance Calculation
It is the pilot’s responsibility to ensure that the maximum gross weight and
center of gravity limitations are observed throughout each flight. In the case of
the 206 B3, improper cabin loading combined with fuel consumption may create
a situation where the center of gravity could travel outside CG limits during the
flight. For this reason, it is important to load the helicopter as evenly as possible,
both longitudinally and laterally.
Longitudinal Center of Gravity Limits
The longitudinal center of gravity limits are from station 106.0 to 114.2 inches;
however the aft limit is variable depending upon the gross weight.
When operating with any combination of forward and aft door(s) off; the CG vs
GW limit is station 106.0 to 110.0
NOTE
Station 0 (datum) is located 55.16 inches forward of the forward jack point
center line.
Lateral Center of Gravity Limits
The lateral center of gravity limits are:
3.0 Inches (76.2mm) left of the aircraft center line.
4.0 Inches (101.6mm) right of the aircraft center line.
Doors Off
There is no change from the basic helicopter CG with only the aft cabin doors
removed.
The actual weight change shall be calculated after the doors have been removed.
Ballast may have to be readjusted to maintain the empty weight center of gravity
within allowable limits.
Front Seat Weight
Minimum – 170 pounds (77.1kg)
A minimum crew weight of 170 pounds (77.7kg) is required in the cockpit. Crew
and passengers may be loaded in any sequence without exceeding the gross
weight center of gravity limits if the forward doors are installed. If the forward
doors are removed, the CG must be calculated.
Example
Weight Limits
Maximum gross weight with standard configuration landing gear: 3200 pounds
Maximum gross weight with high skids and emergency floats:
3200 pounds
Maximum gross weight with external loads (up to 1500 pounds):
3350 pounds
Maximum gross weight with standard fixed floats: 3000 pounds
Plot the zero fuel weight and arm on the “Center of Gravity Vs Gross Weight”
chart (see diagram below).
Plot the gross weight and arm on the “Center of Gravity Vs Gross Weight” chart
(see diagram below).
The points plotted should be within the appropriate envelope.
Plot the Longitudinal arm and the Lateral arm for zero fuel and gross weight on
the “Lateral Vs Longitudinal CG Limits” chart (see below).
The points plotted should be within the appropriate envelope.
NOTE
Use ONLY the tables in the flight manual to calculate the center of gravity. It
should be noted that the centre of gravity position changes depending on the fuel
load.
11
PERFORMANCE
Section 4 of the Rotorcraft Flight Manual contains the Bell 206 performance
information and related charts. All performance charts are based on an engine
meeting minimum Rolls-Royce specifications. The data shown is derived from
actual flight tests and are intended to provide information to be used in
conducting flight operations. This performance data is applicable to the 250C20B/C20J engine.
NOTE
The diagrams displayed in this section of the book are for demonstration
purposes only. The performance figures are estimated and should never be used
for planning purposes. For all planning refer to the helicopter flight manual.
Power Check Procedures
A power check chart is provided for the Rolls-Royce 250C-20J engine. This
chart indicates the minimum percent torque that must be available for an engine
that meets minimum specifications. The engine must develop these values in
order to meet performance data contained in this section.
The takeoff power limits of the 250-C20B, C20J are:
Maximum Torque – 100% (5 minutes)
Maximum TOT (Turbine Outlet Temperature) - 810°C (5 minutes)
Maximum Gas Producer RPM (N1) – 105%
NOTE
Accurate power checks may also be accomplished in a stabilized 52 knot IAS
climb by adding 2% to the chart percent torque reading. If this power check
proves unacceptable then perform the power check from a hover when altitude,
temperature and gross weight permit a safe hovering height. (Refer to the
Height/Velocity diagram.)
More accurate power checks are achieved above Maximum Continuous TOT
(738°C). To avoid exceeding engine torque limits, this procedure will generally
require a climb above 5000 feet.
On cold days the torque pressure limit may be reached before the TOT limit is
reached. On hot days or at high altitudes, the TOT will be the limiting factor. To
perform a power check, ensure the anti-ice and generator switches are OFF.
Raise the collective to increase power until a stabilized TOT or torque pressure
limit is reached. Record the Outside Air Temperature (OAT), TOT, Pressure
Altitude, Torque and N1.
Refer to the Power Check Chart.
NOTE
The power check is acceptable when the chart percent torque reading is equaled
or exceeded.
Rate of Climb
The rate of climb data shown in the graphs is termed “TAPE-LINE RATE OF
CLIMB” which means actual rate of climb. The rate of climb as measured with
an altimeter will show tape-line rates of climb only on a standard day with a
standard lapse rate which is a very rare occurrence.
The following example is for use with Rate of Climb – Maximum with takeoff
power. The example is typical for use with Rate of Climb Charts.
EXAMPLE:
Assume an ambient OAT of 10°C, a pressure altitude of 12,000 feet and a gross
weight of 3000 pounds. Determine the maximum rate of Climb.
PART 1 – ANTI-ICE OFF
Enter the temperature scale at 10°C and proceed vertically to the intersection of
the 12,000 feet pressure altitude curve. From this point move horizontally to the
right to intersect the 3000 pound gross weight line. Drop vertically and read antiice OFF rate of climb of 860 feet per minute.
PART 2 – ANTI-ICE ON
From the intersection of the horizontal example line at 3000 pounds gross
weight, proceed vertically to the upper R/C Ft/MIN section of the chart diagonal
line and then move horizontally to the right and read 250 ft/min. Subtract 250
ft/min from the 860 ft/min (anti-ice OFF R/C) and the anti-ice ON R/C is
determined to be 610 ft/min for the same 3000 pound gross weight.
Rate of Climb – Door(s) off
Reduce the basic rate of climb chart data by 350 ft/min when operating with one,
all or any combination of cabin doors off.
