Miscellaneous Instruments............................................................... 1-13-D..... 1-28Outside Air Temperature Gauge .................................................. 1-13-D-1.. 1-28Eight Day Clock or Digital Chronometer ..................................... 1-13-D-2.. 1-28Hourmeter....................................................................................... 1-13-D-3.. 1-29
Electrical System Indicators ............................................................. 1-25-G..... 1-42Electrical System Emergencies and Malfunctions ......................... 1-25-H ..... 1-42
Turbine Outlet Temperature (TOT)............................................... 1-28-C-1.. 1-58Power and Accessory Gearbox........................................................ 1-28-D..... 1-58
Engine Torquemeter ...................................................................... 1-28-D-1.. 1-59Engine Accessories....................................................................... 1-28-D-2.. 1-59Engine Oil System ......................................................................... 1-28-D-3.. 1-60Engine Fuel and Control System ................................................. 1-28-D-4.. 1-60N2 Droop Compensator System ................................................... 1-28-D-5.. 1-62
Retirement Index Number (RIN) ........................................................... 1-29......... 1-63Drivetrain System .................................................................................. 1-30......... 1-63
General Helicopter and Main Components ......................................... 1-1 ........... 1-8Principal Dimensions ............................................................................ 1-2 ........... 1-11Cabin and Baggage Compartment ....................................................... 1-3 ........... 1-14Instrument Panel.................................................................................... 1-4 ........... 1-19Pitot-static System................................................................................. 1-5 ........... 1-22Cyclic and Collective Control Switches............................................... 1-6 ........... 1-35DC Power System .................................................................................. 1-7 ........... 1-38Left Fuel Boost Pump Alternate Circuit Schematic............................ 1-8 ........... 1-41Fuel System............................................................................................ 1-9 ........... 1-44Fuel Burn Sequence .............................................................................. 1-10 ......... 1-46Particle Separator and Vortex Generator ............................................ 1-11 ......... 1-53Power Plant ............................................................................................ 1-12 ......... 1-55Engine Oil System ................................................................................. 1-13 ......... 1-61Drivetrain System .................................................................................. 1-14 ......... 1-64Freewheel Shaft Assembly ................................................................... 1-15 ......... 1-65Transmission Oil System...................................................................... 1-16 ......... 1-67Cyclic Controls....................................................................................... 1-17 ......... 1-74Collective Controls ................................................................................ 1-18 ......... 1-75Elevator Controls ................................................................................... 1-19 ......... 1-76Tail Rotor Controls ................................................................................ 1-20 ......... 1-78Swashplate and Support ....................................................................... 1-21 ......... 1-79Hydraulic System................................................................................... 1-22 ......... 1-81Hydraulic System Schematic................................................................ 1-23 ......... 1-83
23 OCT 2008—Rev. 1———1-5/1-6
MANUFACTURER’S DATA BHT-206L4-MD-1
Section 1SYSTEM DESCRIPTION
11-1. INTRODUCTION
The 206L4 LongRanger helicopter, its primaryand auxiliary systems, as well as emergencyequipment are described in this section. Themain helicopter and its components areshown in Figure 1-1.
The information in this Manufacturer’s Data isalso appl icable to 206L1+ and 206L3+helicopters (Increased Gross Weight Upgradek i t 206-706-530 insta l led perBHT-206-S I -2052) . In format ion that isspecifically applicable to 206L1+ and 206L3+helicopters will be identified as “206L1+ and206L3+ IGW Upgrade”.
1-2. HELICOPTER DESCRIPTION— GENERAL
The LongRanger (Figure 1-1) is a singleengine, seven place helicopter designed totakeoff and land on any reasonably levelterrain. Standard configuration provides forone pilot and six passengers.
The fuselage consists of three main sections:the Forward Section which extends from thecabin nose to the bulkhead aft of thepassenger compartment, the IntermediateSection which extends from the bulkhead afto f the passenger compar tment to thetailboom, and the Tailboom Section.
The helicopter is equipped with a 28 VDCelectrical system. The helicopter electricalpower is provided by a nickel-cadmiumbattery, a starter/generator and an externalpower receptacle.
The basic 206L4 as well as the 206L3+ IGWUpgrade helicopters have a fuel capacity of112.0 U.S. gallons (110.7 gallons usable) [99.4U.S. gallons (98.4 gallons usable) for the206L1+ IGW Upgrade] distributed into threecells. The helicopters are designed to operateon standard aviation jet fuels.
The helicopter is powered by a Rolls-Royceturboshaft engine, model 250-C30P, thatprovides 490 Shaft Horsepower (SHP) fortakeoff and 370 SHP for maximum continuousoperation.
The drivetrain system provides a means oftransmitting power from the engine to themain and tail rotor assemblies. The main rotoris a semi-rigid, seesaw, two blade design thatemploys a preconed and unders lungfeathering axis to ensure smooth operation.The tail rotor is a semi-rigid, delta hinged, twoblade design.
The flight controls are mechanical linkagesthat are actuated by conventional controlsand are used to control flight attitude anddirection. Both the cyclic and the collectivecontro ls incorporate hydraul ic se rvoactuators. The main rotor controls consist ofthe swashplate, drive link and pitch links.
The hydraulic system provides pressurizedfluid to operate the cyclic and collective flightcontrol servo actuators. Operation of thesystem is electrically controlled by means ofthe hydraulic system switch.
The helicopter can be equipped with variouslanding gear configurations including: lowskid landing gear (standard), high skidlanding gear, float landing gear, or emergencyflotation gear.
23 OCT 2008—Rev. 1———1-7
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-1. General Helicopter and Main Components (Sheet 1 of 2)
1
2
13
3
11
9
10
FORWARDSECTION
8
7
5
4
206L4_MD_01_0001
12
28
32
2526
22 192021
27
31
TAILBOOMSECTION
INTERMEDIATE SECTION
15
16
17
18
6
14
23
24
29
30
33
34
35
36
1-8———29 SEP 2006
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-1. General Helicopter and Main Components (Sheet 2 of 2)206L4_MD_01_0026
Hydraulic servo actuators and reservoirMain rotor hub and bladesTail rotor gearbox and servicing doorTail rotor hub and bladesTail rotor driveshaft and coverHorizontal stabilizer and end platesEngine and transmission oil cooler blowerEngine and access doorBaggage compartment doorMain input driveshaftLitter doorCabin ventilation inletAft fairingEngine cowlingAir inlet induction cowlingTransmission cowlingForward cowlingPitot tubeExternal power receptacleBatteryStatic portCrew doorPassenger doorLanding gearFuel cell filler capFuel sump drain buttonFuel shutoff valve panelMain transmissionIntake cowling access doorTailboom attachment access panelEngine oil tank and access panelNavigation lightTailboomTail skidVertical finStrobe (anti-collision) light
29 SEP 2006———1-9
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-3. PRINCIPAL DIMENSIONS
Principle exterior dimensions are shown inFigure 1-2. All height dimensions must beconsidered approximate due to variations inloading and landing gear deflection. Principleinterior dimensions of the cabin and baggagecompartment are shown in Figure 1-3.
1-4. LOCATION REFERENCES
Locations on and within the helicopter can bedetermined in relation to fuselage stations,waterlines and buttock lines, measured fromknown re fe rence po ints . Refe r to theBHT-206L4-MM, Chapter 6 (BHT-206L1-MM/BHT-206L3-MM for 206L1+ and 206L3+ IGWUpgrade) for a lines drawing of the helicopter.
1-4-A. FUSELAGE STATIONS
Fuselage stations (FS or STA) are verticalplanes perpendicular to, and measured alongthe longitudinal axis of the helicopter. Stationzero is the reference datum plane and islocated 55.16 inches (1401 mm) forward of theforward jack point centerline.
1-4-B. WATERLINES
Water l ines (WL) are horizontal p lanesperpendicular to, and measured along thevertical axis of the helicopter. Waterline zerois a plane 20 inches (508 mm) below thelowest point on the fuselage of the helicopter.
1-4-C. BUTTOCK LINES
Buttock l ines (BL) are vert ica l p lanesperpendicular to, and measured to the left andright along the lateral axis of the helicopter.Buttock line zero is a plane at the longitudinalcenterline of the helicopter.
1-5. FUSELAGE — GENERALARRANGEMENT
The fuselage (Figure 1-1) consists of threemain sections: the Forward Section, the
Intermediate Section, and the TailboomSection.
1-5-A. FUSELAGE — FORWARDSECTION
The forward section (Figure 1-1) utilizesaluminum honeycomb structure and providesthe major load carrying elements of theforward cabin. It provides for pilot andpassenger seating, fuel cell enclosure, andpylon support. The cabin area accommodatesup to seven persons or 80 cubic feet (2.2 m3)of cargo. The pilot's station is situated on theright side of the crew compartment. The leftside can accommodate a passenger, or acopilot if the dual controls are installed. Floorloading in the cabin area is limited to 75pounds per square foot (3.7 kg/100 cm2).
The crew and passenger doors contain athree-component wedge shaped window ofblue tinted plastic with the lower forwardsection being used as a sliding window forventilation. This sliding window is adjustableand the window handle functions not only asa handle to open, but also as a retainer tokeep the window from sliding out of the track.The windshields are fabricated of tintedacrylic plastic and are supported by formedaluminum alloy sections. Two blue-tintedtransparent lower windows (chin bubbles)fabricated of plastic are located in the lowercabin nose section. Two cabin roof windows(skylights) are provided in the roof of forwardcompartment. The right side cabin panelwindow is fixed and made of tinted plastic.
The helicopter is equipped with five entrance/exit doors for crew and passengers, as well asa variety of hinged doors and panels, whichprovide access for inspection and servicing.Crew and passenger doors are located onboth sides of the fuselage, while a single litterdoor and a baggage compartment door arelocated on the left side of the helicopter.
1-10———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-2. Principal Dimensions (Sheet 1 of 3)206L4_MD_01_0002_c1
6 FT. 4.2 IN.
(1.93 m)4 FT. 4.0 IN.
(1.32 m)
NO LOAD ON GEAR
7 FT. 5.0 IN.
(2.30 m)AT GROSS WT .OF 4000 LBS.
7 FT. 8.1 IN. (2.34 m)
6 FT. 2.7 IN.
(1.90 m)
42 FT. 8.5 IN.
(13.02 m)
12 FT. 10.2 IN. (3.92 m)
3 FT. 8.0 IN.
(1.12 m)
9 FT. 10.9 IN.
(3.02 m)
GROUND LINE AT 4000 LBS. GW
9 FT. 6.1 IN.
(2.90 m)
2 FT. 10.5 IN.
(0.88 m)
1 FT. 1.0 IN.
(0.33 m)
4 FT. 4.0 IN.
(1.32 m)
2°15'
PRECONE
23 FT. 6.2 IN.
(7.17 m)
10 FT. 0.4 IN.
(3.05 m)
5°
5°
5.5°
3 FT. 2.9 IN. (0.988 m)
CLEARANCE
1 FT. 11.3 IN.
(0.59 m)
DOOR OPENING (APPROXIMATE)
3.33 FT. HEIGHT (1.02 m)
5.0 FT. WIDTH (1.52 m)
6 FT. 0.0 IN.
(1.83 m)
1 FT. 3.0 IN.
(0.38 m)
6 FT. 5.7 IN.
(1.97 m) PITCH CHANGE
AXIS
C
10 FT. 2.4 IN.
(3.11 m)
8° 30' FLAPPING (UP + DOWN)
206L4
(206L1+ AND 206L3+ IGW UPGRADE)
STANDARD LOW GEAR
37 FT. 0.0 IN.
(11.28 m)
11 FT. 8.3 IN. (3.56 m)
PIVOT POINT
TURNING RADIUS
23 FT. 10.0 IN.
(7.26 m)
6 FT. 10.7 IN.
(2.10 m)
Ø5 FT. 5.0 IN.
(1.65 m)
5.25 IN.
(0.13 m)
23 OCT 2008—Rev. 1———1-11
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-2. Principal Dimensions (Sheet 2 of 3)206L4_MD_01_0052_c1
6 FT. 4.2 IN.
(1.93 m)4 FT. 4.0 IN.
(1.32 m)
NO LOAD ON GEAR
7 FT. 5.0 IN.
(2.30 m)AT GROSS WT .OF 4000 LBS.
7 FT. 9.6 IN. (2.38 m)
7 FT. 1.2 IN.
(2.16 m)
42 FT. 8.5 IN.
(13.02 m)
12 FT. 10.2 IN. (3.92 m)
4 FT. 0.0 IN.
(1.21 m)
9 FT. 10.8 IN.
(3.01 m)
GROUND LINE AT 4000 LBS. GW
10 FT. 10.8 IN.
(3.32 m)
3 FT. 10.8 IN.
(1.18 m)
1 FT. 1.0 IN.
(0.33 m)
4 FT. 4.0 IN.
(1.32 m)
2°15'
PRECONE
23 FT. 6.2 IN.
(7.17 m)
10 FT. 0.4 IN.
(3.05 m)
5°
5°
5.5°
3 FT. 2.9 IN. (0.988 m)
CLEARANCE
1 FT. 11.3 IN.
(0.59 m)
DOOR OPENING (APPROXIMATE)
3.33 FT. HEIGHT (1.02 m)
5.0 FT. WIDTH (1.52 m)6 FT. 0.0 IN.
(1.83 m)
2 FT. 1.2 IN.
(0.64 m)
6 FT. 5.7 IN.
(1.97 m) PITCH CHANGE
AXIS
C
11 FT. 0.9 IN.
(3.38 m)
8° 30' FLAPPING (UP + DOWN)
206L4
(206L1+ AND 206L3+ IGW UPGRADE)
HIGH SKID GEAR
37 FT. 0.0 IN.
(11.28 m)
12 FT. 6.8 IN. (3.83 m)
PIVOT POINT
TURNING RADIUS
23 FT. 10.0 IN.
(7.26 m)
6 FT. 10.7 IN.
(2.10 m)
Ø5 FT. 5.0 IN.
(1.65 m)
5.25 IN.
(0.13 m)
1 FT. 6.0 IN.
(0.46 m)
10 FT. 0.0 IN.
(3.05 m)
1-12———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-2. Principal Dimensions (Sheet 3 of 3)206L4_MD_01_0053_c1
6 FT. 4.2 IN.
(1.93 m)4 FT. 4.0 IN.
(1.32 m)
NO LOAD ON GEAR
7 FT. 5.0 IN.
(2.30 m)AT GROSS WT .OF 4000 LBS.
8 FT. 1.2 IN. (2.47 m)
7 FT. 1.2 IN.
(2.16 m)
42 FT. 8.5 IN.
(13.02 m)
12 FT. 10.2 IN. (3.92 m)
2 FT. 6.0 IN.
(0.75 m)
14 FT. 4.8 IN.
(4.39 m)
GROUND LINE AT 4000 LBS. GW
10 FT. 10.8 IN.
(3.32 m)
3 FT. 10.8 IN.
(1.18 m)
1 FT. 1.0 IN.
(0.33 m)
4 FT. 4.0 IN.
(1.32 m)
2°15'
PRECONE
23 FT. 6.2 IN.
(7.17 m)
10 FT. 0.4 IN.
(3.05 m)
5°
5°
5.5°
3 FT. 2.9 IN. (0.988 m)
CLEARANCE
1 FT. 11.3 IN.
(0.59 m)
DOOR OPENING (APPROXIMATE)
3.33 FT. HEIGHT (1.02 m)
5.0 FT. WIDTH (1.52 m)
6 FT. 0.0 IN.
(1.83 m)
2 FT. 1.2 IN.
(0.64 m)
6 FT. 5.7 IN.
(1.97 m) PITCH CHANGE
AXIS
C
11 FT. 0.9 IN.
(3.38 m)
8° 30' FLAPPING (UP + DOWN)
206L4
(206L1+ AND 206L3+ IGW UPGRADE)
EMERGENCY FLOAT KIT
37 FT. 0.0 IN.
(11.28 m)
12 FT. 6.8 IN. (3.83 m)
PIVOT POINT
TURNING RADIUS
23 FT. 10.0 IN.
(7.26 m)
6 FT. 10.7 IN.
(2.10 m)
Ø5 FT. 5.0 IN.
(1.65 m)
5.25 IN.
(0.13 m)
1 FT. 6.0 IN.
(0.46 m)
10 FT. 0.0 IN.
(3.05 m)
1 FT. 7.2 IN.
(0.49 m)
23 OCT 2008—Rev. 1———1-13
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-3. Cabin and Baggage Compartment206L4_MD_01_0003
3.7 FT
(1.1 m)
2.1 FT
(0.6 m)
4.2 FT
(1.3 m)
5.0 FT
(1.5 m)
4.2 FT
(1.3 m)
3.6 FT(1.1 m)
1.8 FT
(0.5 m)
3.1 FT
(0.9 m)
1.3 FT
(0.4 m)
Aft cabin space 80 cubic feet (2.2 m³)Left forward cabin 20 cubic feet (0.6 m ³)
Cabin floor stressed for 75 pounds/square foot (3.7 kg/100 cm ²)
Baggage compartment 16 cubic feet (0.45 m ³)
Baggage compartment floor stressed for 86 pounds/square foot (4.2 kg/100 cm ²)
NOTE
1-14———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
A crew door is installed on each side of theforward fuselage to provide access to thecockpit area. Each door is equipped with alatch assembly, which may be operated fromeither side of the door, and a lock, installed inthe exter io r door hand le . Each doorincorporates a s l id ing w indow and astationary window.
A passenger door is installed on each side ofthe fuselage to provide access to the cabinarea. Each door is equipped with a latchassembly, which may be operated from eitherside of the door, and a lock installed in theexterior door handle. The left passenger dooris hinged on the litter door so that the twomay be opened together to aid in loading thehelicopter.
The litter door, located between the crew andpassenger doors on the left side of thehelicopter, is hinged forward and is opened inconjunction with the left passenger door toprovide increased access for loading thecabin area. The door must be closed andlatched properly, or the LITTER DOOR OPENwarning light will illuminate.
The baggage compartment door, located aft ofthe passenger door on the left side of thehelicopter, is hinged forward and is securedby two pushbutton latches and a keyed lock.
Door locks are installed in the crew, litter,passenger, and baggage compartment doorsto provide security to the cockpit, cabin, andbaggage areas . Locks are s imi lar inconstruction, and are keyed alike.
The battery access door, located on the noseof the helicopter, provides access to theba t te ry, ba t tery re lay, Turb ine Out le tTemperature (TOT) light reset key switch,hourmeter circuit breaker, hourmeter, and leftfuel boost pump circuit breaker. The accessdoor is hinged aft, and two camloc fastenerssecure the forward edge of the door to thefuselage.
Doors and panels are also provided at variouslocations in the cowling and fairings forservicing and inspection. The engine oilreservoir access door and the oil cooleraccess door are located on the aft fairing. Theengine cowling has side panels which arehinged for easy access, and the air inductioncowl ing has doors on bo th s ides forinspection of the transmission area. Mostmiscellaneous access doors open on pianohinges and are secured with flush latches andwinghead stud fasteners.