Operation in Allowable Relative Wind
Satisfactory stability and control has been demonstrated in relative winds of 20
MPH (17 KTS) sideways and rearwards at all loading conditions within Area A
of the Hover Ceiling Charts. Note that this is not a limitation.
IGE and OGE Hover Ceiling Charts
(Refer to Fig 28 on following pages)
NOTE
The Hover Ceiling Charts presented in the following pages reflect performance
with the 65 inch diameter tail rotor installed. For performance with the 62 inch
diameter tail rotor refer to the relevant section in the helicopter flight manual.
The hover Ceiling In Ground Effect charts and Hover Ceiling Out of Ground
Effect charts present hover performance (allowable gross weight) for conditions
of pressure altitude and OAT. The charts are divided into two areas. They must
be read with reference to the Critical Azimuth Chart.
AREA A
(White Area) as shown on the Hover Ceiling Charts present hover performance
for which controllability has been demonstrated regardless of the wind direction.
CAUTION
THE ENGINE TOT WILL RISE NOTICEABLY WHEN HOVERING
DOWNWIND. AVOID HOVERING DOWNWIND WHEN OPERATING NEAR
TOT LIMITS.
AREA B
(Yellow Area) as shown in the Hover Ceiling Charts.
The helicopter can hover in any wind direction (wind speed up to 17kt) if using
the 65” tail rotor.
If using the 62” tail rotor, hover performance can be achieved in calm winds.
Extra performance can be achieved by not having the wind in the critical area.
i.e. 050° to 210°.
AREA C
(Blue Area) as shown in the Hover Ceiling Charts.
Presents hover performance that can be achieved in CALM WINDS or winds
outside the CRITICAL RELATIVE WIND AZIMUTH AREA with either the
65” or 62” diameter tail rotor.
What all of this means is that if you are expecting to end up in the hover with the
relative wind coming from 050 - 210°, only use area A of the chart.
Hover Ceiling
The following example is for use with the Hover Ceiling Out of-Ground-Effect
(OGE), Takeoff Power, Anti-ice Off chart and is typical for use of Hover Ceiling
charts.
For actual figures – refer to the relevant chart in the flight manual.
EXAMPLE
Determine the gross weight OGE hover capability at a site having the following
conditions:
Pressure altitude
= 10000 feet
Outside Air Temperature
= 20°C
For the above example the pilot must refer to the OGE 0°C to 46°C Hover
Ceiling chart.
From the appropriate IGE chart obtain:
A maximum of 2250 pounds for all allowable wind conditions.
A maximum of 2530 pounds in calm winds or outside the critical wind azimuth
area if operating with the 62” diameter tail rotor.
A maximum of 2920 pounds in calm winds or outside the critical wind azimuth
area if operating with either the 62” or 65” diameter tail rotor installed.
Height Velocity Envelope
The height velocity envelope chart defines the conditions from which a safe
landing can be made on a smooth, level, firm surface following an engine
failure. The chart is based on zero wind conditions. A sample diagram is shown
below. This diagram should not be used for planning. Refer to the helicopters
flight manual for exact figures.
Power checks
When working above 6000 feet the torque will not normally limit you as the
TOT will be the limiting factor.
Depending on the AUW of the helicopter, pilots are often unsure if they have
enough power available to come to the hover when making an approach to land.
The landing site may be at a different elevation than your departure point and
you cannot normally refer to the helicopter flight manual in flight as you are
using both hands.
A reasonable power check can be achieved by flying straight and level at 70 kts
and approximately 500’ above the landing site elevation. Take a note of the
torque. You will need 20% more than the noted value to hover in ground effect.
You will need 30% more than the noted value to climb back out.
If you are taking on more passengers or luggage then add 1% torque for every
15kg loaded. (e.g. A 200 lb person is equivalent to 6% torque.)
12
NORMAL PROCEDURES
Introduction
This section contains instructions and procedures for operating the helicopter
from the planning stages, through actual flight conditions, to securing the
helicopter after landing.
Normal and standard conditions are assumed in these procedures. Pertinent data
in other sections is referenced when applicable.
Instructions and procedures contained herein are written for purpose of
standardisation and are not applicable to all situations.
Operating Limitations
The minimum and maximum limits and normal and cautionary operating ranges
for the helicopter and its subsystems are indicated by instrument markings and
placards. The instrument markings and placards in the aircraft represent careful
aerodynamic calculations that were substantiated by flight test data. Refer to the
LIMITATIONS section in the aircraft Flight Manual, for a detailed explanation
of each operating limitation.
Any time an operating limitation is exceeded, an appropriate entry shall be made
in the helicopter logbook. The entry shall state which limit was exceeded, the
duration of time, the extreme value attained and any additional information
essential in determining the maintenance action required.
Flight Planning
Each flight should be planned adequately to ensure safe operations and to
provide the pilot with data to be used during the flight.
Check type of mission to be performed and destination.
Select the appropriate performance charts to be used from Section 4,
PERFORMANCE, of the helicopter flight manual.
Take Off and Landing Data
Refer to Section 10 (in this book) for takeoff and landing weight limits, and to
Section 11 (in this book) for performance information.
Weight and Balance Checks
Determine the exact weight and balance of the helicopter as follows:
1.
Consult the helicopter flight manual for instructions.
2.
Compute the takeoff and anticipated landing gross weight. Check the
helicopter center of gravity (CG) locations and determine the weight of the fuel,
oil, payload etc.
3.
Ensure the weight and balance limits listed in Section 1 of the flight
manual have not been exceeded.
Preflight Check
The pilot is responsible for determining whether the helicopter is in condition for
safe flight.
NOTE
The pre-flight check is not intended to be a detailed mechanical inspection, but
simply a guide to help the pilot check the condition of the helicopter. It may be as
comprehensive as conditions warrant at the discretion of the pilot.
All areas checked shall include a visual check for evidence of corrosion,
particularly when the helicopter is flown near or over salt water or in areas of
high industrial emissions.
Before Exterior Check
1. Flight planning - Completed.
2. Publications - Checked.
3. GW and CG - Computed.
4. Helicopter servicing - Completed.
5. Battery - Connected.
Exterior Check List
1. Fuselage – Cabin Right Side
Right static port - Condition.
Cabin doors - Condition and security.
Windows - Condition and security.
Landing gear - Condition and ground handling wheel removed.
Forward and aft cross tube fairings (if installed) - Secured, condition, and
aligned.
2. Fuselage – Center Right Side
Cabin roof, transmission fairing and engine air inlet area – Cleaned of all debris
and accumulated snow and ice.
Hydraulic reservoir – Oil level.
Hydraulic system filter – Bypass indicator retracted.
Hydraulic actuators and lines – Condition, security, interference and leakage.
Forward fairing – Secured.
Access door – Secured.
TRANSMISSION AREA
Transmission – Oil level and area for leaks.
Transmission and mounts – Security and condition.
Isolation Mount – Condition.
Drag pin – Security and evidence of contact with static stop plate.
Main driveshaft forward coupling – Condition and grease leakage. Check paint
strips for evidence of overheating indicated by brown colour.
Access door – Secured.
Air induction Cowling – Secured.
Engine inlet and plenum – Condition and clear of obstructions.
Fuel filler cap – Visually check level. Cap secure.
Fuel sump – Drain fuel sample as follows:
FUEL BOOST pump circuit breakers – OUT.
BATtery switch – ON.
FUEL VALVE switch – OFF.
Fuel drain or button – Depress, drain sample and then release.
NOTE
Apply the following procedure to the airframe fuel filter kit and/or engine fuel
pump filter.
A/F fuel filter (if installed) – Drain and check before the first flight of the day as
follows:
FUEL VALVE switch – ON.
FUEL BOOST pump circuit breakers – IN.
CAUTION LT circuit breaker – IN
Fuel filter drain valve – Open, drain sample, then close.
NOTE
Filter test button is located on top of fuel filter.
Fuel filter test button — Press and check A/F FUEL FILTER caution light
Illuminated. Release switch and check light extinguished.
FUEL VALVE switch — OFF.
BAT switch — OFF.
POWER PLANT AREA
Main driveshaft aft coupling – Condition and grease leakage.
Check paint strips for evidence of overheating indicated by brown colour.
Engine – Condition. Security of attachments.
Engine mounts – Condition and security.
Throttle linkage – Condition, security and freedom of operation.
Fuel control and governor – Evidence of leakage.
Hoses and tubing – Chafing, security and condition.
Exhaust stack and clamp – Security and condition.
Engine cowl – Secure.
Generator cooling scoop – Clear of debris.
Right exhaust cover – Removed.
Oil tank – Oil level, leaks, security and cap secure.
Access door – Secured.
Aft fairing – Secured.
3. Fuselage – Aft Right Side
Fuselage – Condition.
Tail rotor driveshaft cover – Condition and secured.
Tail boom – Condition.
Horizontal stabiliser and position light – Condition and security.
Main rotor blade – Condition.
4. Fuselage – Full Aft
Vertical fin – Condition.
Tail rotor guard – Condition and security.
Anti-collision light – Condition.
Aft position light – Condition.
Tail rotor gearbox – Oil level, leaks and security.
Tail rotor – Tie down removed, condition and free movement.
Tail rotor controls – Condition and security.
Tail rotor blades – Condition. Tip block for rivet damage, corrosion and seal
condition.
5. Fuselage – Aft Left Side
Main rotor blade – Tie down removed. Condition.
Tail boom – Condition.
Tail rotor driveshaft cover – Condition and secure.
Horizontal stabiliser and position light – Condition and security.
Fuselage – condition.
Baggage compartment – Cargo tied down – Door secure.
Forward tail rotor driveshaft coupling – condition of splined adapter and
freedom of rotation of witness pin.
Oil cooler blower shaft hanger bearings – Grease leakage and overheating.
Oil cooler blower – Clear of obstruction and condition.
Oil cooler – Condition and leaks.
Oil cooler access door – Secured.
Aft fairing – Secured.
Left exhaust cover – Removed.
POWERPLANT AREA
Engine – Condition. Security of attachments.
Engine mounts – Condition and security.
Exhaust stack and clamp – Condition and security.
Evidence of fuel and oil leaks.
Hoses and tubing for chafing and condition.
Linear actuator and governor control linkage – Condition and security.
Engine anti-ice valve and linkage – Condition and security.
Tail rotor driveshaft – Condition of splines, couplings and freedom of
movement.
Engine cowling – Secured.
Air induction cowling – Secured.
Engine inlet and plenum – Condition. Clear of obstructions.
TRANSMISSION AREA
Transmission and mounts – Condition and security. Check area for leaks.
Isolation mount – Condition.
Access door – Secured.
6. Fuselage – Cabin Left Side
Cabin roof, transmission fairing and engine air inlet area – Cleaned of all debris
and accumulated snow and ice.
Rotor head reservoirs (oil lubricated) – Visible oil levels.
Main rotor hub and yoke – condition and cracks.
Pitch horn trunnion bearing – Wear and security.
Main rotor pitch links – Condition, cracks and security of attachment bolts and
locking hardware.
Swashplate assembly – Condition, security of attached controls and boot
condition.
Control linkages to swashplate – Condition, security of attachment bolts and
locking hardware.
Forward fairing and access door – Secured.
Cabin doors – Condition and security.
Windows – Condition and security.
Landing gear – Condition. Ground handling wheel removed.
Left static port – Condition.
7. Fuselage – Front
Exterior surfaces – Condition.
Windshield – Condition and cleanliness.
Battery – Connected.
Battery and vent lines – Condition and security.
Battery access door – Secured.
Pitot tube – Cover removed. Clear of obstructions.
External power door – Condition and security.