1-5-B. FUSELAGE — INTERMEDIATESECTION
The intermediate section (Figure 1-1) utilizesan aluminum semi-monocoque constructionand provides a deck for engine installation, abaggage compartment, and a compartmentunder the engine deck for heater andelectrical equipment.
A titanium engine pan, located below theengine, acts as both a drip pan and firewall. Itprovides clearance to remove accessorieswithout having to remove the engine. Theengine pan along with the titanium fore andaft firewalls are fitted within the top part ofthis section to form the engine compartment.
Located directly below the engine area is thebaggage compartment. This compartment hasa volume capacity of 16 cubic feet (0.45 m3)and can carry a maximum of 250 pounds(113.4 kg) not to exceed 86 pounds per squarefoot (4.2 kg/cm2). The baggage door, locatedon the left side of the fuselage and hinged atthe forward end, opens the full width andheight of this compartment and it is securedby means of two pushbutton latches and akeyed lock. Within the compartment are 10tie-down footman loops to secure cargo andequipment.
The equipment compartment is the areabehind the parcel shelf and above thebaggage compartment. It contains the DCpower system electrical relays and voltageregulator and has space for the optional
23 OCT 2008—Rev. 1———1-15
BHT-206L4-MD-1 MANUFACTURER’S DATA
heater or environmental control unit andprovides a convenient location near thesource of bleed air.
1-5-C. FUSELAGE — TAILBOOMSECTION
The tailboom (Figure 1-1) is attached to the aftfuselage by means of four bolts. It consists ofan aluminum alloy monocoque tailboomwhich supports the tail rotor drivetrain as wellas a horizontal stabilizer with end plates,vertical fin, and fairings. Modifications of thetailboom, including installation of antenna,lights or equipment that change the mass ormass distribution of the helicopter, are notpermitted without appropriate Bell Helicopterapproval.
The horizontal stabilizer is constructed ofaluminum honeycomb. The stabilizer is aninverted airfoil that provides a downwardresultant lift on the tailboom to maintain thecabin in a nearly level attitude throughout thecruise airspeed range and to aerodynamicallystreamline the fuselage to reduce drag.Leading edge slats are installed to improvepitch stability during climbs. A controllabletrailing edge synchronized elevator modifiesthe downforce transmitted to the tailboombased on airflow across the unit.
The leading edges of the end plates ordynamic dihedrals, located on each end of thehorizontal stabilizer, are offset 5° to the left ofcenterline (Figure 1-2). The offset improvesdihedral (roll) stability of the helicopter inforward flight.
The vertical fin, constructed primarily ofaluminum and honeycomb material, providesdirectional (yaw) stability. The leading edge ofthe fin is canted 5.5° outboard to reduce tailrotor thrust requirements during forwardflight (Figure 1-2). It contains a top fairing tomount the anticollision light and a rubberbumper and tail skid to protect the tail rotorand warn the pilot of a tail-low landingattitude.
A one-piece cowling is used to cover the tailrotor driveshafts and is secured with quickrelease fasteners.
The tail rotor gearbox fairing encloses the tailrotor gearbox and is attached to the tailboomand vertical fin. It incorporates a whiteposition light and three screens. The screensprovide cooling airflow for the tail rotorgearbox. The oil level in the gearbox is visiblethrough one of the screened openings.
1-6. CABIN HARD POINTS
Two airframe hard points are located on thebelly of the helicopter forward of the frontcrosstube and one hard point is located onthe belly aft of the baggage compartment.These hard points are used for jacking andmooring. Clevises may be attached to thesehard points for securing helicopter tie-downs.
The hard points are not to be used for towing.Refer to Section 2 for additional information.
1-7. COWLINGS
The forward, transmission, air inlet induction,engine, and aft fairing cowlings (Figure 1-1)are constructed of aluminum alloy, fiberglass,or honeycomb materials. Each section isreadi ly removable for engine andtransmission maintenance. Access doors areprovided with flush-type latches that permitinspections and servicing without removingthe cowling.
The forward cowling incorporates a singlepoint hinge fitting assembly that is located atthe forward end of the fairing. Two supportrods are stowed inside to support the fairingin a raised position. Two toggle hook latchfittings secure the cowling in the closedposition.
The transmission cowling encloses theforward portion of the main transmission andcontrol l inkage from the cockpit to theswashplate.
1-16———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
The air inlet induction cowling encloses theaft portion of the main transmission. Inletducts are provided on each side of the airinduction cowling to direct the airflow into theinduction screen or particle separator. Twodoors are also provided for maintenance andinspection of the induction area and maintransmission. The induction screen is locatedinside and attached to the air induction cowland the forward side of the forward enginefirewall.
The engine cowling, which includes theengine upper cowling and engine sidecowling doors, is constructed of aluminumalloy. It contains two screen vents forward ofthe cutout for the exhaust stack. Hinged to theengine upper cowling are two engine sidedoors. The side doors are constructed ofaluminum alloy and incorporate five screenvents, three flush latches, and two camlocfasteners. The right engine side door includesa screen vent for the starter/generator inlet aircooling duct. Each engine side door may behinged open and supported by a rod foraccess into the engine compartment.
The aft cowling is constructed of aluminumalloy and incorporates two screen vents, oilcooler and oil tank access doors, and a cutoutfor the oil cooler blower exhaust.
1-8. LANDING GEAR
The skid landing gear (Figure 1-1) consists oftwo skid tubes attached on the ends of twoarched crosstubes that are secured to thefuselage by means of four strap assemblies.Each skid tube is designed with a forward endstep, a tow fitting, two saddle fittings forcrosstube attachment, six separate shoesalong the bottom, a rear cap, and two eyeboltfittings for mounting of ground handlingwheels. The airfoil type skid landing gearfairings (optional) enclose the forward and aftcrosstubes. These fairings are constructedfrom white thermoplastic with aluminum alloystiffeners and supports.
When the opt iona l h igh sk id gear oremergency lightweight flotation gear is
installed, the front mounting hardware mustbe modified to include a spacer bar, springs,and two lower supports per s ide. Thisadditional attachment hardware will preventmedium frequency vibrations from beingtransmitted from the landing gear to theairframe.
It is recommended that no components beattached to the landing gear assembly exceptas des ignated by the manufac turer.Unauthorized attachments may lead to failureof the crosstubes.
1-9. GROUND HANDLING
Ground handling wheels may be attached tothe two lugs on the skid tubes. Care shouldbe exercised to make sure that groundhandling wheels are properly maintained andsecurely attached to the skids, prior to liftingthe helicopter.
Tow rings are located on the inside of eachskid. Never attach a towing device to anyother locat ion . Refer to Sect ion 2 foradditional information.
1-10. CREW AND PASSENGERSEATING
All seat back cushions are constructed of softsponge with styrene backing. Seat cushionsare constructed of soft urethane foam. Backand seat cushions are bonded to tubularframes and upholstered with fabric resistantto flame, fluids, and climatic conditions. Asmall storage area is located behind eachcrew seat. Aft facing seats are equipped withan adjustable headrest.
1-10-A. SEAT RESTRAINT ASSEMBLIES
Each crew and passenger seat is equippedwith a restraint assembly that consists of aninert ia reel , shoulder harness, and anadjustable seat belt. The inertia reel isprovided with an anti-rebound lock featureand is capable of retracting 22 inches(558.8 mm) of web belt.
23 OCT 2008—Rev. 1———1-17
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-11. INSTRUMENT PANEL ANDCONSOLES
The instrument panel (Figure 1-4) is mountedon a center pedestal forward of the two crewseats and the top is tilted forward at a 5° anglefor maximum visibility. Flight instruments arelocated on the right side of the panel andsystem instruments are in two vertical rows tothe left of the flight instruments. Caution andWarning panel lights are mounted just belowthe glare sh ie ld across the top of theinstrument panel.
The pedestal extends aft from the instrumentpanel between the seats to form a console forthe radios and miscellaneous controls.
The overhead console is centrally positionedaft of the windshields on the cockpit ceiling.This console contains most of the circuitbreakers and electrical switches.
1-12. PITOT/STATIC SYSTEM
The pitot tube is mounted on the mostforward part of the cabin nose structure(Figure 1-5). The tube supplies impact air tothe airspeed indicator. Turning the PITOTHEAT switch to the ON position activates theheater element for the pitot tube.
Static air pressure for the airspeed, altimeterand vertical speed indicator operation isobtained from two non-heated static ports onthe left and right sides of the helicopter, aft ofthe cabin lower chin bubble windows.
1-13. INSTRUMENT SYSTEM
The instrument system (Figure 1-4) is dividedinto four separa te categor ies : f l ight ,navigation, propulsion, and miscellaneous.All indicators are installed in the hingedinstrument panel except the OAT gauge,compass, and hourmeter. The OAT gauge is
mounted in the upper left corner of the pilot'swindshield. The magnet ic compass ismounted to the r ight side of the cabinstructure slightly forward of the instrumentpanel. The hourmeter is mounted in thebattery compartment.
1-13-A. FLIGHT INSTRUMENTS
The flight instruments include the airspeedindicator, altimeter, turn and slip indicator(optional), vertical speed indicator (optional),attitude indicator (optional) or inclinometerand directional gyro (optional).
1-13-A-1. AIRSPEED INDICATOR
The airspeed indicator is a standard pitot/static instrument. This indicator provides anairspeed reading in knots (KTS) and miles perhour (MPH) by measuring the differencebetween impact air pressure from the pitottube and the static air pressure from the staticports.
The indicator presents airspeed from 0 to 150knots and is scaled in 5-knot increments at 20knots and above. A maximum autorotationspeed blue line is located at 100 knots and amaximum speed (VNE) red line is located at130 knots.
206L4_MD_01_0007
1-18———29 SEP 2006
23 OCT 2008 Rev. 1 1-19
TURER’S DATA BHT-206L4-MD-1
Figure 1-4. Instrument Panel (Sheet 1 of 3)206L4_MD_01_0043_c1
58
10 9
11
CAUTION
WRN HORN MUTE
LT TEST
12
7
6
3
TRANSOIL TEMP
ROTORLOW RPM
PTRANS
OIL PRESSBATTERY
HOTENG OUT
BATTERYRLY
PULL
FOR
QUICK
ERECT
ON
FUEL
VALVE
OFF
2 MIN TURN
MANUFAC
AVOID CONT OPS 71.8% TO 91.5% N2
ENG
OIL
X 10
P T
15 15
1010
5 5
0
PSI °C
OIL
X 10
P T
15
10
5 5
0
PSI °C
XMSN15
10
30
80
90
100
70
20
10
50
40
60GAS
PRODUCER
%RPM
1 09
8
6
2
3
45
7
TORQUE
PERCENT
010
20
30
4050
60 7080
90
100
110120
TEMP°CX100
TURBOUT
12
34 5
6
7
89 ROTOR
POWER
TURBINE
PERCENTRPM
120
110
90
807060
50
40
30
20
100
100
T
R
AIRSPEED
MPHX 10
150 200
120
100
80
60
40174
6
810
12
14
KNOTS
FUEL
QTYLBS
x1000
1
2
34 5
6
7
8
0
1
2
310
LOAD FUEL
% DC PSI P
X 10
5
00
20
17
16 14
1
23
21
TOTAL
FWD
TOT
LT TEST
(TYPICAL)
FUEL QTY
22
2 34
115
OBS
1819
GENFAIL
R/FUELPUMP
L/FUELPUMP
BAGGAGEDOOR
TRANSCHIP
FUELFILTER
ENG CHISPARESPARESPARE
SPARE SPARE SPARE SPARE SPARE T/R CHIP
FUELLOW
LITTERDOOR OPEN
THIS HELICOPTER MUST BE OPERATED IN
COMPLIANCE WITH THE OPERATING
LIMITATIONS SPECIFIED IN THE APPROVED
HELICOPTER FLIGHT MANUAL.
MINIMUM COCKPIT WEIGHT 170 LBS.
SELECTIVE PASSENGER LOADING
WHEN BOTH CREW SEATS ARE OCCUPIED
ONLY ONE (1) MID PASSENGER IS
PERMITTED UNLESS THERE ARE TWO (2)
AFT PASSENGERS.
WHEN ONLY ONE (1) CREW SEAT IS
OCCUPIED, NO MORE THAN TWO (2) AFT
PASSENGERS ARE PERMITTED UNLESS
THERE IS ONE (1) MID PASSENGER.
ABOVE 4,150 LB GW, ALTERNATE
PASSENGER LOADING FROM SIDE TO SIDE.
REFER TO RFM WEIGHT AND BALANCE
FOR ADDITIONAL LOADING INFORMATION.
1-20
BHT-2
25
Rev. 1 23 OCT 2008
06L4-MD-1 MANUFACTURER’S DATA
Figure 1-4. Instrument Panel (Sheet 2 of 3)206L4_MD_01_0044_c1
29
30
3135
24
36
37
38
32
OVERHEAD PANEL(TYPICAL)
MISCELLANEOUS CONTROL PANEL(TYPICAL)
34
26 27 28
33
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-4. Instrument Panel (Sheet 3 of 3)
Engine oil temperature/pressure gauge Torquemeter indicator Airspeed indicator RPM warning horn mute button Attitude indicator (optional)/inclinometer (not shown) Caution and warning panel Caution and warning panel light test button Altimeter indicator Vertical speed indicator (optional) Directional gyro/horizontal situation indicator (optional) Turn and slip indicator (optional) Fuel valve switch and guard Radio magnetic indicator (RMI) (optional) Dual tachometer indicator VOR indicator (optional) Turbine outlet temperature gauge Gas producer indicator ClockFuel pressure/loadmeter gauge Fuel quantity gauge Forward fuel cells quantity switch Turbine outlet temperature gauge light test button Transmission oil temperature/pressure gauge Miscellaneous controls panel Particle separator purge switch (optional) Caution and warning panel Bright/Dim switch Engine anti-icing switch Hydraulics switch Overhead panel Anti-collision light switch Generator switch Battery switch Defog blower switch Pitot heat switch Attitude indicator switch (optional) Heater/air conditioning switch (optional) Position light switch Instrument light intensity rheostat
The barometric pressure altimeter presentsan altitude reading in feet above mean sealevel (MSL) based on the relationship betweenthe static air pressure and the barometricsetting on the altimeter. The barometricsetting may be adjusted to reflect the currentbarometric pressure corrected to sea level ininches of mercury or in millibars, dependingon the instrument installed.
1-13-A-3. TURN AND SLIP INDICATOR
The turn and slip indicator (optional) turnneedle is powered by the 28 VDC bus throughthe TURN IND circuit breaker. The turn needleindicates the rate at which the helicopter isturning about the vertical axis in degrees persecond. The ba l l ind icates when thehelicopter is in a coordinated turn or balanced
straight and level flight, i.e. "in trim". If thehelicopter is in a slip or skid, the ball willmove off center.
1-13-A-4. VERTICAL SPEED INDICATOR
The vertical speed indicator (optional) utilizesstatic air pressure to present rate of climb ordescent information from 0 to 4000 feet perminute.
1-13-A-5. ATTITUDE INDICATOR
The attitude indicator (optional) is powered bythe 28 VDC bus and presents pitch and rollattitudes of the helicopter in relation to thehorizon. A fail flag will be visible when powerto the indicator is lost, warning that attitudeinformation is unreliable.
206L4_MD_01_0008
206L4_MD_01_0010
206L4_MD_01_0015
206L4_MD_01_0056
29 SEP 2006———1-23
BHT-206L4-MD-1 MANUFACTURER’S DATA
The manual caging knob should be pulled outto erect the gyro when power is applied. Thepi tch t r im knob is used to ad just thehelicopter symbol vertically.
1-13-A-6. INCLINOMETER
The inclinometer consists of a curved glasstube, ball, and dampening fluid. The ballind icates when the he l icopter is in acoordinated turn or balanced straight andlevel flight, i.e. "in trim". If the helicopter is ina slip or skid, the ball will move off center.
1-13-A-7. DIRECTIONAL GYRO
The directional gyro (optional) is powered bythe 28 VDC bus. It is not slaved to a magnetic
flux valve and therefore, must be set prior totakeoff and periodically during flight tocorrect for magnetic variation and gyroscopicprecession. A fail flag will be visible whenpower to the indicator is lost , warningdirectional information is unreliable.
1-13-B. NAVIGATION INSTRUMENTS
The navigation instrument on the basichelicopter is the magnetic compass.
All other navigation instruments are not partof the basic helicopter. See the avionicsmanufacturers documentation for applicableinformation.
1-13-B-1. MAGNETIC COMPASS
The magnet ic compass is a standard,non-stabilized, magnetic-type instrumentmounted on a support attached to the pilotside of the forward cabin next to the doorpost. A compass correction card is locatedbelow the instrument.
1-13-C. PROPULSION INSTRUMENTS
The propulsion instruments consist of thedual tachometer, gas producer tachometer,torquemeter ind ica tor, tu rb ine out le ttemperature indicator, engine oil pressure/temperature indicator, transmission oilpressure/temperature indicator, fuel pressure/DC loadmeter indicator and fuel quantityindicator.
206L4_MD_01_0050
206L4_MD_01_0049
206L4_MD_01_0009
1-24———29 SEP 2006
MANUFACTURER’S DATA BHT-206L4-MD-1
1-13-C-1. DUAL TACHOMETER TURBINE/ROTOR (N2/NR)
The dual tachometer is equipped with twopointers, one displaying main rotor RPM onthe inner scale and one that displays enginepower turbine RPM on the outer scale. Allscales are presented in percent of rated RPM.
NOTE
The rotor (NR), and turbine (N2),tachometer genera tors areself-generating and will operatewithout electrical bus power.
The T (engine) indication is powered by thepower turbine (N2) tachometer generator thatis mounted on the left front side of the engineaccessory gearcase.
The R (rotor) indication is powered by therotor tachometer generator mounted onto thet ransmiss ion o i l pump pad on thetransmission lower case.
1-13-C-2. GAS PRODUCER INDICATOR (N1)
NOTE
An alternate indicator to that shownmay be instal led. Refer to theBHT-206L4-FM-1, Section 1.
NOTE
The gas producer (N1) tachometergenerator is self-generating and willoperate without electrical bus power.
The gas producer indicator (N1) displaysengine gas producer speed in percent of ratedRPM. The outer scale of the gauge displays in2% increments and the smaller dial scaledisplays in 1% increments. The indication ispowered by the gas producer tachometergenerator that is mounted onto the right frontof the engine accessory gearcase.
206L4_MD_01_0011206L4_MD_01_0012
23 OCT 2008—Rev. 1———1-25
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-13-C-3. TORQUEMETER INDICATOR
NOTE
The engine torquemeter indicatoroperates without electrical buspower.