Landing light glass – Condition.
Antennae – Condition and security.
Main rotor blade – Condition.
External power – Connect (if desired). BATtery OFF.
Interior Check List
Cabin interior – Cleanliness and security of equipment.
Fire extinguisher – Secure. Charge indicator in green arc.
First aid kit – Secure.
Co-pilot controls – Secure and properly installed.
Co-pilot seat belt – Secured (if solo).
Cabin loading – Refer to Section 6 of the flight manual (Weight and Balance).
Cabin doors – Secured.
Engine Pre-start Check List
NOTE
Helicopters serial numbers 3567 and subsequent are equipped with ENGINE
ANTI-ICING and HYDRAULIC SYSTEM switches. On helicopters prior to this,
switches are labeled ENGINE DE-ICING and CONTROL BOOST, respectively.
Flight controls – Release friction. Check freedom of movement and adjust to
(cyclic) neutral; (collective) flat pitch position and pedals neutral.
Throttle – Check freedom of full travel and flight idle stop operation. Check copilot throttle if installed. Return to closed position.
A good visual check that the throttle is closed is to note the position of the screw
on the throttle in front of your hand. It should be to the right of the Idle Release
Button.
LDG LTS switch – OFF.
ENGINE DE-ICING or ENGINE ANTI-ICING switch – OFF.
CONTROL BOOST or HYDRAULIC SYSTEM switch – ON.
FUEL VALVE switch – ON.
Altimeter – Set to field elevation.
Instruments/Gauges – Static position at zero.
Overhead switches – OFF.
NOTE
Effective helicopter S/N 4128 and prior: for daylight operations, ensure INST LT
switch (rheostat) is OFF. If the INST LT switch is on, the caution lights can be
dimmed and may not be visible.
Effective helicopter S/N 4129 and subsequent: with the INST LT switch
(rheostat) ON and caution light selector positioned to DIM, the caution lights
are dimmed to a fixed intensity and can not be adjusted by the INST LT switch.
Generators switch – OFF.
Circuit breakers – In (as required).
BATtery switch – ON (OFF if external power used).
Observe TRANS OIL PRESS, ENG OUT, and ROTOR LOW RPM
caution/warning light segments illuminated and applicable audio signal(s)
operative.
WRN HORN MUTE button (if installed) – Press to mute.
NOTE
Engine out audio may be deactivated.
Warning and caution segments – test to check illumination of each segment
utilised.
Turbine outlet temperature (TOT LT TEST) button (if installed) – Press. Check
TOT light illuminates.
ROTOR LOW RPM system – Check as follows: (if WRN HORN MUTE button
is installed, the following does not apply).
Collective pitch – Increase. Check ROTOR LOW RPM light and audio ON.
Collective pitch – Full down. Check ROTOR LOW RPM light ON and audio
OFF.
Flight controls – Neutral/flat pitch position. Apply friction (if needed).
FUEL BOOST pump circuit breakers – IN. Check fuel pressure within limits and
FUEL PUMP caution light OFF.
ANTI-COLLision light switch – ON (if required).
Engine Starting
Collective pitch – Full down.
Throttle – Full Closed.
Rotors – Clear.
STARTER – Engage (see starter limits, Section 1).
Engine oil pressure – Indication of increase.
Throttle – Open to flight idle at 15% gas producer (N1) RPM with Turbine
Outlet Temperature (TOT) at or below 150°C.
CAUTION
A start should not be attempted at N1 speeds below 12%.
Use the following guide for desired N1 starting speed versus outside air
temperature:
Above 7°C (45°F) N1 RPM 15%
-18°C to +7°C (0° to 45°F) N1 RPM 13%
Below -18°C (0°F) N1 RPM 12%
NOTE
During the first few seconds of the start the TOT will accelerate at a fairly rapid
rate and must be pilot monitored.
Turbine outlet temperature (TOT) – Monitor to avoid hot start. Abort the start if
either the 927°C maximum or the 810 to 927°C MAXIMUM 10 SECONDS
transient limitation is about to be exceeded by depressing the engine IDLE
RELease button. CLOSE THROTTLE and continue to motor the starter until
TOT decreases to less than 810°C. Some helicopters may be equipped with a red
warning light on the TOT gauge. If limits are exceeded or light illuminates,
consult Allison Engine Operation and Maintenance Manual.
CAUTION
If the main rotor is not rotating by 25% gas producer speed (N1), abort the start.
STARTER – Release at 58% gas producer RPM (N1).
Engine and transmission oil – Check pressures increasing.
CAUTION
If the engine has been shut down for more than 15 minutes, stabilise at idle
speed for one minute before increasing power.
NOTE
During cold temperature operations stabilize engine at idle speed of 60 to 62%
gas producer RPM (N1) until oil temperature reaches 0°C.
Gas producer RPM (N1) – Check for 60 to 62%.
External power – Remove – BATtery ON.
Throttle – Open to 70% gas producer RPM.
GENerator switch – ON.
Radio equipment – ON.
ELT (if installed) – Check for inadvertent transmission.
POSition LT switch – ON.
ANTI COLLision LT switch – ON.
NOTE
Uncommanded control movement or motoring with hydraulic system off may
indicate hydraulic system malfunction.
HYDRAULIC SYSTEM or CONTROL BOOST switch – OFF then ON.
Smoothly and firmly advance throttle at a continuous rate to full open position
maintaining collective pitch down and cyclic control in neutral.
Power turbine (N2) governor – Check range 97 to 100% RPM.
Engine Run-up Check
NOTE
If temperature is 4.4°C (40°F) or below and visible moisture is present, the
engine anti-icing system shall be ON.
ENGINE DEICING or ENGINE ANTI-ICING switch – ON (if conditions
warrant and observe TOT rise).
PITOT HEAT switch – ON (if installed) in visible moisture with temperature
below 4.4°C (40°F).