The engine torquemeter is a wet oil gauge.The system utilizes engine oil that has beenmetered by the torquemeter gear shaftassembly within the engine to indicate poweroutput of the engine. The gauge displays inpercent torque.
1-13-C-4. TURBINE OUTLETTEMPERATURE INDICATOR
NOTE
Alternate indicators to that shownmay be installed on 206L1+ and206L3+ IGW Upgrade helicopters.Refer to the BHT-206L4-FM-1 ,Section 1.
Turbine Outlet Temperature (TOT) indicatordisplays engine turbine gas temperature indegrees Celsius. Four probes measure gastemperature between the gas producerturbine and the power turbine. The TOT gaugeis powered by the 28 VDC bus through the3-amp TOT IND circuit breaker. A TOT LightTest button is located on the left side of theinstrument panel.
The pressure side of the gaugeoperates without electrical buspower.
The engine oil pressure gauge is a wet oilgauge. The engine oil pressure/temperatureindicator displays oil pressure in PSI on theleft side of the instrument and oil temperaturein degrees Celsius on the right side.
The engine oil temperature input signal isprovided by a thermo-bulb installed on the
206L4_MD_01_0013
206L4_MD_01_0018
206L4_MD_01_0017
1-26———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
engine oil tank. The temperature side of thegauge is powered by the 28 VDC bus throughthe 3-amp ENG TEMP IND circuit breaker inthe overhead panel.
The pressure side of the gaugeoperates without electrical buspower.
The transmission oil pressure/temperatureindicator displays oil pressure in PSI on theleft side of the instrument and oil temperaturein degrees Celsius on the right side. Thetransmission oil pressure gauge is a wet oilgauge.
The transmission oil temperature input signalis provided by a thermo-bulb installed on thet ransmiss ion o i l f i l te r assembly. Thetemperature side of the gauge is powered bythe 28 VDC bus through the 3-amp XMSNTEMP IND circuit breaker in the overheadpanel.
1-13-C-7. FUEL PRESSURE/DCLOADMETER
NOTE
Alternate indicator to that shownmay be installed on 206L1+ and206L3+ IGW Upgrade helicopters.Refer to the BHT-206L4-FM-1 ,Section 1.
The fuel pressure/DC loadmeter displays fuelpressure in PSI on the right side of theinstrument and the DC current load of thegenerator on the left side of the instrument.
Fue l p ressure is der ived f rom apressure-operated transducer that measuresthe output pressure of the boost pumps. Thetransducer is located just below the fuelshutoff valve. The pressure side of the gaugeis powered by the 28 VDC bus through the3-amp QTY PRESS circuit breaker in theoverhead panel.
The loadmeter ind icates the load inpercentage that is supplied to the 28 VDC busfrom the generator. The loadmeter signal isprovided from the loadmeter shunt located inthe electrical equipment shelf. The circuit iscompleted through two circuit breakers inc lose prox imity to the shunt . The DCloadmeter indicates the load in amperes thatis being supplied to the 28 VDC bus by the
206L4_MD_01_0019
206L4_MD_01_0021
23 OCT 2008—Rev. 1———1-27
BHT-206L4-MD-1 MANUFACTURER’S DATA
generator. The maximum continuous load is90% (180 amps).
1-13-C-8. FUEL QUANTITY INDICATOR
NOTE
Alternate indicator to that shownmay be installed on 206L1+ IGWUpgrade helicopters. Refer to theBHT-206L4-FM-1, Section 1.
The fuel quantity indicator displays total fuelquantity in pounds. The gauge will indicate752.8 pounds (669.1 POUNDS, 206L1+ IGWUpgrade) (JP-5 at 15°C) with a full fuel load.The signal is generated by the fuel quantitytransmitter that is connected to three fuelprobes located in the fuel system. The fuelquantity indicator is powered by the 28 VDCbus through the 3-amp QTY PRESS circuitbreaker on the overhead panel.
The FUEL QTY switch has two positions (FWDand TOTAL). The switch is spring loaded tothe TOTAL (center) position. Holding theswitch in the FWD position causes the fuelgauge to indicate fuel quantity in the forwardcells only. Allowing the switch to return to the
TOTAL position allows the gauge to displaythe fuel quantity of the entire system.
1-13-D. MISCELLANEOUSINSTRUMENTS
The miscellaneous instruments are the OATgauge, the clock, and the hourmeter.
1-13-D-1. OUTSIDE AIR TEMPERATUREGAUGE
The Outside Air Temperature (OAT) gauge ismounted in the upper left corner of the pilot’swindshield. It provides OAT information inboth degrees Celsius and Fahrenheit.
1-13-D-2. EIGHT DAY CLOCK OR DIGITALCHRONOMETER
206L4_MD_01_0020
206L4_MD_01_0055
206L4_MD_01_0014
1-28———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
The eight day clock or digital chronometer(shown), is mounted in the lower left area ofthe instrument panel.
Features of the digital clock include UniversalCoordinated Time (GMT), Local Time (LT) in24-hour format, or Elapsed Time. ElapsedTime feature includes a count up timer to amaximum of 99 hours, 59 minutes.
The eight day clock provides information inhours, minutes, and seconds and needs to bewould mechanically.
1-13-D-3. HOURMETER
The hourmeter is mounted on the aft bulkheadof the battery compartment. The hourmeter ispowered by the 28 VDC bus through a 1-ampremote circuit breaker that draws power fromthe 5-amp CAUTION circuit breaker. Thesystem is activated when the gas producerRPM (N1) is greater than 55%.
1-14. CAUTION AND WARNINGSYSTEM
The caution and warning system includes thecaution and warning light panel (Figure 1-4),engine RPM sensor and warning horn, lowrotor RPM sensor and warn ing horn ,associated components, and interconnectingwiring.
The caution and warning l ight panel ispowered by the 28 VDC bus and protected bythe 5-amp CAUTION circuit breaker. Each
segmented indicator light is in series with itsrespective system. The warning lights arecolored RED while caution lights are coloredAMBER. Pressing the CAUTION LT TESTbutton will illuminate all the lights.
1-14-A. ENGINE OUT
The ENGINE OUT warning system providessimultaneous visual and audible indicationsof an engine out condition. This includes anENGINE OUT warning light, located on thecaution and warning light panel, and theENGINE OUT warning horn (intermittentsound), located on the overhead trim panel.
When gas producer RPM (N1) drops below 55±3%, the engine RPM sensor c loses ,activating the warning light and audio.
1-14-B. ROTOR LOW RPM
The rotor low RPM warning system providessimultaneous visual and audible indicationsof a low rotor RPM condition. The rotor lowRPM light and audio (continuous tone) areactivated when the NR sensor detects therotor RPM dece lera t ing be low 90 ±3%regardless of collective position.
1-14-C. WARNING AUDIO MUTE SYSTEM
The warning audio mute system is installedon all 206L4 helicopters as well as 206L3+IGW Upgrade helicopters S/N 51190 andsubsequent or S/N 51001 through 51189 PostTB 206L-87-137 and 206L1+ IGW Upgradehelicopters S/N 45154 and subsequent PostTB 206L-87-137. A button is located in theupper center area of the instrument panel tosilence the audio for ENG OUT and ROTORLOW RPM. Pressing the HORN MUTE button
206L4_MD_01_0022
23 OCT 2008—Rev. 1———1-29
BHT-206L4-MD-1 MANUFACTURER’S DATA
will temporarily silence both warning horns.The warning horns are automatically reset tooperate when preset operational parametersare re-established.
1-14-D. BATTERY HOT
A temperature sensor installed within thebattery case activates the BATTERY HOTwarning light when preset temperature limits(approximately 140°F/60°C) are reached.
1-14-E. BATTERY RELAY
The BATTERY RLY caution light illuminateswhen the battery switch is placed in the OFFposition and the battery relay remains closed(battery still connected to the 28 VDC bus).
1-14-F. GENERATOR FAILURE
The GEN FAIL caution light illuminates whenthe generator is not online or has failed.
1-14-G. TRANSMISSION OIL PRESSURE
The transmission oil pressure warningsystem includes the TRANS OIL PRESScaution light which is activated by an oilpressure switch that is mounted in the lowerpedestal area. The light will illuminate whenpressure decreases to 30 PS I ±2 andextinguishes when pressures increases to36 PSI.
1-14-H. TRANSMISSION OILTEMPERATURE
A thermal swi tch is located on thetransmission oil filter assembly. If the oiltemperature rises above 110°C ±5.5°C, theTRANS OIL TEMP caution light will illuminate.
1-14-I. ENGINE CHIP DETECTOR
The engine chip detector caution systemincludes the ENG CHIP caution light whichmay be activated from either of two enginegearbox mounted magnetic chip detectors. Ifferrous metal particles are found in the engineoil, the magnets in the detectors will attractthem, completing the circuit and illuminatingthe ENG CHIP caution light.
1-14-J. TRANSMISSION CHIP DETECTOR
The transmission chip detector cautionsystem includes the TRANS CHIP cautionlight, two transmission chip detectors, and afreewheeling unit chip detector. If ferrousmetal particles are found in the transmissionor freewheel oil, the magnets in the detectorswill attract them, completing the circuit andilluminating the TRANS CHIP caution light.
1-14-K. TAIL ROTOR GEARBOX CHIPDETECTOR
1-30———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
The tail rotor gearbox chip detector cautionsystem includes the T/R CHIP caution lightand the tail rotor gearbox chip detector. Ifferrous metal particles are found in thegearbox oil, the magnet in the detector willattract them, completing the circuit andilluminating the T/R CHIP caution light.
1-14-L. FUEL LOW
The FUEL LOW caution light illuminates whenapproximately 50 to 75 pounds (10 gallons) offuel remains in the aft (main) cell. A fuel lowsensor located in the aft fuel cell, independentof the fuel quantity indicating system,controls the FUEL LOW caution light.
1-14-M. FUEL BOOST PUMPS
The L/FUEL PUMP and R/FUEL PUMP cautionlights are controlled by their respective fuelflow switch located near the ejector pump,between the forward fuel cells. The flowswitches are activated when fuel flow isinsufficient to operate the ejector pump.
If both fuel pump lights are illuminated,unusable fuel may be as high as 86.3 pounds(39.2 kg) with FUEL QTY switch installed or156.9 pounds (70.9 kg) on 206L3+ IGWUpgrade helicopters without FUEL QTYswitch (154.4 pounds (70.1 kg) for 206L1+ IGWUpgrade) due to inability to transfer fuel.
1-14-N. FUEL FILTER
The FUEL FILTER caution light illuminateswhen the a i r f rame fue l f i l te r is in animpending bypass condition (approximately 1PSI differential). The airframe fuel filter willbypass at approximately 4 PSI differential.
1-14-O. BAGGAGE DOOR
A micro-switch located in the upper rightcorner of the baggage compartment openingilluminates the BAGGAGE DOOR caution lightwhen the door is opened or is not securelyfastened.
1-14-P. LITTER DOOR
The litter door includes the LITTER DOOROPEN warning light, micro-switches for theupper and lower door strikers, and relatedwiring. If the litter door is not securelyfastened or if it has been removed, thecaution light illuminates.
1-14-Q. SPARE CAUTION LIGHT
The spare caut ion l ights are normallyilluminated only during the testing of thepress to test feature for the caution andwarning system.
23 OCT 2008—Rev. 1———1-31
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-15. FUEL VALVE SWITCH
A fuel valve switch (Figure 1-4) is located onthe lower right side of the instrument paneland electrically operates a motor-driven fuelvalve providing a means of shutting off fuelflow to the engine. The FUEL VALVE switch iscovered by a red locking device to preventinadvertent closing of the fuel valve.
1-16. MISCELLANEOUS CONTROLPANEL
The miscellaneous control panel (Figure 1-4)contains switchology that allows the pilot toactivate certain helicopter systems. Switcheson this panel include the Particle SeparatorPurge switch (optional), the Caution andWarning Panel Bright/Dim switch, the EngineAnti-icing switch, and the Hydraulic Systemswitch.
1-16-A. PARTICLE SEPARATOR PURGESWITCH
When the Particle Separator Purge switch(optional) (Figure 1-4) is selected to the ONposition, engine bleed air is used to purgedebris from the particle separator.
1-16-B. CAUTION AND WARNING PANELBRIGHT/DIM SWITCH
When the INST LT rheostat is ON (Figure 1-4),positioning the BRT/DIM CAUTION LIGHTSswitch momentari ly to the BRT or DIMposition will dimly or brightly illuminate thecaution panel lights. The caution lights willremain at the selected intensity until the INSTLT rheostat is turned OFF or another positionis selected.
1-16-C. ENGINE ANTI-ICING SWITCH
The engine anti-ice solenoid valve is poweredfrom the 28 VDC bus through the 5-ampENGINE ANTI-ICING circuit breaker and theENGINE ANTI-ICING switch (Figure 1-4). Whenthe solenoid valve is de-energized (ON), hotair passes from the compressor diffuser
through the anti-ice valve to the engine inlethousing. This hot air aids in preventing iceformation on the hollow inlet guide vanes.Power is only provided to the solenoid whenthe switch is positioned to OFF. In the event oftotal electrical failure, the system is fail-safeON (no electrical power = engine anti-icingON).
1-16-D. HYDRAULIC SYSTEM SWITCH
The HYDRAULIC SYSTEM switch (Figure 1-4)controls the operation of the hydraulic bypasssolenoid. The solenoid is normally open whende-energized. The switch provides power tothe solenoid in the OFF position only. In theevent of total electrical failure, the system isfail-safe ON (no electrical power = hydraulicsON).
1-17. OVERHEAD CONSOLE
The overhead console (Figure 1-4) is centrallypositioned aft of the windshields on thecockpit ceiling. This console contains most ofthe helicopter electrical system circuitbreakers and electrical switches and is easilyaccessible from the pilot's station. The circuitbreakers are installed to protect a specificsystem from excessive amperage. Circuitbreakers may be either the pop-out or switchstyle. The electr ic toggle switches aredesigned to connect/disconnect a specificsystem to the 28 VDC bus.
1-17-A. GENERATOR SWITCH
The GEN swi tch (F igure 1-4 ) is athree-position switch that controls a normallyopen line control relay (Figure 1-7) thatconnects the generator to the 28 VDC bus.Switch positions are OFF, GEN (ON), andRESET. When the GEN switch is in the OFFposi t ion , the l ine cont ro l re lay isde-energized, interrupting the circuit from thegenerator to the 28 VDC bus. If electricalpower is applied to the bus through thebattery or external power with the GEN switchin the OFF position, the GEN FAIL cautionlight will be illuminated on the caution and
1-32———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
warning panel. When the switch is selected tothe GEN (ON) position with the electricalsystem operating normally, the line controlrelay is energized closed, completing thecircuit and connecting the generator to the28 VDC bus and extinguishing the GEN FAILcaution l ight. The RESET posit ion is amomentary spring-loaded position that allowsthe opera tor the a t tempt to reset thegenerator should a generator malfunctionoccur. Th is RESET fea ture is used inconjunction with the 10-amp GEN FIELD and5-amp GEN RESET circuit breakers locatedimmediately aft of the GEN switch on theoverhead panel.
1-17-B. BATTERY SWITCH
The BAT switch (Figure 1-4) is a two-positionswitch that controls a normally open batteryrelay (Figure 1-7) in the nose of the helicopterthat connects the battery to the 28 VDC bus.Switch positions are OFF and BAT. When theswitch is in the OFF position, the battery relayis de-energized, interrupting the circuit fromthe battery to the 28 VDC bus. When theswitch is selected to the BAT (ON) position,the bat tery re lay is energized c losed,completing the circuit and connecting thebattery to the 28 VDC bus.
1-17-C. DEFOG BLOWER SWITCH
The DEFOG BLOWER switch (Figure 1-4) is atwo-position, 5-amp circuit breaker styleswitch that connects the two electric blowermotors to the 28 VDC bus. When the switch ispositioned to the aft position (OFF), the circuitis de-energized and the two blower motorsare OFF. When the switch is positionedforward, the circuit is completed, connectingthe two blower motors to the bus, blowingambient air on the windscreens to assist infog removal.
1-17-D. PITOT HEAT SWITCH
The PITOT HEAT switch (Figure 1-4) is atwo-position, 5-amp circuit breaker styleswitch that connects the electric heating
element in the pitot tube to the 28 VDC bus.When the switch is positioned to the aftposition (OFF), the circuit is de-energized andthe pitot heat element is OFF. When the switchis pos i t ioned forward , the c i rcu i t iscompleted, connecting the pitot heat elementto the 28 VDC bus, heating the element toprevent ice formation and removing moistureinside the pitot tube. The static ports are notheated.
1-17-E. POSITION LIGHT SWITCH
The POS LT swi tch (F igure 1-4 ) is atwo-position, 5-amp circuit breaker styleswitch that connects the five helicopterposition lights to the bus. When the switch ispositioned to the aft position (OFF), the circuitis de-energized and the position lights areOFF. When the switch is positioned forward(ON), the circuit is completed, connecting theposition lights to the 28 VDC bus. Positionlights are located on the outboard side of theauxiliary fins on the horizontal stabilizer, thelower forward fuselage left and right side, andon the tail rotor gearbox fairing.
1-17-F. ANTI-COLLISION LIGHT SWITCH
The ANTI-COLL LT switch (Figure 1-4) is atwo-position, 5-amp circuit breaker styleswitch that connects the anti-collision light ontop of the vertical fin to the bus. When theswitch is positioned to the aft position (OFF),the c i rcu i t is de-energ ized and theanti-collision light is OFF. When the switch isposit ioned forward (ON), the circuit iscompleted, connecting the anti-collision lightto the bus.
1-17-G. INSTRUMENT LIGHT RHEOSTAT
The INST LT rheostat (Figure 1-4) controls thebrightness of the background lighting on theinstrument panel and l it port ion of theoverhead console. This rheostat is also usedalong with the caution and warning panelBRT/DIM switch on the miscellaneous controlpanel to control the brightness of the cautionand warning lights.
23 OCT 2008—Rev. 1———1-33
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-18. LIGHTING SYSTEMS
The lighting system includes both interior andexterior lighting. The interior lighting systemincludes the cockpit map reading light,instrument and control panel lighting, andcabin lighting. The exterior lighting systemincludes anti-collision, position, and landinglights.
1-18-A. INTERIOR LIGHTING SYSTEM
The cockpit map reading light features anarrow spotlight or wide floodlight beam. It isa multipurpose utility light designed toprovide blue or white illumination and isprotected by its own circuit breaker. Controlsfor ON/OFF, BRIGHT/DIM, BLUE/WHITE, andSPOT/FLOOD are incorporated into thecockpit light body. The 28 VDC power issupplied through the COCKPIT LIGHTS 5-ampcircuit breaker located on the overheadconsole.
Instrumentation lighting 28 VDC power issupplied through the INSTR LIGHTS 5-ampcircuit breaker located on the overheadconsole. The INST LT rheostat knob, alsolocated on the overhead console, adjusts lightintensity. Rotation of this knob operates thepower OFF/BRT switch that provides power toboth the 28 and 5 VDC instrument panellighting systems. The INSTR LT rheostat alsoenables the BRT/DIM swi tch on themiscellaneous control panel.
Four individual white cabin lights providepassenger area illumination at the pilot'sdiscretion. These four lights are protected bytheir own circuit breaker. They are regulatedby the CABIN LT circuit breaker switch and byindividual switches located near each light.When the CABIN LT switch is placed in theCABIN LT position, all four cabin lightsilluminate. When the switch is placed in PASS
position, the passengers can individuallycontrol each light. When the switch positionis OFF (centered), the cabin lights system isde-energized.
1-18-B. EXTERIOR LIGHTING SYSTEM
The anti-collision light is mounted on the topportion of the vertical fin. Position lights arelocated on the outboard side of the auxiliaryfins on the horizontal stabilizer, the lowerforward fuselage, left (red), and right (green)side, and on the tail rotor gearbox (white)fairing.
Dual fixed landing lights are located in thelower portion of the helicopter nose sectionand are controlled by the LDG LTS switchlocated on the collective switch box.
1-19. CYCLIC FLIGHT CONTROLSWITCHES
The cyclic control (Figure 1-6) incorporates aseries of switches. A two-position transmitswitch is mounted on the forward side thecyclic grip. Depressing the switch to the firstde tent pos i t ion w i l l ac t iva te in te rna lcommunication system (ICS) between thepilot and passengers. Depressing the switchto the second detent allows for the pilot totransmit on the radio selected on the audiopanel. Additional switches are used to controloptional equipment that may be installed onthe helicopter (i.e. cargo hook).
1-20. COLLECTIVE FLIGHTCONTROL SWITCHES
The co l lec t ive contro l (F igure 1-6 )incorporates a series of switches. Thecollective switches are the GOV RPM INCR/DECR switch, LDG LTS switch, and the STARTswitch.
1-34———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-6. Cyclic and Collective Control Switches
The GOV RPM INCR/DECR switch (Figure 1-6)is a spring-loaded, momentary contact switchthat enables the pilot to change the governorRPM actuator setting. With the switch held inthe INCR position, the actuator motor movesthe governor arm and the N 2 /NR RPMincreases. With the switch held in the DECRposition, the actuator motor moves thegovernor arm in the opposite direction anddecreases the N2/NR RPM. With the switch inthe normal (centered) position, the circuit isde-energized.
1-20-B. LANDING LIGHT SWITCH
The LDG LTS switch (Figure 1-6) located inthe middle of the collective switch boxcontrols the two fixed landing lights locatedunder the nose of the helicopter. The switch isa three-position switch, OFF/FWD/BOTH. Inthe OFF position, the circuit from the 28 VDCbus to the lights is disconnected and thelights are extinguished. In the FWD position,the most aft mounted forward facing light isenerg ized whi le the forward mounteddownward facing light is de-energized. In theBOTH position, both lights are energized.
1-20-C. START SWITCH
The STARTER switch (F igure 1-6) is aspr ing- loaded, momentary contactpushbutton switch that enables the pilot toengage the starter through the starter relay,when 28 VDC power is appl ied to thehelicopter electrical system. With the switchin the normal (extended) position, the circuitis de-energized.
1-21. ROTOR BRAKE
When the optional rotor brake is installed,application is limited to ground operationafter the engine has been shut down and NRhas decreased to between 38 to 30%. Enginestarts with the rotor brake engaged areprohibited.
1-22. VENTILATION SYSTEM
Air for cabin ventilation is obtained byopening the sliding window in each of theentrance doors. Additional ventilation to thecockpit may be obtained by pulling out theVENT control knobs and positioning theDEFOG BLOWER switch on the overheadconsole to the forward (ON) position.
When the VENT control knobs below theinstrument panel are pulled out, ram air isforced into the air intake grills (Figure 1-1) anddirected through a plenum and flapper valve.The flapper valve assembly is opened orclosed with the VENT control knob.
1-23. DEFOG SYSTEM
Two electrically-driven axial flow blowers areinstalled in the inlet end of the defrosternozzles. The blowers are controlled by aDEFOG BLOWER circuit breaker switch onthe forward row of the overhead console. Thedefog system is primarily used for ventilationand defogging during ground operation of thehelicopter.
1-24. PORTABLE FIREEXTINGUISHER
A manually-operated fire extinguisher isfurn ished and is located in the crewcompartment between the pilot and copilot
206L4_MD_01_0016
1-36———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
seats. A mounting bracket with a quickrelease mechanism is provided for rapidremoval of the extinguisher.
1-25. ELECTRICAL SYSTEM —GENERAL
The helicopter is equipped with a 28 VDCelectrical system (Figure 1-7). The helicoptere lec t r ica l power is prov ided by anickel-cadmium battery, a 200-amp starter/generator, and an external power receptacle.
The major components of the DC powersystem include the battery, starter/generator,DC voltage regulator, relays, 28 VDC bus, andcircuit breakers. All circuits in the electricalsystem are single wire with fuselage commonground return.
Controls for the electrical system are locatedon the overhead console and instrumentpanel. The battery and associated relays arelocated in the battery compartment in thenose of the helicopter. The starter/generatoris mounted on the right side of the engineaccessory gearcase. The DC voltage regulatorand electrical control panel (start relay,generator relay and shunt) are located on theaccessory she l f above the baggagecompartment. The main 28 VDC bus andcircuit breakers are located in the overheadconsole.
1-25-A. BATTERY SYSTEM
The battery system (Figure 1-7) includes thebattery, a battery relay, and a battery switchwith related wiring. The nickel-cadmium 24volt 17 ampere-hour battery is located in thenose of the helicopter.
During engine operation, a fully chargedbattery can be determined by moving the BATswitch from BAT to OFF and observing theeffect on the generator loadmeter. If thechange in indication is less than 2.5% on the
loadmeter (20 cell battery), the battery ischarged.
Battery can supply emergency DC load forapproximately 1 hour with 17 ampere drain.
The battery relay, located in the nose sectionforward of battery, is an electrically operatedswitch controlling battery current to the mainbus bar. The relay is actuated by the batteryswitch, located in the overhead console,which opens and closes the circuit to therelays energizing coil.
1-25-B. EXTERNAL POWER SYSTEM
The external power system (Figure 1-7)includes the external power receptacle,external power relay, and related wiring. Theexternal power receptacle is located in thelower front center of the nose section.
A 28 VDC Ground Power Unit (GPU) shall be500 amperes or less to reduce risk of starterdamage from overheating.
1-25-C. BATTERY CHARGING
As a maintenance function, battery chargingmay be accomplished with battery installed inhelicopter, due to minor depletion, using aGPU. The GPU must incorporate a goodquality constant voltage regulator, a variablevoltage selector, and an amperage indicator.
NOTE
It is recommended that the chargingprocedure does not exceed 45minutes in duration. Frequent use ofthis procedure may cause loss ofelectrolyte due to gassing throughelectrolysis. Reduced levels ofe lect ro ly te can cause ce l limbalances and lower overall batterycapacity. This procedure is notin tended to take the p lace o fscheduled battery maintenanceprocedures.
23 OCT 2008—Rev. 1———1-37
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-7. DC Power System206L4_MD_01_0005
EXTERNAL
POWER RELAY
2K2
2CR4
EXTERNAL
POWER RECEPTACLE
2J2
BATTERY
RELAY
2K1
BATTERY
SWITCH
2S1
BAT (ON)
OFF
BATTERY
2BT1
28
VD
C B
US
2TB1
GEN
RESET
5
2MG1
STARTER
GENERATOR
D
E
A
C
B
2VR1
DC CONTROL UNIT/
VOLTAGE REGULATOR
J
M
G
L
D
H
2S2
GEN SWITCHGEN
FIELD
10
RESET
OFF
GEN
STARTER
RELAY
1K1
STARTER SWITCH
4A1S3
1CB1 ENG START
BUS BAR
LINE CONTROL
RELAY
2K3
BUS BAR
LOADMETER
SHUNT
2R1
1-38———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
CAUTION
BATTERY HOT CAUTION LIGHTMUST BE CONTINUOUSLYMONITORED DURING THISPROCEDURE. IF BATTERY HOTLIGHT ILLUMINATES, BATTERYSWITCH SHALL BE SET TO OFF ANDGPU POWER REMOVED TO REDUCEPOSSIB IL ITY OF THERMALRUNAWAY.
BAT switch may be set to BAT (ON) after GPUpower is applied and voltage adjusted to 28.5VDC (do not exceed 28.5 volts). Batterycharging is supplied by the GPU and must bemonitored. Charging will be completed whenGPU output indicates approximately 8amperes. In the event BATTERY HOT warninglight illuminates, battery shall be turned OFFand GPU power removed.
1-25-D. GENERATOR SYSTEM
The generator system (Figure 1-7) includesthe generator portion of the starter/generator,DC control unit/voltage regulator, line controlrelay, generator reset switch, and generatorshunt.
The 30 volt, 200 amp generator is derated to180 amps (90% loadmeter indication). It islocated on the lower right aft side of theengine accessory gearbox and furnishespower at a regulated voltage for all DCelectrical components on the helicopter. Thegenerator output connects to the main buswhen the generator switch is positioned toGEN (ON) and generated voltage exceeds thevoltage on the bus by 0.30 to 0.42 volt.
The DC control unit/voltage regulator controlsthe operation of the electrical system. Itoperates the line control relay, regulatesgenerator voltage, and provides protectionagainst overvoltage, low voltage, and reversecurrent.
The line control relay is installed on theelectrical panel assembly located on theequipment she l f above the baggagecompartment. It opens or closes the powercircuit between the DC bus and the generator,and is controlled by the presence or absenceof a proper output voltage from the voltageregulator.
The generator shunt is also installed on theelectrical panel assembly on the equipmentshelf above the baggage compartment andprovides a voltage drop proportional to thecurrent to operate the loadmeter.
1-25-E. STARTER/IGNITER SYSTEM
The starter/igniter system includes the starterportion of the starter/generator, starter relay,field/igniter relay, igniter, and starter switch.
The starter/generator is energized by thestarter relay to spool the engine.
The starter relay is installed on the electricalpanel assembly located on the electrical shelfabove the baggage compartment. Directcurrent is supplied to the starter through thestarter relay when the start switch is engaged.
The field/igniter relay is also installed on theelectrical panel assembly located on thee lect r ica l she l f above the baggagecompartment. Direct current is supplied to theigniter and start field suppress section of theDC control unit through the field/igniter relaywhen the starter is engaged.
The tension capacitor discharge ignitionexciter (igniter exciter) is located on the lowerleft section of the engine accessory case. Thisexciter provides increased voltage to theigniter plug to insure fuel ignition during theengine start cycle.
1-25-F. CIRCUIT BREAKERS ANDSWITCHES
The majority of circuit breakers and switchesare mounted in the overhead console(Figure 1-4). Circuits are protected by the
The BAT switch (Figure 1-4) is installed on theoverhead console and controls the batteryrelay that connects the battery to the DC bus.The switch positions are OFF and BAT.
The BAT switch also configures the DC powerfeed to the left fuel boost pump between theDC bus (battery switch positioned to BAT),and the helicopter battery (BAT switchpositioned to OFF).
In the event the battery switch is positioned toOFF dur ing he l icopter operat ions , analternate circuit (Figure 1-8) is provided toallow operation of the left fuel boost pump.During this condition, with the FUEL VALVEswitch positioned to ON, battery voltage issupplied through the 10-amp left hand fuelboost circuit breaker, the fuel valve switch,the battery switch, and the 7.5-amp left fuelboost circuit breaker to the left fuel boostpump.
1-25-F-2. GENERATOR SWITCH
The GEN switch (Figure 1-4) is installed on theoverhead console and controls generatoroutput by opening and closing the generatorfield circuit. The switch is a double pole,double throw, spring loaded design with onlymomentary contact in the RESET position.The switch positions are GEN, OFF, andRESET.
With the generator switch positioned to GEN,its function is to complete the generator fieldcircuit between the starter generator and thegenerator control unit/voltage regulator.Under normal operating conditions, this willallow the generator control unit/voltageregulator to monitor and control the outputvoltage of the starter generator and in turnconnect the output of the generator to the 28VDC bus through the generator line controlrelay.
Positioning the generator switch to OFF,opens the generator field circuit whichremoves control of the generator control unit/voltage regulator from the generator andgenerator line control relay. The generatorline control relay will open removing thegenerator from the 28 VDC bus.
In the event an over voltage condition isdetected by the generator control unit/voltageregulator, an internal regulator trip relaycircuit will be activated. Positioning thegenerator switch to RESET, provides buspower from the battery to the generatorcontrol unit/voltage regulator, which will resetthe in terna l t r ip re lay c i rcu i t . I f themalfunction condition persists following theRESET, further attempts to reset should notbe made.
Additionally, following a start using a GPU,ensure battery switch is positioned to BAT(ON) and GPU is disconnected prior topositioning the generator switch to GEN.Positioning the generator switch to GEN withthe GPU connected may cause a reversecurrent situation and trip the generator offline.
1-25-F-3. GENERATOR FIELD AND RESETCIRCUIT BREAKERS
These 10- and 5-amp circuit breakers arelocated on the overhead panel (Figure 1-4)and protect the generator line control relayand the generator field relay. If either of thesecircuit breakers has tripped, they should bereset (once only). If the circuit breaker tripsagain, leave it out. A generator reset will notbe possible under these conditions.
1-25-F-4. START SWITCH
The START switch (Figure 1-6) is located onthe collective switch box. It contains a springloaded contact that provides momentarycontact in the START position.
1-40———Rev. 1—23 OCT 2008
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-8. Left Fuel Boost Pump Alternate Circuit Schematic
29 SEP 2006———1-41
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-42———Rev. 1—23 OCT 2008
1-25-G. ELECTRICAL SYSTEMINDICATORS
The DC loadmeter (Figure 1-4) indicates theload in amperes that is being supplied to the28 VDC bus by the generator. The maximumcontinuous load is 90% (180 amps).
The electrical system has one warning lightand two caution lights associated.
1-25-H-1. BATTERY HOT WARNING LIGHT
The 17 amp/hour nickel cadmium batteryincorporates an internal thermal switch thatwill illuminate the warning light if the internalbattery temperature reaches 60°C (140°F).Turn BAT switch OFF and land as soon aspractical.
1-25-H-2. BATTERY RELAY CAUTIONLIGHT
The BATTERY RLY light will illuminate if thebattery relay has remained in the closed(energized) position after the battery switchhas been selected to OFF (1-25-H-1). If thebattery relay remains closed after the batteryswi tch has been se lec ted to OFF, adisagreement circuit will advise the pilot thatthe relay is not in the selected position. Ifbattery relay fails to open, battery willcontinue to receive a charge from generator.To prevent this, pilot should turn GEN switchOFF. Battery will continue to run all electricalsystems through closed battery relay. Toreduce load on overheated battery, pilotshould pu l l c i rcu i t breakers on a l l
non-essential systems. LEFT BOOST circuitbreaker should be left in to ensure fueltransfer from front tanks to main tankcontinues. If BATTERY RLY light illuminateswhen the battery switch is placed in the OFFposition and BATTERY HOT warning light ison, land as soon as possible.
1-25-H-3. GENERATOR FAIL CAUTIONLIGHT
The GEN FAIL light is controlled by the linecontrol relay. The light will illuminate whenthe generator relay is de-energized and thegenerator output is not connected to the DCbus. The line control relay is energized by theDC control unit/voltage regulator whengenerator output climbs through a thresholdof 24 ±2.4 VDC. Prior to the generator relaybeing energized, the GEN FAIL light is on.Once the generator relay is energized, theGEN FAIL light is extinguished. If this cautionlight illuminates in flight, the pilot shouldfollow the generator failure proceduresprescribed in the BHT-206L4-FM-1, Section 3and land as soon as practical.
1-25-I. ELECTRICAL SYSTEM —COMPLETE ELECTRICALFAILURE
The electrical system design has provided away for the helicopter to function (fly) withoutelectrical bus power. If circumstances dictatethat the pilot turns off both the generator andbattery in flight, the electrical system isdesigned to operate the left boost pumpdirectly from the battery (paragraph 1-25-F-1).In the event of a complete electrical failure, afully charged battery will continue to operatethe left boost pump for more than 3 hours.This will provide boosted fuel pressure to theengine and transfer fuel from the forwardcells, making all of the fuel available andmaintaining longitudinal CG on the helicopter.
23 OCT 2008 Rev. 1 1-43
MANUFACTURER’S DATA BHT-206L4-MD-1
In the event of a complete electrical failure, thefollowing systems will be ON regardless ofsystem switch position (fail-safe on).
• Engine Anti-ice
• Hydraulics
• Particle Separator Purge (if applicable)
The following will continue to operate with thebattery and generator switched off:
• Turbine Engine
• Engine Oil Pressure Indication
• Left Boost Pump (as long as the batteryis not completely depleted)
• Transmission Oil Pressure Indication
• Torquemeter Indication
• N1 Tachometer Indication
• N2 Tachometer Indication
• NR Tachometer Indication
• Altimeter
• Airspeed
• Vertical Speed
• Inclinometer
• Compass
NOTE
A shut down helicopter in which theleft hand fuel boost pump circuitbreaker is pushed in, and the fuelvalve has been left ON, will duplicatethe total electrical failure situation.The left boost pump will operate anddeplete the battery. For normalshutdown, the fuel valve should beturned OFF to prevent bat terydepletion.
1-26. FUEL SYSTEM — GENERAL
The fuel system (Figure 1-9) comprises threeinterconnected crash-resistant fuel cells. The
two forward cells are located under the midpassenger seats and the aft fuel cell is locatedbelow and behind the aft passenger seat.
Helicopters equipped with the FUEL QTYswitch provide the operator with the capabilityto monitor fuel quantity in the forward fuelcells independently of the main fuel cell.