NOTE
The Hydraulic Systems Check is to determine proper operation of the hydraulic
actuators for each flight control system. If abnormal forces, unequal forces,
control binding or motoring are encountered it may be an indication of a
malfunction of a control actuator.
Collective – Full down. Friction removed.
Rotor RPM (Nr) – Set to 100%.
HYDRAULIC SYSTEM or CONTROL BOOST switch – OFF.
Cyclic – Centered. Friction OFF.
Check normal operation of cyclic control by moving cyclic in an “X” pattern
right forward to left aft then left forward to right aft (approximately one inch).
Center cyclic.
Collective – Check for normal operations by increasing collective control
slightly (1 to 2 inches). Repeat 2 to 3 times as required. Return to full down
position.
Pedals (if hydraulically boosted) – displace slightly left and right. Note an
increase in force required to move pedal in each direction.
HYDRAULIC SYSTEM or CONTROL BOOST switch – ON.
Cyclic and collective friction – Set as desired.
The Startup in More Detail
When converting to a JetRanger having only ever flown piston engine
helicopters, the startup can seem quite intimidating at first. Most pilots have
heard lots of stories about how startups can go wrong. However the startup is a
simple process as long as certain procedures are followed and these will be
described below. It is important that you use the checklist for startup however
you should know this checklist from memory. You should also be very
comfortable in handling a startup emergency (as in any helicopter) and therefore
you should memorize the appropriate emergency procedures.
The following startup sequence assumes that all prior checks have been carried
out.
After securing your safety harness look outside and make sure that the main
rotor blades are in the three o’clock and nine o’clock position. This will ensure
that the aft tie-down has been removed.
Release the cyclic and collective frictions and check the controls for freedom of
movement. They will feel quite stiff as you force the hydraulic oil through the
system.
Check the pedals are in the correct position for your length of legs and also
check freedom of movement.
Open the throttle for full travel and make sure it is returned to flight idle stop.
Center the cyclic and apply friction.
Collective fully down and apply friction.
Pedals centered.
The landing lights should be turned off as these will drain the battery when the
BATtery switch is turned on.
Rotor brake should be OFF.
The ENGINE ANTI-ICING switch (located below the forward end of the
collective) should be in the OFF position.
The HYDRAULIC SYSTEM switch (located below the forward end of the
collective) should be in the ON position.
Turn on the FUEL VALVE (located on the instrument panel beneath a red
guard).
All of the overhead switches should be in the OFF position.
Normally all the circuit breakers should be in the IN position but this will be as
required for your aircraft.
Turn the BAT switch (located in the overhead console) ON. Press the WRN
HORN MUTE button to silence the warning horn. If this button is not fitted then
pull the caution circuit breaker.
If a battery cart is being used for the startup then the BAT switch should be left
in the OFF position.
Press the CAUTION LT TEST button (located to the right of the caution warning
panel) and check that all segments illuminate.
If installed, check the TOT LT TEST by pressing its associated button. The
position of this button can vary but is normally near the TOT gauge.
Pull each of the FUEL BOOST circuit breakers to the out position. Push one IN
and check for fuel pressure on the fuel pressure gauge. (The fuel pump caution
light should illuminate with one fuel boost pump inoperative). Pull this CB to the
OUT position.
Repeat for the other CB.
Push both FUEL BOOST circuit breakers to the IN position.
If required, turn on the avionics and request permission for startup.
Turn on the ANTI COLL LT located behind the BAT switch.
Double check that the throttle is fully closed and not at the flight idle stop. If the
throttle is not fully closed then fuel will be introduced to the engine before it is
up to a speed to draw sufficient air through the engine for cooling. This is one of
the very few mistakes that could lead to a hot start.
Before starting the engine you must check the TOT gauge. The temperature must
be below 150°C. If the helicopter has been recently shut down then the
temperature may be above 150°C.
DO NOT ATTEMPT TO START IF TOT IS GREATER THAN 150°C.
Check the surrounding area is clear.
If the engine needs to be cooled you must “vent” the engine by pressing the
starter button while holding the throttle in the fully closed position. This will
blow cooler air through the engine and the TOT will drop.
Wait until the gas producer (N1) returns to zero before continuing with the
startup.
Hold the starter button down. You will not release this button again until the gas
producer gauge reads 58%.
At 15% N1, gently roll on the throttle. Fuel will then be introduced to the engine
and ignition will begin.
All of your attention must now be fully on the TOT gauge. This needle will rise
very quickly and may go into the red band. It will normally have two peaks
during the startup.
You must remember the limitations for startup. It is normal for the needle to
reach the top of the green band and enter the yellow band. The needle may go
above 810°C and up to 927°C (red band) for a maximum of 10 seconds. It must
not be allowed to go above 927°C.
NOTE
If you think that the 10 second limit or the 927°C limit is going to be exceeded
then you must continue to hold in the starter button and press the idle release
button on the throttle and close the throttle. Continue motoring the engine until
the temperature drops.
The TOT gauge needle will peak and then start to lower.
You must then bring your attention back to the N1 gauge. At 25% the blades
should be turning. If the blades are not turning then shut down the engine. This
would only happen if the main rotor blades were still tied down.
A quick check of the engine oil pressure should show it rising.
Now focus on the N1 gauge again. At 58% you must release the starter button as
the combustion is self sustaining.
Open the throttle slightly to accelerate the turbine needle (N2) through the
yellow band and then close throttle to flight idle.
N1 should be 60 to 62%. If N1 is less than 60% then get an engineer to look at
the engine as there is a possibility of a flameout during flight especially if
practicing autorotations.
NOTE
If external power is being used, it should now be disconnected and the battery
switch should be turned ON.
Start the clock and wait one minute. (If it is less than 15 minutes since the last
shutdown then the one minute warm up does not apply). Check the temperatures
and pressures.
Increase power to 70%N1. Do not exceed 40% torque as the helicopter may start
to slip on the helipad.