All fuel cells are serviced through the singlepoint filler port located on the right side of thehelicopter. Fuel will fill the aft cell until itreaches the top of the gravity feed standpipe,as follows:
At this point, fuel will begin to fill the twoforward fuel cells through the standpipe untilthe fuel level in the forward cells is equal tothe level in the aft fuel cell, as follows:
At this point, the forward fuel cells are full on206L3+ IGW Upgrade helicopters S/N 51001through 51243 (without FUEL QTY switch) andon 206L1+ IGW Upgrade helicopters. For206L4 and 206L3+ IGW Upgradehelicopters S/N 51244 and subsequent (withFUEL QTY switch), fuel will then fi ll theforward and aft fuel cells together until theforward cells are full (407.3 pounds). Theremainder of the aft fuel cell will then be filled,as follows:
206L3+ IGW Upgrade S/N 51001 through 51243 without FUEL QTY switch
319.1 pounds
206L4 and 206L3+ IGW UpgradeS/N 51244 and subsequent with FUEL QTY switch
109.5 pounds
206L1+ IGW Upgrade 270.6 pounds
206L3+ IGW Upgrade S/N 51001 through 51243 without FUEL QTY switch
476.0 pounds
206L4 and 206L3+ IGW UpgradeS/N 51244 and subsequent with FUEL QTY switch
195.8 pounds
206L1+ IGW Upgrade 425.0 pounds
1-44 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-9. Fuel System206L4_MD_01_0028
17
1
2
3
6
7
8
10
9
4
5
11
14
15
16
21
20
22
23
1819
13
RIGHT FUEL CELLVENT (TOP)
LEFT FUEL CELLVENT (TOP)
PROBE
DUAL ELEMENT EJECTOR PUMP
FUEL CELLINTERCONNECT TUBEAND GRAVITY STANDPIPE
CHECKVALVES
INLINEFUEL FILTERS
PROBE
FILLER CAP
VENT(TOP)
UPPER FUEL QUANTITYPROBE AND TRANSMITTER
A/F MOUNTEDFUEL FILTER
TO ENGINE
SHUT-OFFVALVE
FUEL PRESSURETRANSDUCER
ENGINE SUPPLY LINE
MANIFOLD
TWO SINGLE ELEMENT BOOSTPUMPS(ELECTRIC)
THERMALRELIEF VALVES
FUEL TRANSFERFEED LINES
FUEL FLOWSWITCHES
12
Right forward fuel cell Forward fuel cells vent line Fuel flow switches Forward to aft fuel cell transfer line Fuel cell interconnect tube and gravityfeed standpipe Aft fuel cell Fuel filler cap Fuel pressure transducer Fuel shutoff valve Airframe fuel filter Aft fuel cell vent line Engine fuel supply line
1.2.3.4.5.
6.7.8.9.
10.11.12.
Aft fuel cell upper fuel quantity probe and transmitter Aft fuel cell lower fuel quantity probe Fuel boost pumps Fuel manifold including check and thermal relief valves Low fuel level detector and electric sump drain valve Fuel transfer feed lines In-line fuel filters and check valves Left forward fuel cell Forward fuel cell fuel quantity probe Forward fuel cell manual sump drain valve Dual element ejector pump
13.14.15.16.17.18.19.20.21.22.23.
BOOST PUMP PRESSURE
INTERCONNECT FUEL
TRANSFER FUEL PRESSURE
EJECTOR PUMP FUEL PRESSURE
FUEL CELL VENT
TO AFTVENT
TO ENGINE
TO FWDVENT
AFT FUEL CELL
ONE-WAY CHECK VALVES
FORWARD TO AFT CELL TRANSFERLINE
23 OCT 2008 Rev. 1 1-45
MANUFACTURER’S DATA BHT-206L4-MD-1
All fuel cells are vented overboard by acommon vent line.
Fuel from the aft fuel cell is provided to theengine by two electric boost pumps which aremounted on a sump plate assembly at thebottom of the aft cell. Pressure fuel from thepumps is fed through a fuel manifold, fuelshutoff valve and airframe mounted fuel filterbefore reaching the engine. Pressure fuel isalso directed by the manifold forward to a dualelement ejector pump which in turn transfersfuel from the forward fuel cells to the aft fuelcell.
A low fuel level detector and electricallyoperated sump drain valve are also mountedon the sump plate assembly at the bottom ofthe aft cell. The low level detector activates theFUEL LOW caution light (50 to 75 pounds) andthe drain valve is activated by a switch on thelower right side of the aft fuselage. The switchis deactivated when the fuel valve switch isON to prevent inadvertent activation duringflight.
A capacitance type fuel quantity gaugingsystem using three probes is used to providea signal to the fuel quantity indicator. Oneprobe is located in the left forward fuel cell andtwo are located in the aft fuel cell. The upperprobe in the aft fuel cell is also utilized as atransmitter which sends the electrical signalsfrom the three probes to the fuel quantityindicator.
1-26-A. FUEL SYSTEM CIRCULATION
Fuel flow to the engine — When the BATswitch is turned ON, the boost pumps(Figure 1-9) begin to operate, pumpingpressurized fuel into the fuel manifold. Thefuel enters the manifold through the boostpump one-way check valves and is directed
into the engine supply l ine and thepressurized fuel transfer feed lines servicingthe ejector pump. The engine supply line fuelflows past the fuel pressure transducer, thento the fuel valve. When the fuel valve isopened, the fuel flow continues through to theairframe fuel filter, into the engine driven fuelpump, and into the fuel control unit. If the pilothas opened the throttle, fuel will flow from thefuel control unit, through the fuel nozzle, andinto the combustion chamber.
Fuel flow to the ejector pump — When the BATswitch is turned ON, the boost pumps begin tooperate, pumping pressurized fuel into the fuelmanifold. The fuel enters the manifold throughthe boost pump one-way check valves and isdirected into the engine supply line and thepressurized fuel transfer feed lines servicingthe dual element ejector pump. There are twopressurized lines moving fuel forward, one foreach fuel boost pump, connected to itsrespective side of the ejector pump. Thatportion of the fuel flowing to the ejector pumpflows through a screen mesh filter, an in-linecheck valve, through a flow switch and intothe dual element ejector pump. The movementof the pressurized fuel through the ejectorpump creates a venturi effect inside the pump.This venturi effect causes the transfer of fuelfrom the forward cells to the aft cell throughthe transfer return line. If the gravity standpipeis submerged in fuel and the boost pump(s)are runn ing , there w i l l be a constantcirculation of fuel between the forward and aftcells.
1-26-B. FUEL BURN SEQUENCE
When the engine is started and fuel isconsumed, there a re burn sequencescontrol led by the height of the gravitystandp ipe (F igure 1-10) . These burnsequences reverse the fill sequence andmanage the transfer of fuel to the aft cellwhere it can be consumed, and also managethe shift in longitudinal CG that results fromfuel burn.
206L4 and 206L3+ IGW UpgradeS/N 51001 and subsequent
752.8 pounds
206L1+ IGW Upgrade 669.1 pounds
1-46 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-10. Fuel Burn Sequence (Sheet 1 of 3)
With boost pumps operating, fuel is circulating from the forward cells through the ejector pump to the aft cell, then returning to the forward cell. See position 1 on fuel C of G burn curve.
1. FULL FUEL – 752.8 LBS
Fuel is consumed from forward cells until depleted. Fuel center of gravity moves from position 2 to position 3 on fuel C of G burn curve.
FUEL BURN ZONE
FUEL REMAINING
3. PARTIAL FUEL – BURN FROM 476 TO 319.1 LBS
Fuel is consumed from aft cell until depleted.Fuel center of gravity moves forward from position3 to position 4 on fuel C of G burn curve.
4. PARTIAL FUEL – BURN FROM 319.1 TO 0 LBS
Fuel is consumed from aft cell down to the level of the gravity standpipe. Fuel center of gravity moves from position 1 to position 2 on fuel C of G burn curve.
2. PARTIAL FUEL – BURN FROM 752.8 TO 476 LBS
206L4_MD_01_0057
BURNZONE
FUEL BURN
START
FUEL BURN
START FINISH
FUEL BURN
START FINISH
752.8LBS
752.8LBS
476LBS
BURNZONE
476 LBS
319.1 LBS
BURNZONE
0
319.1 LBS
3
4(0 LBS)
(319.1 LBS)
(476 LBS)
(752.8 LBS)
0
10
20
30
40
50
60
70
80
90
100
110
120
120 124 128 132 136 140 144
Fuselage Station - Inches
Fu
el-
US
Gallo
ns
0
50
100
150
200
250
300
350
400
450
3050 3150 3250 3350 3450 3550 3650
Fuselage Station - millimeters
Fu
el-
Lit
ers
2
1
FUEL CENTER OF GRAVITY (C of G) BURN CURVE (NO SWITCH)
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1 78QTY
LBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
EMPTY
FUEL BURN SEQUENCE(206L3+ IGW UPGRADE S/N 51001 THROUGH 51243)
WITHOUT FUEL QUANTITY SWITCH
23 OCT 2008 Rev. 1 1-47
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-10. Fuel Burn Sequence (Sheet 2 of 3)
With boost pumps operating, fuel is circulating from the forward cells through the ejector pump to the aft cell, then returning to the forward cell. See position 1 on fuel C of G burn curve.
1. FULL FUEL – 752.8 LBS
FUEL BURN SEQUENCE(206L4 AND 206L3+ IGW UPGRADE S/N 51244 AND SUBSEQUENT)
WITH FUEL QUANTITY SWITCH
Fuel is consumed from forward cells until depleted. Fuel center of gravity moves from position 3 to position 4 on fuel C of G burn curve.
FUEL BURN ZONE
FUEL REMAINING
4. PARTIAL FUEL – BURN FROM 195.8 TO 109.5 LBS
Fuel is consumed from aft cell until depleted.Fuel center of gravity moves forward from position4 to position 5 on fuel C of G burn curve.
5. PARTIAL FUEL – BURN FROM 109.5 TO 0 LBS
C of G moves forward until the aft level is equal to the forward level (407.3 pounds). Fuel center of gravity moves from position 1 to position 2 on fuel C of G burn curve.
2. PARTIAL FUEL – BURN FROM 752.8 TO 407.3 LBS
Fuel is consumed equally from aft and forward cells down to the level of the gravity standpipe. Fuel center of gravity moves from position 2 to position 3 on fuel C of G burn curve.
3. PARTIAL FUEL – BURN FROM 407.3 TO 195.8 LBS
206L4_MD_01_0058
BURNZONE
FUEL BURN
START
FUEL BURN
START FINISH
FUEL BURN
START FINISH
FUEL BURN
START FINISH
BURNZONE
752.8LBS
752.8LBS
407.3LBS
195.8 LBS
407.3 LBS
BURNZONE
195.8 LBS
109.5 LBS
BURNZONE
0
109.5 LBS
1
2
4
5
110 114 118 122 126 130 134 138
FUSELAGE STATION - INCHES
FUEL CENTER OF GRAVITY (C of G) BURN CURVE (WITH SWITCH)
FUSELAGE STATION - MILLIMETERS
0
10
20
30
40
(0 LBS)
(109.5 LBS)
(195.8 LBS)
50
60
70
80
90
100
110
120
FU
EL
- U
.S.
GA
LL
ON
S
FU
EL
- L
ITE
RS
0
50
100
150
200
250
300
350
400
450
2800 2900 3000 3100 3200 3300 3400 3500
(407.3 LBS)
(752.8 LBS)
3
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1 78QTY
LBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
EMPTY
1-48 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-10. Fuel Burn Sequence (Sheet 3 of 3)
With boost pumps operating, fuel is circulating from the forward cells through the ejector pump to the aft cell, then returning to the forward cell. See position 1 on fuel C of G burn curve.
1. FULL FUEL – 669.1 LBS
Fuel is consumed from forward cells until depleted. Fuel center of gravity moves from position 2 to position 3 on fuel C of G burn curve.
FUEL BURN ZONE
FUEL REMAINING
3. PARTIAL FUEL – BURN FROM 425 TO 270.6 LBS
Fuel is consumed from aft cell until depleted.Fuel center of gravity moves forward from position3 to position 4 on fuel C of G burn curve.
4. PARTIAL FUEL – BURN FROM 270.6 TO 0 LBS
Fuel is consumed from aft cell down to the level of the gravity standpipe. Fuel center of gravity moves from position 1 to position 2 on fuel C of G burn curve.
2. PARTIAL FUEL – BURN FROM 669.1 TO 425 LBS
206L4_MD_01_0059
BURNZONE
FUEL BURN
START
FUEL BURN
START FINISH
FUEL BURN
START FINISH
669.1LBS
669.1LBS
425LBS
BURNZONE
425 LBS
270.6 LBS
BURNZONE
0
270.6 LBS
4
(0 LBS)
(425 LBS)
(669.1 LBS)
(270.6 LBS)
1
0
10
20
30
40
50
60
70
80
90
100
115 120 125 130 135 140 145
Fuselage Station - Inches
Fu
el-
US
Gallo
ns
0
50
100
150
200
250
300
350
2920 3020 3120 3220 3320 3420 3520 3620
Fuselage Station - milimeters
Fu
el-
Lit
ers
FUEL CENTER OF GRAVITY (C of G) BURN CURVE(206L1 ENGINE UPGRADE)
2
3
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1 78QTY
LBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
0
1
23 4 5
6
78
FUEL
QTYLBSx 100
EMPTY
FUEL BURN SEQUENCE(206L1+ IGW UPGRADE)
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-49
1-26-C. FUEL CELLS
The helicopter has three crash-resistantbladder type fuel cells (Figure 1-9) that areserviced through a single point filler portlocated on the right side of the helicopter.These tanks have a combined capacity of112.0 U.S. gallons (110.7 U.S. gallons usable)for 206L4 and 206L3+ IGW Upgrade and 99.4U.S. gallons (98.4 U.S. gallons usable) for the206L1+ IGW Upgrade. The aft cell is locatedbelow and behind the aft passenger seats andhas a total capacity of 89 gallons for 206L4and 206L3+ IGW Upgrade and 75.7 gallons forthe 206L1+ IGW Upgrade. The forward cellsare located under the mid passenger seats.Each forward cell has a capacity of 11.5gallons (43.5 L). A standpipe permits fuel togravity transfer from the aft to forward cellswhenever the level of fuel in the aft cell isabove the height of the standpipe. The gravityand pressure return lines that connect the aftand forward cells are routed along the floor ofthe main cabin and a re covered by aprotective floor panel.
1-26-D. FUEL BOOST PUMPS
Two electric boost pumps (Figure 1-9) aremounted on a common sump plate on thelower section of the aft fuel cell . Eachindividual pump is capable of sufficientpressure to supply fuel to the engine-drivenfuel pump. A portion of the output from eachpump (left or right) is also used for operatingits respective side of the dual element ejectorpump. Each boos t pump de l ive rs i tspressurized fuel flow into the fuel manifold.
The LEFT FUEL BOOST circuit breaker in theoverhead panel is outlined by a yellow borderto identify that it has an alternate electricalcircuit (Figure 1-8). If a malfunction requiresthat OFF be selected for both the GEN andBAT switches, all DC bus power is lost. In thiscase, the only electrical component thatcontinues to operate is the left fuel boostpump. This pump receives power from thebattery through the fuel valve switch eventhough the BAT switch is OFF. This enables
fuel transfer to continue from the forwardcells and all fuel remains usable, assumingthat the battery is not depleted.
The recommended engine shutdownprocedure includes turning OFF the fuelvalve. This opens the alternate circuit andprevents power from reaching the left fuelboost pump while the BAT switch is OFF. Ifthe fuel valve is not turned off, the left boostpump continues to operate and depletes thebattery. Pulling the LEFT BOOST PUMP circuitbreaker is not recommended.
1-26-E. FUEL MANIFOLD
The fuel manifold (Figure 1-9) directs fuel flowthrough a one-way check valve (which willclose and prevent re-circulation in the eventof a single boost pump failure), combines theflow from each pump, and delivers fuelthrough a common line to the engine. Aportion of the fuel from each boost pump isrouted out of the manifold through a transferfeed line on the floor, an in-line filter, an in-linecheck valve, a flow switch, and into therespective side of the dual element ejectorpump.
1-26-F. FUEL PRESSURE TRANSDUCER
The fuel pressure transducer (Figure 1-9)converts a wet line pressure into an electricalsignal. The fuel pressure transducer ispowered by 28 VDC power and measures thecombined fuel boost pump pressure to theengine. This data is displayed on the fuelpressure gauge in the cockpit. The transduceris installed in the engine supply line prior tothe fuel shutoff valve. The fuel pressuregauge will display the proper pressureindication with the BAT switch ON and boostpump(s) energized, regardless of fuel valveposition.
1-26-G. FUEL VALVE
The fuel valve (Figure 1-9) is located alongwith the transducer behind the inspectionpanel on the right side of the helicopter above
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-50———Rev. 1—23 OCT 2008
the refueling port. The FUEL VALVE switch islocated on the lower right corner of theinstrument panel and electrically controls thefuel valve. When the switch is in the ONposition, fuel flows through the open valveand into the airframe fuel filter. When theswitch is placed into the OFF position, thefuel valve closes and stops fuel flow to theengine.
1-26-H. AIRFRAME FUEL FILTER
The airframe fuel filter (Figure 1-9) is locatedon the right side of the forward enginefirewall. It is equipped with a built-in filterbypass valve and impending bypass switchthat is electrically connected to an ambercaut ion l ight on the caut ion panel . Atapproximately 1 PSID (differential fuelpressure), the FUEL FILTER caution lightilluminates. The illumination of this lightindicates that the filter is in impendingbypass. The filter bypasses at approximately4 PSID, however, there are no further cockpitindications. A red press-to-test button islocated on the top of the filter head to checkthe electrical connection to the fuel filtercaution light.
An in-line fuel filter (Figure 1-9) with a meshscreen is installed in the pressure transferfeed line from each boost pump to the ejectorpump. These filters protect componentsdownstream in the fuel flow from debris andpossible blockage. The in-line check valvesare immediately downstream from the filtersand are opened by fuel flow (pressure). Thepurpose of the in-line check valves is toprevent air from entering the fuel system inthe event of a dual boost pump failure. To bemore specific, if a check valve is defectiveand both boost pumps are inoperative, airmay be drawn into the fuel system by theengine driven fuel pump once the fuel level inthe aft cell is below the top of the fuel gravityfeed standpipe.
1-26-J. FUEL FLOW SWITCHES
A fuel flow switch (Figure 1-9) in each fuelfeed transfer line continuously monitors fuelflow to the dual element ejector pump. Theseflow switches will activate the L/FUEL PUMPor R/FUEL PUMP caution light if fuel flow isinadequate in the fuel feed transfer line.Inadequate flow may be the result of boostpump failure, in-line filter clogging, or ejectorpump clogging. If fuel flow is restricted, thesystem may not have sufficient pressure tooperate the ejector pump.
1-26-K. DUAL ELEMENT EJECTOR PUMP
The dual element ejector pump (Figure 1-9) isconnected to each of the forward fuel cellsand to the pressurized fuel feed transfer linesfrom the boost pumps. The ejector pump willutilize this pressurized fuel flow to create lowpressure inside the ejector pump. This lowpressure (venturi) effect will suck the fuelfrom the respective forward cell and move thefuel aft through a pressure transfer linebetween the ejector pump and the aft cell.
1-26-L. FUEL LOW LEVEL DETECTOR
A float type low-level detector (Figure 1-9) ismounted on the sump plate in the aft fuel celland is calibrated to activate the FUEL LOWcaution light on the caution panel whenapproximately 10 gallons (50 to 75 pounds) offuel remains in the aft tank.