Turn the GENerator switch on and check the load meter reads less than 70%.
If the CAUTION circuit breaker has been pulled then now is the time to push it
back IN.
Turn on the GYRO switch and cage the gyros (if fitted).
Turn on radio equipment.
Turn on lights as desired.
Hold the cyclic firmly and rest your left wrist on the collective while you turn off
the HYDRAULIC SYSTEM switch. You are feeling for a “kick” on the controls
which would indicate a problem. Turn the switch back on again.
Without exceeding 40% torque, advance the throttle continuously to the fully
open position (100% N2)
Check the operation of the governor by moving the “beeper” switch on the
collective. Make sure the N2 needle is set at 100%.
Release the cyclic and collective frictions.
Hold the cyclic firmly and rest your wrist on the collective again in preparation
for another hydraulic check.
Turn the HYDRAULIC SYSTEM switch OFF.
Move the cyclic approximately one or two inches in a “X” pattern. There should
be no uneven forces or binding of the controls. If there is a problem with the
movement of the controls at this stage then shut the helicopter down in the
normal manner. By informing the engineer of the direction you were moving the
cyclic when the problem occurred, he will be able to determine which servo
requires attention.
Move the collective up and down approximately 1 to 2 inches and check for
motoring or unusual forces. Repeat 2 or 3 times.
Turn the HYDRAULIC SYSTEM switch ON.
Set the cyclic and collective frictions as desired.
The startup sequence is now complete.
Before Takeoff
Electrical equipment – Check. Reset as required.
Lighting – As required.
Radio – Check as required.
Throttle – Full open.
Power and flight instruments – Normal operating range.
Generator load – Below 70% (Note – normal load is 10 to 20%).
Power turbine (N2) – Set for 100% in flat pitch.
Takeoff
Collective pitch – Increase to hover.
Directional control – As required to maintain desired heading.
Cyclic control – Apply as required to accelerate smoothly.
Collective – Apply as required to obtain desired rate of climb and airspeed.
Monitor engine limits and adjust collective as necessary.
In Flight Operations
Airspeed – As desired (not to exceed Vne at flight altitude or maximum
allowable for door(s) off flight configuration).
PITOT HEAT switch – ON (if installed) in visible moisture with temperature
below 4.4°C (40°F).
ENGINE DEICING or ENGINE ANTI-ICING switch – ON in visible moisture
when temperature is below 4.4°C (40°F).
NOTE
TOT will increase when ENGINE ANTI-ICING is switched ON.
Descent and Landing
Flight controls – Adjust friction as desired.
Throttle – Full open.
Power Turbine RPM (N2) – 97 to 100%.
NOTE
Decreasing the collective pitch into the low power realm may result in an RPM
overspeed condition. For prolonged low power approaches the RPM can be
controlled by a small amount of collective pitch increase (no significant torque
increase) and/or by beeping down the N2 governor speed controller to 100%
RPM. This will maintain governing within limits (see diagram below) during low
power descents however the GOV RPM switch should be beeped to INCR as
collective is applied.
Flight path – As required for type of approach being made.
LDG LTS switch – On as required.
Engine Shutdown
Throttle – Flight idle – Check engine deceleration time.
WRN HORN MUTE button (if installed) – Press to mute.
Flight controls – Position for shutdown. Apply frictions.
ENGINE DEICING or ENGINE ANTI-ICING switch – OFF.
TOT – stabilised at flight idle for two minutes.
ELT (if installed) – Check for inadvertent transmission.
IDLE RELease button – depress and roll throttle firmly to full closed position.
CAUTION
To ensure engine cutoff, hold throttle in closed position until N1 decelerates to 0
and TOT is stabilizing. Do not turn BATtery off until N1 is 0 and TOT stabilised.
TOT – Check decreasing.
During rotor coast down apply cyclic to minimise static stop contact.
Radio equipment – OFF.
FUEL VALVE switch – OFF
GENerator switch – OFF.
All switches – OFF.
BATtery switch – OFF after N1 is zero and TOT stabilised.
Pilot should remain at flight controls until rotor has come to a complete stop.
After Exiting Helicopter
Install the main and tail rotor tie downs if any of the following conditions exist:
High or gusting winds are predicted.
Other helicopters are operating or expected to be operating in the
immediate area.
Any time the helicopter is to be left unattended.
Install protective covers (engine inlet, exhaust and pitot tube).
Check fluid levels.
Perform a general walk round of the helicopter.
Covers and Tie downs
Protective covers and tie downs are furnished as loose equipment and are used
for the parking and mooring of the helicopter. Additional equipment such as
ropes, cables, devices, or ramp tie downs will be required during mooring.
For winds up to 50 knots
Secure the main and tail rotor blades. (Do not tie the blades to the extent that
they become flexed).
Insert the engine inlet plugs and the engine exhaust covers.
Place the pitot tube cover over the pitot tube.
Tighten the flight control friction locks.
Make sure that all switches are in the OFF position.
Disconnect the battery.
Make sure all doors, windows, vents and access panels are closed.
If the aircraft is to be parked in a heavy dew environment then all exposed
control bearings will need to be purge lubricated every seven days to ensure no
voids exist that could trap moisture.
For winds greater than 50 knots the helicopter should be moored.
There is a risk of damage to the helicopter due to objects blown through the air
by the strong wind. For this reason the helicopter should be hangered. If the
winds are forecast to be greater than 75 knots, the helicopter should be flown to
a safe weather area.
If there is a suitable, paved parking area with built in tie-down rings then
position the helicopter between these with the nose pointing in the direction from
which the highest forecast wind is expected. Tie the helicopter to the ramp tiedowns.
If suitable tie-down rings are not available then park the helicopter in an
unpaved area; again with the nose pointing in the direction from which the
highest forecast wind is expected. Remove the ground handling wheels and use
heavy weights to help secure the lines to the helicopter.