1-26-M. FUEL SUMP DRAIN VALVES
An electrically activated sump drain valve(Figure 1-9) is located on the sump plate inthe aft fuel cell and is electrically operated bymeans of the drain switch located aft of theright side passenger door. The drain switch iswired through the FUEL VALVE switch on theinstrument panel. The FUEL VALVE switchmust be in the OFF position to operate thedrain valve.
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-51
Two manually-operated sump drain valves aremounted in the forward tanks and may beaccessed from below the fuselage.
1-26-N. FUEL QUANTITY TRANSMITTER
A fuel quantity transmitter (Figure 1-9) islocated on the electrical shelf behind the rearforward facing seat. Three separate fuelquantity probes (one located in the forwardleft cell and two located in the aft cell) sendsignals that indicate fuel quantity (based onfuel levels and density) to the fuel quantityt ransmi t ter. A microprocessor in thetransmitter uses the information provided bythe three probes to compute the weight of thefuel. The fuel quantity indicator then displaysthe combined fuel quantity in pounds.
The pilot can monitor fuel quantity in theforward fuel cells independently of the aft fuelcell by selecting FWD on the FUEL QTYswitch (206L4 and 206L3+ IGW Upgrade S/N51244 and subsequent) located to the left ofthe fuel quantity indicator. Under normalcondi t ions, in level f l ight , fuel beginstransferring from the forward fuel cell to mainfuel cell at the following approximate total fuelquantities:
Transfer is complete and forward cellsemptied at the following approximate totalfuel quantities:
1-26-O. FUEL SYSTEM — EMERGENCYOPERATION
There are four caution lights associated withthe fuel system:
L FUEL PUMP/ R FUEL PUMP — When theflow switches detect a problem in flow to theejector pump, they illuminate the respectiveL/FUEL PUMP or R/FUEL PUMP cautionlight(s). A flow problem may indicate that thepump has failed or that the in-line filter isclogged.
In the case of a single boost pump failure,pressurized fuel continues to flow to theengine and the ejector pump from theremaining boost pump. The pilot shoulddescend below 6000 HP and land as soon aspractical. If a dual boost pump failure shouldoccur, the ejector pump will not function andcannot transfer fuel from the forward cells tothe aft cell. In this case also, the pilot shoulddescend below 6000 feet HP and land as soonas possible. All of the fuel remaining in theforward cells is trapped and should beconsidered unusable. Depending upon thelevel of fuel in the forward cells at the time offailure, this quantity may be as much as:
206L3+ IGW UpgradeS/N 51001 through 51243 without FUEL QTY switch
476.0 pounds
206L4 and 206L3+ IGW UpgradeS/N 51244 and subsequent with FUEL QTY switch
407.3 pounds
206L1+ IGW Upgrade 425.0 pounds
206L3+ IGW UpgradeS/N 51001 through 51243 without FUEL QTY switch
319.1 pounds
206L4 and 206L3+ IGW Upgrade S/N 51244 and subsequent with FUEL QTY switch
109.5 pounds
206L1+ IGW Upgrade 270.6 pounds
206L3+ IGW UpgradeS/N 51001 through 51243 without FUEL QTY switch
156.9 pounds
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-52———Rev. 1—23 OCT 2008
FUEL FILTER — The fuel filter caution lightilluminates when the differential pressureswitch senses a difference between input andoutput pressure greater than 1 PSI. Thispressure difference is an indication that thefilter is becoming clogged. The filter willcontinue to function normally until thepressure differential reaches approximately 4PSI. At this time, the bypass valve will openand permit fuel to bypass the filter enroute tothe engine fuel pump. Proper pilot action is toland as soon as practical. The filter must beinspected/replaced prior to the next flight.
FUEL LOW — The FUEL LOW light illuminateswhen approximately 50 to 75 pounds of fuel(approximately 10 gallons) remain in the aftcell. This caution light is activated by the lowlevel detector which is mounted on the sumpplate in the a ft cel l and is complete lyindependent of the fuel quantity system. If thelight illuminates, the pilot should verify thefuel quantity, then land as soon as practical. Ifthere is a fuel transfer problem, the fuelquantity indicated may be higher than normal,indicating fuel is trapped in the forward cellsand not available to the engine.
1-27. PARTICLE SEPARATOR
The optional particle separator (Figure 1-11)may be installed in place of the standard inletscreen. The particle separator providescontinuous protection for the compressoragainst damage from the ingestion of sand,dust, and other foreign material. The unit
consists of the separator, bleed air tubing andhoses, compressor wash fittings, and otherrequired hardware.
The separator section has 281 filter elementsand is positioned so that all inlet air mustpass through the filter elements beforeenter ing the engine. Each of the f i l terelements or tube assemblies in the separatorconsists of a vortex generator bonded into aninlet tube and a second smaller tube to form ascavenge chamber.
Particles such as dust, dirt or sand, enteringthe filter elements, are spun in the vortexgenera tor (F igure 1-11 ) and hur ledcentrifugally into the scavenge chamber. Theyare ejected overboard by the venturi effectcreated by engine bleed air as it is dischargedthrough the ejector tubes. This scavengeflow, which carries the particulate materialwith it, accounts for approximately 8% of theinlet air flow.
A window is installed on each side of thecowling to permit visual inspection of theseparator plenum chamber. The ejector tubesare mounted on each side of the cowling justbelow these windows. A compressor washfitting permits the introduction of wash waterdirectly into the engine inlet.
Insta l la t ion imposes no aerodynamicrest r ic t ions and has no e f fect on themaximum gross weight. The kit is equippedwith a particle separator purge switch locatedon the misce l laneous contro l pane l(Figure 1-4). With the switch ON, engine bleedair is used to purge the particle separatorplenum. Due to the pressure drop at the inletand the use of bleed air to purge the system,there is a slight reduction of power available.With the switch OFF, slightly more enginepower is available because of the reduction inbleed air use.
On 206L1+ and 206L3+ IGW Upgradehelicopters without a particle separator purgeswitch, bleed air is continually provided topurge the particle separator plenum.
206L4 and 206L3+ IGW UpgradeS/N 51244 and subsequent with FUEL QTY switch
86.3 pounds
206L1+ IGW Upgrade 154.4 pounds
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-53
Figure 1-11. Particle Separator and Vortex Generator
206L4_MD_01_0027
DIRTY AIR IN
VORTEX GENERATION
VORTEX TUBE
SCAVENGE FLOWWITH CONTAMINANT
TO ENGINEINLET
OUTLETTUBE
CLEAN AIR
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-54———Rev. 1—23 OCT 2008
Flight in falling or blowing snow requires theinstallat ion of BHT-206L4-FMS-7 SnowDeflector Kit, either with or without theparticle separator installed.
1-28. POWER PLANT
The Rol ls -Royce 250 -C30P eng ine(Figure 1-12) is an internal combustionturboshaft engine featuring a free powerturbine. The gas generator is composed of asingle stage, single entry centrifugal flowcompressor directly coupled to a two-stagegas generator turbine. The power turbine is atwo-stage free turbine which is gas coupled tothe gas generator turbine. The integralreduction gearbox has front and rear drivesplines to mate the helicopter drives. Theengine has a single combustion chamber withsingle ignition. The output shaft centerline islocated below the centerline of the engineturbine shafting and the single exhaust outletis directed upward.
The Model 250-C30P is a thermodynamicallyrated 650 shaft horsepower (SHP) engine. Forinstallation in the 206L4 and on 206L1+ and206L3+ IGW Upgrade he l icopters , themaximum fuel flow the engine will receive hasbeen set to 356 PPH. This fuel flow willproduce approximately 575 SHP or 117%torque at sea level on a standard day. This flatrating is set lower than the thermodynamiclimit to protect the drivetrain from excessivetorque. Torque gauge markings create aderated limit. At takeoff the pilot may use 490SHP or 100% torque (5 minute limit), and 370SHP for continuous operation 75% torque.This combination of flat rating and deratedlimits will enable the helicopter to haveconsistent lifting performance from sea levelto moderate altitudes (see performance chartsin BHT-206L4-FM-1, Section 4 for detailedinformation).
Air is supplied to the engine through airscoops on either side of the fuselage locatedimmediately aft of the transmission cowling.The air is directed through a protective inlet
screen (or particle separator), the air inlet, andthen into the compressor.
The compressor compacts and increases thevelocity of this incoming air, then dischargesit into two external ducts that direct the air tothe combustion section of the engine. Anautomatic bleed air system is incorporated toallow a portion of the compressed air to ventoverboard during the start cycle. This willallow the engine speed to accelerate at adesired rate and minimize the opportunity forcompressor stall to occur. As compressorspeed increases, the pneumatically controlledbleed valve gradually closes.
The combustion section burns the fuel atpeak efficiency. After the igniter is energizedand when fuel is introduced into the chamberat a precisely controlled rate, the mixtureignites. Combustion is even and continuousas long as the proper air/fuel mixture ismaintained.
The turbine section consists of a two-stagegas producer turbine, gas coupled to thetwo-stage power turbine. The gas producerturbine (N1) rotates at 50,940 RPM at 100% N1speed and the two-stage power turbine (N2)rotates at a constant 30,650 RPM at 100% N2speed. The turbine section is designed toutilize the energy from combustion to drivevarious gear trains. These gear trains drivethe components that will sustain the engineoperation and provide an output to thehelicopter drivetrain.
The engine is horizontally mounted aft of thetransmission and above the fuselage tosimplify the drive system, improve the inlet/exhaust arrangement, and to reduce cabinnoise. The engine is supported by three bipodmounts on the engine deck to providestructural integrity. The transmission iscoupled to the engine by means of the maininput driveshaft and freewheeling unit.
The major engine components are acompressor, combustion section, turbinesection, and power and accessory gearbox.
23 OCT 2008 Rev. 1 1-55
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-12. Power Plant (Sheet 1 of 2)206L4_MD_01_0023
Figure 1-12. Power Plant (Sheet 2 of 2)206L4_MD_01_0031
INLET AIR
COMPRESSOR BLEED AIR
COMPRESSOR DISCHARGE AIR
COMBUSTION GASES/EXHAUST GASES
FUEL
COMPRESSOR DISCHARGE AIR TUBE
COMBUSTIONLINER
TURBINE TOCOMPRESSOR COUPLING
EXHAUST AIR OUTLET
SPUR ADAPTERGEAR SHAFT
FUEL
NOZZLE
IGNITER
PLUG
POWER OUTPUT GEARPOWER
OUTPUT
PINION GEAR
INLET
AIR
COMPRESSOR
SECTION
ACCESSORY
GEARBOX
SECTION TURBINE SECTION
AIR FLOW
COMBUSTION SECTION
TORQUEMETER
GEAR
N1
N2
COMPRESSORROTOR
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-57
1-28-A. COMPRESSOR
The compressor assembly (Figure 1-12)consists of a compressor front support,shroud housing, diffuser, rear supportassembly, centr i fuga l impel ler, scrol lassembly, mount assembly and bearings. Thefront support encloses the front bearing andsupports it through five hollow inlet struts.The struts are hollow to provide an anti-icingcapability using compressor discharge air. Oilpassages are incorporated to providelubrication of the compressor front bearing.The compressor rear bearing is mounted inthe rear support assembly and is lubricatedfrom the gearbox.
The compressor impeller is machined from asingle piece of forged titanium. The vanest rans i t ion f rom ax ia l to cent r i fuga l ,eliminating the need for stators. At 100% N1the compressor is rotating at 50,940 RPM. Thecompression ratio is approximately 8.5:1 (8.5Bars or 123 PSI). This rapid compression willincrease the temperature of the compressedair to approximately 291°C (555°F).
The compressor impel le r requ i res aconsiderable amount of shaft horsepower topump air to provide the required air flow andpressure. The shaft horsepower required bythe compressor rotor varies directly with airdensity and N1 RPM, and the gas producerturbine rotor must develop the horsepowerrequired. As a general rule, a compressor willconsume approximately 2 HP for eachhorsepower at the output shaft.
1-28-A-1. COMPRESSOR BLEED SYSTEM
Two features are used to control pressures inthe compressor. The compressor liner has anarrow slot machined into the perimeter nearthe front. This slot is known as the inducerbleed port. During low RPM operation, airexits through the port, reducing drag. Thisreduction of drag allows rapid acceleration. Ath igher RPM, when maximum eng ineef f ic iency is requ i red , a i r enters thecompressor through this slot, increasing the
volume of air entering the engine. Thisinducer bleed port works in conjunction witha pressure-activated, diffuser scroll mountedbleed valve, to complete the system design.
The bleed valve (Figure 1-12) is open duringstarting and idle operation and remains openuntil a predetermined pressure ratio isobtained. At this pressure ratio, the valvebegins to modulate from open to the closedposition. It reaches the full closed positionduring acceleration to full power. When poweris reduced, the bleed valve will open. Openingand closing of the bleed valve is controlled bycompressor discharge pressure in thediffuser.
1-28-A-2. ANTI-ICING SYSTEM
Opera t ion o f the eng ine dur ing ic ingconditions could result in ice formations onthe compressor front support. If ice wereallowed to build up, airflow to the enginecould be a f fec ted and per formancedecreased. The engine anti-icing system isdesigned to prevent ice formation on thecompressor front support. The pilot activatesthe system by p lac ing the ENGINEANTI- ICING switch to the ON posit ion(Figure 1-4).
When the system is in operation, compressordischarge air, which has been heated due tocompression, flows through the anti-icingva lve (F igure 1-12 ) and tube to thecompressor front support passages. Hot airflows between the double wall outer shell andinto the five hollow radial struts. The hot airflowing through the radial struts exhaustseither from small slots in the trailing edge ofthe struts or from the double wall bullet nosehub of the compressor front support. Thecompressor inlet guide vanes and frontbearing support hub are the only enginecomponents with anti-icing provisions. Thebleed air shutoff valve is solenoid controlled.In the event of a total electrical failure, thesystem is fail-safe to on (no electrical power =anti-ice ON).
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-58———Rev. 1—23 OCT 2008
1-28-B. COMBUSTION SECTION
The combustion sect ion (F igure 1-12)consists of an outer combustion case and acombustion liner. The combustion liner islocated inside the outer case. The fuel nozzleand igniter plug are mounted in the aft end ofthe outer combustion case. Compressordischarge air enters the combustion case viatwo air tubes connected to the diffuser scroll.Air then enters the combustion liner throughvanes and holes in the liner dome and skin.The air is mixed with fuel sprayed from thefuel nozzle and combustion takes place. Airenters the combustion liner in such a mannerthat the flame pattern is prevented fromtouching the wall of the combustion liner,thus insulating and protecting the liner fromdamage and de format ion caused byexcessive heat. Combustion is even andcontinuous as long as the proper air/fuelmixture is maintained.
Approximately 20 to 25% of the air deliveredto this section is required to burn the fuel. Theother 75 to 80% is used to expand the gasesand cool the internal parts. Approximately 2%of the air is used to seal oil passages.
1-28-C. TURBINE SECTION
As the gas stream leaves the combustionchamber, it passes to the turbine section(Figure 1-12). This high-energy gas streampowers the two turbine sections to sustain theairflow through the engine and provide outputpower. The turbine section consists of a gasproducer turbine support, power turbinesupport, a two-stage gas producer turbinerotor, a two-stage power turbine rotor, and anexhaust collector support. The gas producerturbine drives the compressor and certainengine accessories through the N1 drivetrain.At 100% N 1 , the turb ine ro ta tes a tapproximately 50,940 RPM.
The power turbine drives the power outputshaft and certain engine accessories. At 100%N2 RPM, the power turbine rotates at aconstant 30,650 RPM.
The turbines take advantage of the impact andreaction of the gases passing through the gasproducer and power turbines. The turbinesmay be broadly classified as an impact,reaction type. Varying fuel flow changes thetemperature of the gases passing through theturbine section and the amount of energy inthe gas stream. This variation of gas energywill result in a change in the expansion rate ofthe gases as well as a change in gas velocitythrough the turbines. Consequently, anyincrease in gas temperature will result in anincrease in the speed and torque developedby the turbines.
1-28-C-1. TURBINE OUTLETTEMPERATURE (TOT)
The temperature of the gases passingthrough the turbine is sensed by means offour thermocouples (Figure 1-12) locatedbetween the N1 and N2 turbine wheels. Eachthermocouple p roduces a vo l tageproportional to the gas temperature. Anaverage of the four voltages is displayed onthe cockpit TOT gauge.
1-28-D. POWER AND ACCESSORYGEARBOX
The main power and accessory drive geartrains (Figure 1-12) are enclosed in a singlegearcase. The gearcase serves as thestructural support of the engine. All enginecomponents are attached to the case. Thereare two independent drivetrains in thegearbox, gas producer (N1), and powerturbine (N2).
The gas producer (N1) gear train drives the N1tachometer generator, fuel pump, fuel control,starter generator, and oil pump.
The power turbine (N2) gear train drives the N2tachometer generator, power turb inegovernor, torquemeter, and freewheeling unit.
A two-stage helical and spur gear set is usedto reduce N2 rotational speed from 30,650RPM at the power turbine to 6016 RPM at the
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-59
output drive spline which mates with thefreewheel.
1-28-D-1. ENGINE TORQUEMETER
The torquemeter in the engine gearbox is ahydrau l ic type that uses the eng inelubrication system as its oil (hydraulic)pressure source. The output oil pressure fromthe engine torquemeter is displayed on thecockpit torque indicator. In order to minimizef r ic t ion e f fec ts and prov ide accuratemeasurement of torque, the axial gear thruston the helical torquemeter gear shaft is high.System pressure must always be greater thanthe torquemeter oil pressure. Therefore, it isnecessary to regulate the system oil pressureto the relatively high value of 115 to 130 PSI.
1-28-D-2. ENGINE ACCESSORIES
Informat ion on the var ious eng ineaccessories (Figure 1-12) follows:
The oil filter assembly includes an oil filter, oilpressure regulating valve, and an oil filterbypass valve.
The gas producer N1 tachometer/generatorproduces an electrical signal that is a functionof N1 gas producer turbine rotor RPM. Theoutput of th is tachometer generator isdelivered to the gas producer tachometerindicator and reflects gas producer turbine(N1) RPM.
The power turbine N2 tachometer/generatorproduces an electrical signal that is a functionof N2 power turbine rotor RPM. The output ofthis tachometer generator is delivered to thedual tachometer turbine/rotor indicator andreflects power turbine (N2) RPM.
The ignition exciter converts 28 VDC energyinto high temperature/high amperage arcs atthe igniter plug and is required during thestarting cycle. The igniter plug is threadedinto the combustion outer case. It extendsinto the combustion liner providing ignitionsparks that ignite the fuel/air mixture duringstart . Once the s tar t is complete , the
combustion is continuous and the ignitionsource is no longer required. Electrical poweris supplied from the electrical bus and isprotected by the 5-amp IGNTR circuit breaker.
The start counter (if installed) is energized bythe ignition system and is used to record thenumber of starts or attempted starts.
The oil pump supplies a pressurized volumeof oil for proper lubrication and cooling of thebearings and gears. A scavenge pumpcollects oil from the various sump cavitiesand returns it to the airframe mounted engineoil tank.
The fuel nozzle atomizes and injects fuel intothe combustion liner at the proper sprayangle and pattern.
The burner drain valves drain any unburnedfuel from the combustion section following anengine shutdown or aborted start. Duringstart, the drain valves close when the airpressure within the combustion sectionexceeds the air pressure on the outside of thecombustion section by a predeterminedvalue. The valves open on shutdown bymeans of spring action.
The starter/generator is used as a DC motor tospool the engine during the starting cycle.Once the start is completed, the DC generatorsupplies all the electrical needs of thehelicopter and charges the battery.
The engine-driven fuel pump assembly is asingle element gear type pump that producesapproximately 600 PSI. This high pressure isneeded to operate the fuel control system.The pump receives filtered fuel from theairframe mounted fuel filter, increases thepressure, and delivers the high pressure fuelinto the fuel control unit.
The gas producer fuel control and powerturbine governor serve as the fuel controlsystem to provide speed governing of thepower turbine rotor and overspeed protectionfor the gas producer turbine. The fuel controlsystem senses N1 and N2 RPM, compressor
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-60———Rev. 1—23 OCT 2008
discharge pressure, fuel control leverposition, and collective position to regulateand maintain fuel flow between establishedlimiting values. The system regulates theengine funct ions dur ing s ta r t ing ,acceleration, governing, deceleration, andshutdown.
1-28-D-3. ENGINE OIL SYSTEM
The engine incorporates a dry sump oilsystem with an airframe mounted supply tankand an oil cooler located on the top aft sectionof the fuselage (Figure 1-13). The oil filterassembly, consisting of an oil filter, filterbypass valve, and pressure regulating valve,is located on the top of the gearbox. Magneticchip detectors are installed at the bottom ofthe gearbox and at the engine oil outletconnection on the front of the gearbox. Allengine oil system lines and connections areinternal except the pressure and scavengelines to the front compressor bearing and tothe bearings in the gas producer and powerturbine supports.
The system is designed to furnish adequatelubrication, scavenging, and cooling asneeded to the bearings, splines, and gears,regardless of the helicopter attitude oraltitude. Jet lubrication is provided to allcompressor, gas producer turbine, and powerturbine rotor bearings, and to the bearingsand gear meshes of the power turbine geartrain, with the exception of the power outputshaft bearings. The power output shaftbearings and all other gears and bearings arelubricated by oil mist.
Oil from the tank is delivered to the engineinternal pressure pump. System oil pressureis regulated to 115 to 130 PSI by the pressureregulating valve. Pressurized oil passesthrough the filter and then to various points oflubrication. The scavenge pump returns theheated oil to the engine oil outlet port where itis routed to the oil cooler.
The oil cooler and blower assembly aremounted aft of the engine. The cooler bloweris a part of the tail rotor drive system andforces air upward through the cooler core.Return oil from the engine flows through theoil cooler bypass valve. The bypass valve willdirect oil away from the cooling fin ducts if theoil temperature is low. The valve is designedto regulate oil temperature between 71 and81°C (160 and 178°F).
The normal capacity of the engine oil tank is6.0 US quarts (5.7 L). The oil level is checkedby means of a dipstick mounted on the capand adapter assembly of the oil tank. The oiltank provides port openings for the supplytube, vent tube, scavenge tube, temperaturebulb, and drain valve.
Engine oil pressure is monitored by routingregulated oil pressure through an externalline to the engine oil pressure gauge in thecockpit. Oil temperature is monitored by theuse of a temperature bulb mounted in the oilline at the base of the oil tank. That signal isdelivered to the temperature side of theengine oil gauge and is protected by a 3-ampXMSN TEMP IND circuit breaker.
1-28-D-4. ENGINE FUEL AND CONTROLSYSTEM
The principal components of the engine fuelsystem are the engine driven fuel pump, gasproducer fue l cont ro l , power turb inegovernor, an accumulator, and a fuel nozzle(Figure 1-12).
The engine driven fuel pump assemblyincorporates a single gear type pumpingelement and a bypass pressure regulatingvalve. The pump produces high-pressure fuel(600 PSI) based upon N1 speed and sends it tothe fuel control for metering. The aft face ofthe pump provides a mounting pad for the gasproducer fuel control.
23 OCT 2008 Rev. 1 1-61
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-13. Engine Oil System206L4_MD_01_0032
SUPPLY FROM TANKAND PRESSURE PUMP
PRESSURE OIL
BYPASS OIL
SCAVENGE OIL
SCAVENGE RETURN TO TANK
TORQUE MODULATED PRESSURE
LUBRICATION SYSTEM SCHEMATIC (250-C30P)
TORQUEMETER
GAUGEPRESSURE
REGULATING
VALVE
OIL
FILTER
ASSEMBLY
CHECK
VALVE
FILTER
BYPASS
VALVE
FILTER
ELEMENT
OIL TEMPERATURE
AND PRESSURE
GAUGECOOLING AIR OUTLET
ACCESSORY
GEARBOX
HOUSING
PULLER CAP
AND DIP
STICK
OVERBOARD
VENT
OIL COOLER
CENTRIFUGAL
OIL BREATHER
TEMPBULB
PUMP
PRESSURE
OIL
NOZZLE
OIL
NOZZLE
OIL NOZZLE
PRESSURE
REDUCER
COOLER
BYPASS
VALVE
OIL TANK
AIR-OIL
SEPARATOR
GEAR
OIL COOLER
BLOWER
DRAIN
VALVE
MAGNETICCHIP
DETECTORPLUG
MAGNETICCHIP
DETECTORPLUG
CHIP DETECTORLIGHT
CHIPS
GEARBOXSCAVENGE
PUMP
THREESCAVENGE
PUMPS
GEARBOXHOUSING
COOLING
AIR INLET
1-62 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
The gas producer fuel control and the powerturbine governor provide for a fuel meteringsystem. This system senses gas producerRPM, power turbine RPM, compressordischarge pressure, and throttle position toregulate and maintain fuel flow. The gasproducer fuel control is the device that willmeter fuel into the engine to determine idlespeed, fuel flow for the initial combustionprocess, and start acceleration. When thethrottle is in the full open position, the fuelcontrol unit will control the fuel flow into theengine based upon the pneumatic signalreceived by the power turbine governor. Thefuel control unit has been mechanicallyadjusted (flat rated) to restrict fuel flow to amaximum of 356 pounds per hour to deliverapproximately 575 SHP.
An accumulator is located in the pneumaticline between the power turbine governor andthe gas producer fue l cont ro l . Theaccumulator dampens pressure signals fromthe power turbine governor to the fuel controldue to torsional vibrations encountered fromthe rotor system and maintains a constant N1RPM.
The fuel nozzle has a single entry and a dualoutlet orifice. This nozzle provides a finelyatomized spray of fuel at all flow conditionsthat are required by the engine. It is designedto provide an optimum spray angle forstarting the engine and an even distribution offuel into the combustion liner. The nozzle isequipped wi th a f i l ter to minimize thepossibility of contamination.
The fuel control system maintains enginepower output by regulating the gas producerspeed. Gas producer speed levels areestablished by the action of the power turbinefuel governor, which senses power turbinespeed. Power turbine speed (N2) is selectedby the pilot utilizing the BEEP switch on thecollective. The power required to sustain theselected RPM is maintained by power turbinegovernor action on metered fuel flow. Thepower turbine governor sends pneumaticsignals to the fuel control to increase or
decrease fuel f low to the combust ionchamber based on N2 RPM changes andmovement of the collective.
The engine controls consist of both the N1gas producer controls and the N2 droopcompensator controls. The gas producercontrols are operated by a twist grip on thecollective stick, and the droop compensatorcontrols are operated from a bellcrank in thecollective system.
The twist grip throttle on the collectivecontrols the gas producer fuel control andconsists of a flexible control cable whichextends from the throttle arm on the rear ofthe collective stick to a bellcrank assemblymounted on the engine deck. A control tube isconnected between the bellcrank and a levermounted on the fuel control shaft.
1-28-D-5. N2 DROOP COMPENSATORSYSTEM
206L4_MD_01_0024
206L4_MD_01_0025
23 OCT 2008 Rev. 1 1-63
MANUFACTURER’S DATA BHT-206L4-MD-1
The droop compensator control systemconsists of a mechanical linkage between anidler in the collective system and a levermounted on the power turbine governor shaft.Movement of the collective stick results in arepositioning of the governor shaft. Thisaction provides droop compensation toprevent RPM variations as power changes aremade. The system incorporates a linearactuator that is controlled electrically by aGOVernor RPM INCrease/DECRease switchmounted on the collective stick.
The droop compensator maintains N2 engineRPM as power demand is increased. It is adirect mechanical l inkage between thecollective stick and the speed selector leveron the N2 governor and will maintain N2 RPMwhen properly rigged.
Droop is defined as the speed change in N2engine RPM as power is increased from a noload condition. It is a characteristic designedinto the governor system to prevent instabilityf rom deve lop ing as eng ine ou tput isincreased. Without this characteristic, N1speed would overshoot or hunt the valuenecessary to satisfy the new power condition.I f N 2 is a l lowed to d roop o ther thanmomentarily, the reduction in rotor speedcould become critical.
1-29. RETIREMENT INDEX NUMBER(RIN)
Each component with a ret i rement l i fesensitive to torque events will be assigned amaximum Retirement Index Number (RIN).This RIN corresponds to the maximum allowedfatigue damage resulting from lifts andtakeoffs. A new component will begin with anaccumulated RIN of zero that will be increasedas lifts and takeoffs are performed. Theoperator will record the number of lifts andtakeoffs and increase the accumulated RIN.When the maximum RIN is reached, thecomponent will be removed from service.Certain components may be assigned a life inhours in addition to a RIN.
Pilots are to record torque events for eachflight. Maintenance personnel will convert thetorque events into RIN to track the lives of allthe required components.
A torque event is defined as a takeoff (onetakeoff plus the subsequent landing = 1 RIN)or a lift (internal or external). For example, if anoperator performs six takeoffs and 10 slingloads, this would total 16 torque events (6takeoffs = 6 events, 10 sling loads = 10 events,6 + 10 = 16 events total).
1-30. DRIVETRAIN SYSTEM
The drivetrain system (Figure 1-14) provides ameans of transmitting power from the engineto the main and tail rotor assemblies. Thedrivetrain includes the freewheel assembly,main driveshaft, transmission, mast, tail rotordriveshaft, and the tail rotor gearbox. The rotorsystems include the swashplate assembly, themain rotor hub and blades and the tail rotorhub and blades, and the tail rotor pitch changemechanism.
1-30-A. FREEWHEEL ASSEMBLY
The freewheel assembly (Figure 1-14 andFigure 1-15) is mounted on the engine gearboxand driven under power from the enginepower takeoff (PTO) gear shaft. Engine poweris transmitted to the outer shaft of thefreewheeling unit, through the engaged spragclutch, and then to the freewheeling innershaft . The f reewheel ing inner shaft isconnected to the main input driveshaft on theforward adapter flange. The tail rotor drivesystem is driven through a flexible couplingand a splined adapter mounted on the aft endof the freewheel inner shaft.
Dur ing autorota t ion , the sprag c lu tchdisengages and the rotational forces of themain ro tor are u t i l i zed to dr ive thetransmission accessories and tail rotor drivesystem.
1-64 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
Figure 1-14. Drivetrain System
1
3
2
4 5 6 7
9
1011
13 12
14
15
16
17
25
26
27
29
30
31
32
32
33
34 38
3637
35
39
40
8
18
20
21
22
23
24
19
206L4_MD_01_0038
28
Transmission restraintUp stopStop mountLink attachmentHydraulic pump and NR tachometer generatorLower chip detectorUpper chip detectorNodal beam assemblies and mountsDrag pinLink attachmentMain input driveshaftRotor brake assembly (optional)Spacer
The main driveshaft (Figure 1-14) is designedto transfer power from the engine to thetransmission during normal operations and totransfer drive from the transmission to the tailrotor system during autorotation. The maindriveshaft rotates at 6016 RPM at 100% N2.
The helicopter uses the Kaflex driveshaft,which provides a flexible drive couplingbe tween the f reewheel ing uni t andtransmission. This flexing allows for thesmooth t ransfe r o f torque when thetransmission and freewheeling unit are notaligned. The driveshaft is made up of threerectangular plate sets in each coupling. Eachcoupling can be considered a truss work inwhich torque loads are carried as axial loadsthrough the straight members of each plate.
Preflight inspection consists of checking theflex frames for cracks and for condition andsecurity.
1-30-C. TRANSMISSION
The transmission (Figure 1-14), located on thecabin roof, incorporates a nodal beamsuspension and restraint. It is mounted ontwo nodes, and a flexure assembly absorbsnormal vibrations and vertical movement.Elastomeric bearings are used to minimizevibrations being transferred into the airframe.
The transmission reduces engine RPM,changes the angle of drive from the engine tothe rotor, and serves as the structuralmember that links the mast to the airframe. Itprovides a two-stage reduction of 15.23 to 1.0(6016 to 395 RPM). The first stage is a bevelgear arrangement with a 3.26 to 1.0 reduction;the second stage reduction is obtainedthrough a planetary gear train that provides a4.67 to 1.0 reduction.
The transmission assembly consists of a topsupport case and a lower case. This lowercase contains the input pinion, bevel gear
arrangement, planetary gear train, andaccessory gear drive. Components attachedto the transmission assembly are the maindr iveshaft , t ransmiss ion o i l pump,transmission oil filter housing, hydraulicpump, rotor RPM tachometer generator, andtwo magnet ic ch ip detectors . Thetransmission system is equipped with atemperature switch, temperature bulb,pressure switch, and wet line pressure gaugeto provide system information.
1-30-D. TRANSMISSION OIL SYSTEM
The transmission oil system (Figure 1-16)includes a pump, a pressure regulating valve,an oil cooler, a filter element, and two sprayjets. The pump is a constant volume typedriven by the accessory gear, which deliverspressurized oil externally to the filter andhousing assembly, cooler, and then returns itto the main transmission where spray jetslubricate the internal parts. A sight glass isloca ted on the r ight s ide o f the maintransmission lower case where the oil levelcan be easily checked.
A non-vented filter cap is located on thetransmission top case. The oil system alsoprovides lubrication for the freewheelingassembly mounted in the engine accessorygearcase. Pressurized oil is delivered fromthe transmission into a T-fitting on the cabinroof. Oil flowing out of the left side of thefitting is reduced in pressure and is utilized tolubricate the freewheeling unit. Oil flowing outof the r ight side of the f it t ing del iversregulated pressure into the transmission oilpressure gauge and into the oil pressureswitch for the TRANS OIL PRESS cautionlight.
The oil pump, rotor tachometer generator, andhydraulic pump are mounted together on theforward s ide of the t ransmission. Theaccessory d r ive gear dr ives thesecomponents using a 1.42 to 1 ratio.
1 To pilot oil pressure gauge and low pressure warning switch.
NOTE
1
PLANETARY ASSEMBLY
OIL PUMP
4.5 TO 5 GPM
(17 TO 19 IPM)
DRAIN PLUG AND MAGNETIC
CHIP DETECTOR
INTERNAL JETS TO SUN GEAR
AND PLANETARY
OIL JET NO. 1
OIL JET NO. 2
BLOWER ASSEMBLY
MAST ASSEMBLY
DRAIN PLUG AND MAGNETIC
CHIP DETECTOR
PRESSURE
REG VALVE
FILTER
FREEWHEEL UNIT
ENGINE
FILTER ELEMENT
THERMOBULB
HIGH TEMP WARNING SW
MAST BEARING
MAGNETIC CHIP
DETECTOR
CLOG FILTER INDICATOR
FILTER BYPASS VALVE
OIL COOLER THERMO
BYPASS VALVE
OIL COOLER
RESTRICTORS IN FITTING
1-68 Rev. 1 23 OCT 2008
BHT-206L4-MD-1 MANUFACTURER’S DATA
The pump output is between 6.0 to 6.7 gallonsper minute (22.7 to 25.4 L/min). The pressureis regulated by the transmission oil pressureregulator valve and is set at 50 to 55 PSI. Thepump scavenges oil from the lower casesump through a wire screen and the lowerchip detector. Oil is then directed to thetransmission oil filter assembly.
The transmission oil pressure regulatingvalve, which is located on the left rear cornerof the lower case, is used to adjust the oilpressure to normal operating limits and torelieve excess oil pressure back into thetransmission case.
The transmission oil f il ter and housingassembly, which is mounted on the aft side ofthe transmission, has a filter element, bypassind icato r, bypass va lve , thermosta t ,temperature bulb, and a temperature switch.The filter element, which is contained in thelower housing of the oil filter assembly, is adisposable, pleated paper element.
The oil fi lter bypass indicator, which islocated on the filter housing above the filterelement, incorporates a spring-loaded magnetthat is ported to the bypass valve. If thebypass valve opens due to a clogged filterelement or extreme cold temperatures, theindicator button will extend to indicate animpending bypass. The bypass va lveassembly will open when excessive backpressure is caused by either a clogged filterelement or extreme cold temperatures. It willthen allow oil to bypass the filter element andto be directed to the thermostat and oil cooler.The temperature bulb is located on theforward side of the filter housing and iselectrically connected to the oil temperatureindicator in the instrument panel.
The transmission oil temperature switch islocated on the oil f i l ter housing and iselectrically connected to the TRANS OILTEMP caution light segment. When the oiltemperature reaches 110°C, the temperatureswitch will close and illuminate the cautionlight segment. The TRANS OIL PRESS caution
light segment switch is located below theinstrument panel and is connected into thetransmission oil pressure indicator oil line.During normal operation, the oil pressureswi tch opens a t maximum 36 PSI onincreasing oil pressure and closes at 30±2 PSI on decreasing pressure.
The oil pump inlet screen assembly, which islocated in a boss adjacent to the transmissionoil pump, filters oil prior to its entering the oilpump.
Two o i l je ts are incorporated in thetransmission. The No. 1 oil jet directs alubricating oil spray to the transmission bevelgears, and the No. 2 oil jet lubricates theplanetary pinions and mast bearing.
The three electric chip detectors (two on thetransmission and one on the freewheelingunit) consist of a self-locking, bayonet typeprobe with a permanent ceramic magnet atthe end. In the event that metal particlesshould become free in the oil, the magnet willat tract them. When suff ic ient metal isattracted to complete a circuit between thepole and ground, the TRANS CHIP segmenton the caution panel will illuminate.
The o i l cooler assembly conta ins twoseparate and independent cores weldedtogether. The aft core is for the engine oilsystem and the forward core is for thetransmission oil system. The cooler assemblyis connected to the o i l cooler b lowerassembly by means of a transition duct. Thethermal bypass valve for the transmission oilcooler is located on the transmission oil filterhousing and will begin to regulate oil to thecooler when transmission oil reaches atemperature of approximately 65.6°C (150°F).The thermal bypass wi l l d i rec t a l lt ransmiss ion o i l to the coo ler a t o i ltemperatures of 81.1°C (178°F).