If there is sufficient time, the main rotor blades should be removed and stored in
a secure building. The main rotor hub should then be secured to the mast to
prevent movement in the flapping axis.
Fill the fuel tank to the maximum capacity as this will help make the helicopter
more stable in the strong winds.
Insert the engine inlet plugs and the engine exhaust covers.
Place the pitot tube cover over the pitot tube.
Tighten the flight control friction locks.
Make sure that all switches are in the OFF position.
Disconnect the battery.
Make sure all doors, windows, vents and access panels are closed.
Cover – Engine Inlet and Pitot Tube
The engine inlet plugs are red, flame resistant, and each cover is attached with a
red streamer stenciled in white letters, “REMOVE BEFORE FLIGHT.” A strap
normally joins the pitot tube cover and engine inlet covers into one assembly.
Cover the pitot tube first, then press inlet plugs into engine air inlet. Tie the pitot
tube cover securely with the cord attached.
Cover – Engine Exhaust
The engine exhaust cover is red, flame resistant, and includes a red streamer
stenciled in white letters, “REMOVE BEFORE FLIGHT.” A tie cord (or Velcro)
is stitched to the cover for securing to the engine exhaust.
Tie down – Main Rotor
The main rotor tie down boot includes a bright red nylon strap with shot bags
attached to the ends to assist in retrieving the main rotor blade for tie down. A
strap is sewn to the boot and is stenciled in white on both sides, “REMOVE
BEFORE FLIGHT.”
After the main rotor blades have stopped turning, toss the lead shot bag ends of
the straps over one main rotor blade near the tip. This will aid in pulling the
blade tip down to install the boot. Insert the metal hook into the eyelet on the end
of the blade. Rotate the main rotor blades in the opposite the direction of rotation
until the main rotor blades are aligned with the vertical fin. Bring the strap to the
port side of the vertical stabilizer and then draw the blade down lightly against
the static stop. Wrap the strap around the “stinger” to secure.
Tie down – Tail Rotor
The tail rotor tie down is red and is stenciled in white letters, “REMOVE
BEFORE FLIGHT.” To tie down the tail rotor assembly, rotate the main rotor
until the tail rotor blades are aligned with the tail boom or with the vertical fin
and the main rotor blades are aligned with the tail boom. Tie down the main
rotor first and then secure the tail rotor to the tail boom or to the vertical fin with
the tie-down strap.
CAUTION
DO NOT TIE DOWN THE TAIL ROTOR TO THE EXTENT THAT TAIL ROTOR
BLADE FLEXES.
Ground Handling
Ground handling wheels may be attached to the two lugs on the skid tubes. Care
should be exercised to make sure that the ground handling wheels are properly
maintained and securely attached to the skids, prior to lifting the helicopter.
The wheels must be removed before flight.
13
EMERGENCY PROCEDURES
This section contains fault conditions considered to constitute an emergency
condition. Warning lights are located on the instrument panel and provide the
pilot with a visual indication of a condition, fault or system malfunction by
means of an individual system light. Illumination is an indication that a problem
has occurred, which unless treated properly, could affect flight safety. Therefore
remedial action as described below should be taken with the urgency each
situation warrants. Listed below are the segments that will illuminate individual
warning lights which indicate and emergency condition.
Definitions
The following terms indicate the degree of urgency in landing the helicopter.
Land as soon as possible - Land without delay at nearest suitable area (i.e. open
field) at which a safe approach and landing is reasonably assured.
Land as soon as practicable - The landing site and duration of the flight are at the
discretion of the pilot. Extended flight beyond the nearest approved landing area
is not recommended.
Warning Light (RED) Segments
ENG OUT: Engine power failure (N1 less than 55%). Reduce pitch immediately
to autorotate. If ample altitude remains investigate the failure and attempt engine
relight.
BATTERY HOT: Battery case temperature has reached 140°F (60°C) or higher.
Turn the battery switch OFF. Land as soon as possible. After landing do not use
helicopter the battery for an engine restart as this will cause additional battery
heating.
Caution Light (AMBER) Segments
ROTOR LOW RPM: Rotor RPM is below normal (approximately 90%). Reduce
collective pitch and check that the throttle is fully open.
TRANS PRESS: Main transmission pressure is below minimum. Check gauge.
Land as soon as possible.
TRANS TEMP : Main transmission oil temperature is at or above the red line.
Check gauge. Reducing power will help alleviate the condition. Check the
transmission oil pressure. Land as soon as possible.
BATTERY TEMP: Battery case temperature has reached 130°F (54.5°C) or
higher. Move the battery switch to OFF until battery cools (light extinguishes)
then battery switch ON.
NOTE
Frequent and repetitive BATTERY TEMP indications may be indicative of a
marginal battery condition. It is recommended that if this occurs the battery
should be removed and inspected in accordance with the manufacturer’s
recommendation at the first convenient opportunity.
ENG CHIP: Metallic particles have been detected in the engine oil. Land as soon
as possible.
TRANS CHIP: Metallic particles have been detected in the transmission oil.
Land as soon as possible.
T/R CHIP: Metallic particles have been detected in the tail rotor gearbox oil.
Land as soon as possible.
FUEL FILTER: The engine fuel filter is clogged. Land as soon as practical.
Clean before next flight.
A/F FUEL FILTER: The airframe fuel filter is clogged. Land as soon as
practical. Clean before next flight.
FUEL PUMP: One or both fuel boost pumps are inoperative. Descend to below
6000 feet pressure altitude if flight permits.
WARNING
Operation with both fuel boost pumps inoperative is not approved. Due to
possible fuel sloshing in unusual attitudes or out of trim conditions. If one or
both fuel boost pumps are inoperative, the unusable fuel is 10 US gallons.
NOTE
The engine will operate without boost pump pressure when below 6000 feet
pressure altitude and one boost pump will supply sufficient fuel for normal
engine operations under all conditions of power and altitude. Both fuel boost
pumps should be ON for all normal operations.
FUEL LOW: Approximately 20 US gallons of fuel remaining. Plan landing.
Engine Fire During Start
An engine fire during start could be caused by an overloading of fuel in the
combustion chamber and a delayed ignition of the fuel resulting in flame
emanating from the tailpipe. To extinguish the fire, proceed as follows:
Throttle – Full closed.
Starter – Continue to motor the engine.
FUEL VALVE switch – OFF.
IGN ENG circuit breaker – OUT.
Engine Fire During Flight
Throttle – Closed.
Immediately enter autorotation.
FUEL VALVE switch – OFF.
BATtery switch – OFF.
GENerator switch – OFF.
Execute autorotative descent and landing.
NOTE
Do not restart the engine until corrective maintenance has been performed.
Smoke in Cabin
Ventilation of the cabin to protect the pilot and cabin occupants from the effects
of toxic fumes, smoke etc., shall be immediately performed as follows:
Cabin vents – PULL OPEN.
Open cabin windows for maximum ventilation.
Engine Failure and Autorotation
Collective pitch control – Adjust as required to maintain rotor RPM 90 to 107%
NOTE
Rotor RPM maintained at the high end of the operating range will provide
maximum rotor energy to accomplish the landing but will cause an increased
rate of descent.
WARNING
Reduce forward speed to desired autorotative airspeed for existing conditions
(50 to 60 knots indicated airspeed).
At low altitude, close throttle and flare to lose excessive speed.
Apply collective pitch as flare effect decreases to further reduce forward speed
and cushion landing.
It is recommended that level touchdown be made prior to passing through 70%
rotor RPM. Upon ground contact, collective pitch shall be reduced smoothly
while maintaining cyclic in neutral or centered position.
WARNING
Excessive ground run with collective up or any tendency to float for a long
distance prior to ground contact shall be avoided.
Maximum airspeed for steady autorotation is 100 knots IAS. Autorotation above
this speed results in high rates of descent and low rotor speed. A blue radial is
installed on the airspeed indicator as a reminder of this condition.
Engine Air Start
When the cause of the engine failure is believed to be mechanical – do not
attempt a restart.
Collective pitch control – Adjust as required to maintain rotor RPM (90 to
107%)
Reduce forward speed to desired autorotative airspeed (50 to 60 knots IAS) for
existing conditions.
Generator – OFF.
Perform normal engine start procedure.
CAUTION
Do not attempt an air start above 12,000 feet. (TOT rises too fast to control).
Fuel Control and/or Governor Failure
Engine fuel control and/or governor failure will be apparent by a change of
power or RPM. There is no manual fuel control on the engine. Control power
with throttle if engine overspeeds.
Maintain RPM with collective pitch if engine underspeeds.
Establish autorotative glide if power is very low or if engine must be shut down.
Prepare for a power-off landing.
Tail Rotor Control Failure
In the event of a tail rotor failure the failure can be one of two types. Each type
requires its own procedure and shall be performed as follows:
COMPLETE LOSS OF THRUST
Reduce throttle to flight idle. Immediately enter autorotation and maintain a
minimum airspeed of 50 knots during the descent.
NOTE
The large vertical fin may permit controlled flight at low power levels and
sufficient airspeed when a suitable landing site is not available. However the
touchdown shall be accomplished with the throttle in the full closed position.
FIXED PITCH FAILURE (Pitch change slider, control failure etc.)
Depending on the pitch position of the tail rotor at the time of failure; the engine
power and airspeed shall be varied as follows:
Power – Adjust as required to minimize excessive yawing.
Airspeed – Adjust to determine the best velocity to minimize excessive yawing.
Hydraulic System Failure
The first indication of hydraulic boost failure will be an increase in the force
required for control movement. Feedback forces will be noticed as well as rate
limiting. Control motions will result in normal flight reactions in all respects
except for the increased force required for control movement. In the event of
hydraulic power failure, proceed as follows:
Reduce airspeed to 61 to 69 KIAS.
HYD BOOST circuit breaker – OUT. If power is not restored – IN.
CONTROL BOOST or HYDRAULIC SYSTEM switch – ON. OFF if power is
not restored.
Land as soon as practical and investigate.
A run-on landing at 10 to 15 knots is recommended. Maintain airspeed above
translational lift speed for best control at touchdown.
Audio Warning System
ENGINE OUT WARNING SYSTEM
When this system (if functional) is activated, an intermittent audio signal is
produced and the ENG OUT light is ON (N1 less than 55%).
Rotor Low RPM Warning System
When this system is activated the ROTOR LOW RPM light is ON and a steady
audio signal is produced. The low RPM warning system is activated when the
collective pitch is off the down stop and the rotor RPM is less than 90%.
Electrical Power Failure
Electrical power for flight is furnished by the starter which is utilised as a
generator after the start has been accomplished. Evidence of the main generator
failure will be provided by observing the load meter load. There is no provision
for stand-by operation in the event of generator failure. Necessary power can be
supplied by the battery for short periods of time should the generator fail.
Cycle the generator switch – If power is not restored – OFF.
All electrical equipment – ON. Only as needed.
WARNING
Reduce altitude to below 6000 feet pressure altitude if flight permits because of
possible loss of fuel boost pumps.
Engine Icing
ENGINE DEICING or ENGINE ANTI-ICING switch (anti-ice) – ON (if
conditions warrant).
TOT – Maintain within limits.
NOTE
When the anti-ice system is on, the TOT will rise for the same power setting.
END OF BOOK
14
OTHER BOOKS BY JOHN TOAL
Flying The Robinson R44 Helicopter
EASA Air Law For The PPL
Flying The EC120 Colibri
Radiotelephony 101
So You Want To Get An EASA Pilot Licence
Learning Objectives For The EASA Instrument Rating
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