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-69
The oil cooler blower is mounted aft of theengine firewall on the helicopter structure andsupplies air to the engine and transmissionoil cooler. The cooler blower shaft is mountedon two sealed bearings and is connected tothe forward and the aft short shafts. It passesthrough the blower assembly housing, andthe blower impeller is bolted to the shaft. Thisshaft is part of the tail rotor driveshaft systemand is used to power the blower.
During normal operation, oil flows from thetransmission to the cooler and back to thetransmission. Two transmission oil coolerdrain valves (pressure line/return line) arelocated inside the baggage compartment toallow for the draining of the transmission oilcooler when transmission oil is changed.
1-30-E. TRANSMISSION MOUNTS
The transmission, mast, main rotor hub, andthe blade assembly are supported by andisolated from the fuselage by means of thetransmission mounts (nodal beam) andtransmission restraint (Figure 1-14). This isaccomplished through the use of four linkattachments and two stop mounts bolted tothe transmission. Attached to these are fourlink assemblies, which are secured to the foursupports and the flexure assemblies. Thesupport assemblies are bolted to the cabinroof shell and cabin roof beam. The supportassemblies contain elastomeric bearing andbushing members to isolate and balancevibration inputs from the rotor into the flexureassemblies. These assemblies are theprimary vibration-isolating component thatdampens the 2/rev vibration from the rotorsystem from being transmitted into the cabin,providing for a vibration free ride throughoutthe he l icop ter 's speed range . Thetransmission restraint restricts movement ofthe transmission through the use of anelastomeric bearing and bushing membersecured to the transmission restraint support.This support is then bolted to the cabin roofshell and cabin roof beam along with two dragpins. The two stop mounts bolted to thetransmission are installed over the drag pins.
The up stops mounted on the drag pinsrest r ic t the ver t ica l movement o f thetransmission and all oscillatory movement.Tuning weights, which are bolted to the armand flexure assemblies, provide for thefine-tuning and balance of the transmissionmounting system.
The arm and flexure assembly is the primarynodal beam component that isolates mainrotor and transmission vibrations from thefuselage. The nodal beam instal lat iondevelops a standing wave form to reducefuse lage v ib ra t ions throughout thehelicopter's speed range. The arm assembliesare manufactured from aluminum alloy andare bushing-mounted to the fore and aftsupport assemblies. The flexure assembly,which is contained between the two armassemblies, consists of a laminated elastomerflexure, shims, and shear pads bonded intometal brackets. Tellurium lead tuning weights,which are located on the end of the armassemblies adjacent to the flexure, provide amethod for tuning the nodal beam installation.
The forward and aft support assemblies,which are manufactured from aluminum alloy,contain bonded elastomeric bearing andbushing members to isolate the arm andflexure assemblies from the fuselage.
1-30-F. TRANSMISSION RESTRAINT
The transmission restraint (Figure 1-14) is anelastomeric spring assembly enclosed in asteel housing. The restraint is secured in anisolation support that is mounted to the cabinroof beam. This restraint restricts forward, aft,and lateral transmission movement. Theforward arms of the rest ra in t conta inteflon-l ined spherical bearings that areattached to the stop mounts, which in turn aresecured to the t ransmiss ion . Thetransmission is further restrained through thestop mounts by means of the up stops anddrag pins.
The aluminum alloy isolation support issecured to the upper cabin roof shell and
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-70———Rev. 1—23 OCT 2008
cabin roof beam. It provides a fixed structuralsupport for the transmission restraint.
The two s tee l stop mounts, which areattached to the lower aft portion of thet ransmiss ion case , l im i t ve r t ica l andoscillatory movement of the transmission.
The two steel drag pins limit the oscillatorymovement of the transmission through thestop mounts. Vertical restraint is limited bythe up stops, which consist of a bondedupper and lower disc in contact with the stopmounts.
1-30-G. MAIN ROTOR MAST
The main rotor mast (Figure 1-14) is attachedto the transmission by means of a mastlocking plate and studs in the top of thetransmission case. The mast is a hollow, steelshaft that transmits rotational energy from thetransmission to the rotor system. It has threesets of splines and two threaded areasincorporated into its design. The upper splineis master splined to correctly position themain rotor trunnion. The swashplate drivesplines, located in the mid portion of themast, are machined to receive a collar set towhich the swashplate drive link is attached.The lower splines are the drive splines wherethe rotational energy from the transmission istransmitted to the mast from the planetarygearing. The upper portion of the mast isthreaded to receive the mast retaining nut.
1-31. MAIN ROTOR SYSTEM
The main rotor system (Figure 1-14) is a semirigid, underslung hub, feathering axis rotorwith two metal blades. The hub assembly isattached to the mast by means of a splinedtrunnion that also functions as the rotorflapping axis. The flapping bearings aregrease lubr icated rol ler bear ings thatassemble into the pillow block housing that isattached to the yoke and also serves toprovide hub centering provisions togetherwith the trunnion. The yoke serves as thepitch change axis, and the grips are attached
to the yoke by means of the tension-torsion(T-T) strap assembl ies which t ransfercentrifugal loading from the blades to theyoke and a lso ass is t to counte ractaerodynamic forces. The pitch changebearings are housed in the grips and aregrease lubricated. The change in blade pitchangle is accomplished by rotating the gripson the yoke with an input at the pitch horns.The blades are attached to the grips with boltswhich have hollow shanks that are used toinstall weight for static and dynamic balanceof the hub and blade assembly. Bladealignment is accomplished by adjustment ofthe blade latches which engage the root endof the blade.
The trunnion mates the rotor system to themast, transmits drive to the yoke assembly,and serves as a pivot point for the rotorflapping axis.
The yoke supports the main rotor system. Ithas two hollow and preconed journals torelieve stresses. Each journal has the hollowbore especially machined at the outboard endto fit a seal assembly, and the inboard end ismachined to fit the tension-torsion (T-T) strapend fitting.
The T-T strap, which consists of stainlesssteel wire wrapped around two steel ends,provides a connection between the grips andyoke. During rotor feathering, the flexibility ofthe strap allows the grip bearings to rotate onthe yoke journals.
The grip assembly is the structural memberbetween the yoke and the main rotor blades.The outboard end of each gr ip has amachined hole for the blade retention bolt,and the grip body has drilled tangs to fit theblade latch bolt that retains the outboard endof the T-T strap.
1-31-A. MAIN ROTOR BLADES
The main rotor blades (Figure 1-14) havequalities of advanced aerodynamic efficiency,ruggedness, and simplicity. The blades have a
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-71
-11° twist and an airfoil characterized by asignificant positive camber to the first 1/3 ofthe chord length and by the flat symmetricalsurfaces of the aft 2/3. Through the use of adroop-snoot airfoil section, rotor bladeefficiency is significantly improved comparedto symmetrical blade shapes. A wide thrustmarg in for maneuverabi l i ty and h ighefficiency in a hover are, in part, the result ofthis airfoil design. It is an all metal bondedassembly consisting of three structuralmembers: an aluminum spar, spar closure,and a trailing edge strip. Skins, stabilized byhoneycomb core, are bonded to the majorsection by adhesive applied under heat andpressure. Reinforcing doublers, grip plates,and drag plates are bonded to the blade buttend. The basic portion of the blade has a13.026 inch (33.09 cm) chord.
The flap restraint is installed on the main rotorhub t runn ion and incorpora tescounterweights and springs which serve tolimit flapping freedom during starting andshutdown, but permit normal flapping atoperating RPM.
1-32. TAIL ROTOR DRIVE
The tail rotor driveshaft (Figure 1-14) is madeup of the forward short shaft, the oil coolerblower shaft, the aft short shaft, and the tailrotor driveshaft segments. Steel laminatedflexible couplings (Thomas couplings) areused to connect the shaft sections and the tailrotor gearbox.
The forward and aft short shafts are locatedon either s ide of the oi l cooler blowerassembly. The fo rward shor t sha ft isconstructed of steel and is connected to theaft end of the freewheeling assembly and theforward end of the fan shaft by means ofsplined adapters. The aft short shaft isconstructed of a luminum al loy and isconnected to the aft end of the fan shaft by
means of a splined adapter and the first tailrotor driveshaft segment.
The segmented tail rotor driveshaft consistsof five segments that extend along the top ofthe tailboom. Each segment of the drivesystem is identical and interchangeable withthe others.
1-32-A. TAIL ROTOR DRIVESHAFTHANGER BEARINGS
Seven hanger bear ing assembl ies(Figure 1-14) support the driveshafts andflexible, steel, disc couplings. Thomascouplings are used to both connect the shaftsections and allow continued alignment withthe tailboom.
1-32-B. TAIL ROTOR GEARBOX
The tail rotor gearbox (Figure 1-14), locatedon the aft end of the tailboom, drives the leftside mounted tail rotor. It contains two spiralbevel gears positioned at 90° angles to theother. The direction of drive is changed 90°and there is a speed reduction of 2.35 to 1.0 atthe gearbox to achieve an output shaft speedof 2550 RPM.
The magnesium housing is attached to thetailboom by means of four bolts and twoalignment pins. The assembly includes abreather type filler cap, oil level sight gauge,and a combination electrical chip detectorand self-closing valve.
The chip detector consists of a self-lockingbayonet probe with a permanent magnet atthe end. Free ferrous metal particles in the oilwill be attracted to the magnet, and whensufficient metal is attracted to complete thecircuit between pole and ground, the T/R CHIPdetector segment on the caution panel willilluminate. The valve automatically closes andprevents loss of oil when the electric chipdetector is removed for inspection. It alsoserves as a drain plug.
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-72———Rev. 1—23 OCT 2008
1-33. TAIL ROTOR HUB AND BLADEASSEMBLY
The tail rotor assembly (Figure 1-14) is a deltahinge type that consists of a hub and twointerchangeable blades. The yoke of the hubis made with a 4° twist for each blade.
The aluminum alloy forged yoke is attached tothe tail rotor gearbox shaft by means of asplined trunnion inside the yoke to provide aflapping axis for the assembly. At the time ofassembly, spanwise balance is accomplishedby using washers on the blade bolts at theyoke, and chordwise balance is accomplishedby using weights and washers on the trunnionbearing housing restraining bolts.
The tail rotor blades are all metal assembliesconsisting of a stainless steel shell reinforcedby a honeycomb filler and stainless steelleading edge abrasive strip. Ballast stations,located at the inboard trailing edge and the tipof the blades, are provided for mass balanceof the blades. Weights used in these locationsare dete rmined when the b lade ismanufactured.
The blades are attached to the yoke by meansof spherical bearings (this provides for pitchchange of the blades) that are mounted in thegrip plates of the blades on the pitch changeaxis.
Connecting linkage consists of push-pulltubes, bellcranks, levers, and supports thatconnect the pilot and copilot (if installed) tailrotor control pedals to the tail rotor pitchchange mechanism.
Tail rotor pitch control is accomplished bymeans of bellcrank, rod, and lever assemblymounted on the tail rotor gearbox whichactuates a control tube through the hollowoutput driveshaft to the crosshead and pitchlinks.
1-34. MAIN ROTOR FLIGHTCONTROLS
The flight controls (Figure 1-17 throughFigure 1-20) are mechanical linkages that areactuated by conventional controls and usedto control flight attitude and direction. Boththe cyclic (longitudinal and lateral) and thecollective controls incorporate hydraulicservo actuators. A synchronized elevator islinked into the fore and aft controls at thecyclic torque tube.
The flight controls are routed beneath theforward seats, aft to the vertical controlcolumn, then up to the cabin roof. The verticalcontrol column also serves as a primary cabinstructure. Access doors on the aft side of thecontrol column and seat panels are providedfor inspection of control components andmaintenance accessibility.
Dual controls are installed as an option toprovide dual flight control capability either foroperations requiring a pilot and a copilot orfor pilot training operations. Installation ofdual controls provides a collective stick,cyclic stick, and a tail rotor control pedalassembly for the copilot. The pilot and copilotcontrols are similar in appearance andrelative position and the control input to therotor system is the same. The control feel forthe copilot controls is the same as that for thepilot controls. The copilot controls areconnected to the pilot controls by means of ajackshaft, control tubes, and electrical wiring.Quick disconnects are provided for thecopilot collective and cyclic.
Aluminum alloy control tubes are usedthroughout the collective, cyclic, and tail rotorcontrol system. Some control tubes are fixedin length with bonded end fittings, whileothers may have ad justab le f i t t ings .Bellcranks, levers, and supports are usedthroughout the collective, cyclic, and tail rotorcontrol systems. These parts transmit orcontrol change movements in the particularsystem in which they are installed.
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-73
1-34-A. CYCLIC FLIGHT CONTROLS
The cyclic flight controls (Figure 1-17) consistof a control stick, torque tube, yoke, hydraulicservo actua tors , con tro l tubes , andbellcranks. Movement of the control stick istransmitted through linkage and hydraulicservo actuators to the swashplate, whichactuates the rotating controls to the mainrotor. Servo actuators are incorporated tominimize the effort required to move thecontrols and to reduce main rotor feedbackforces. The cyclic stick extends upward andforward from the front of the pilot seat. Thetorque tube connects to the cyclic sticksupport and provides a mounting point for theelevator controls. The cyclic system yokeextends aft from the cyclic stick support andjackshaft. Movement of the cyclic stick istransmitted to the mixing lever by means ofthe yoke.
This mixing lever transmits cyclic movementto the swashplate through mechanical linkageand servo actuators.
A balance spring is used in the system tominimize the cyclic stick mass imbalanceforces and elevator induced forces in thelongitudinal control system.
For helicopters with dual controls, the copilotcyclic stick is installed in front of the copilotseat with all VFR control functions. A quickdisconnect feature permits rapid removal ofthe stick. I f the electr ical connector isdisconnected, an adapter must be installed toprovide circuit continuity. A spring pinassembly is provided to ensure positiveengagement of the stick.
1-34-B. COLLECTIVE FLIGHTCONTROLS
The collective flight controls (Figure 1-18)consist of a collective stick, jackshaft, controltubes, bellcranks, and a hydraulic servoactuator. Movement of the collective stick istransmitted by means of linkage and the
servo actuator to the swashplate collectivelever. Collective pitch control is transmitted tothe main rotor controls by vertical movementof the swashplate. The servo actuator ismounted on a support, which is located onthe cabin roof d i rect ly forward of thetransmission, along with two servo actuatorsfor the cyclic system.
The collective stick is installed to the left ofthe pilot seat, and it extends upward andforward through a f lexib le cover. Thecollective jackshaft provides a mounting pointfor the collective stick. An adjustable frictionbearing mounted on the jackshaft allows thepi lot to ad just the f r ic t ion to h is ownrequirements. A minimum friction adjustmentclamp located at the left end of the jackshaftensures that the collective stick will alwayshave a preset minimum friction to eliminatecollective bounce.
The collective trunnion and lever is installedbetween the collective jackshaft and controltube. It ties the collective controls to themixing lever of the cyclic controls. When thecollective stick is moved to change rotorpitch, the cyclic control servo actuators andlinkage will move to maintain the swashplatein its relative plane.
For helicopters with dual controls, the copilotcollective stick is installed at the left of thecopilot seat with a fully functioning throttlecontrol. A quick disconnect feature permitsrapid removal of the copilot collective stick. Aspring pin assembly is provided to ensurepositive engagement of the stick.
1-34-C. ELEVATOR FLIGHT CONTROLS
The elevator flight controls (Figure 1-19)consist of a horizontal stabilizer, elevator,supports, bellcranks, walking beam, andcontrol tubes. To safeguard against elevatoror cyclic controls jamming, the control tubeincludes a shear joint in the interconnectlinkage from the fore and aft cyclic controltorque tube to the elevator horn.
Collective jackshaftFriction knobCollective stickTrunnion and leverControl tubeSupportCollective servo actuatorYokeRight cyclic control tubeElevator control tubeLeft cyclic control tubeIdlerControl tubeCollective bellcrankLinkCollective lever
7
15
16
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-76———Rev. 1—23 OCT 2008
Figure 1-19. Elevator Controls
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-77
The horizontal stabilizer/elevator is mountedon the tailboom. The horizontal stabilizersection is stationary while the elevator, whichis attached to the aft side of the stabilizer, issynchron ized to fo re and a ft cyc l icapplication. The elevator horn is connected tothe cyclic torque tube through the elevatorcontrol linkage.
The elevator control linkage consists ofcontrol tubes, bellcranks, and walking beams.The forward control tube, which contains ashear pin, is attached to an eyebolt mountedin the cyclic torque tube. The aft control tubeis attached to the elevator horn.
Should the pin shear, the rod has sufficientlength to slide within the tube to preventseparation that could result in jamming of thecyclic controls.
1-34-D. TAIL ROTOR DIRECTIONALCONTROLS
The tail rotor directional controls (Figure 1-20)include the control pedal assembly, pedaladjuster, control tubes, bellcranks, damperassembly, and a pitch control mechanismmounted through the tail rotor gearbox shaft.Moving the pedals causes pitch change in thetail rotor blades to offset the main rotortorque and to control the directional headingof the helicopter. The tail rotor control pedalsmounted on the pilot's compartment deck areconnected under the center console to abellcrank pedal adjuster, which provides formanual ad justment o f pedal pos i t ionaccording to the pilot's needs. Alternatepedals may also be installed which allow thepilot to manually adjust the position of thepedal foot rests.
For helicopters with dual controls, the copilotfully functional tail rotor control pedalassembly is installed on the floor in front ofthe copilot seat to provide a means for thecopilot to control the tail rotor assembly. Thecontrol pedals are linked to the pilot pedals by
means of control tubes and a bellcrank. Thecopilot pedals can be positioned, as desired,by means of the pedal adjuster. Alternatepedals may also be installed which allow thecopilot to manually adjust the position of thepedal foot rests. Connecting linkage consistsof push-pull tubes, bellcranks, levers, andsupports that connect the pilot tail rotorcontrol pedals to the tail rotor pitch changemechanism.
Tail rotor pitch control is accomplished bymeans of bellcrank, rod, and lever assemblymounted on the tail rotor gearbox whichactuates a control tube through the hollowrotor driveshaft to the crosshead and pitchlinks.
1-34-E. VERTICAL CONTROL COLUMN
The flight controls are routed beneath thepilot and passenger seats aft to the verticalcontrol column then up to the cabin roof. Thiscontrol column also serves as a primary cabinsupport structure. Access panels on the afts ide of the column, the bottom of thehelicopter, and seat panels are provided forinspection of control components andmaintenance accessibility. The cyclic controlsare mixed with collective control through themixing lever bellcrank located at the base ofthe control column.
1-34-F. SWASHPLATE AND COLLECTIVELEVER
The main rotor controls consist of theswashpla te and suppor t assembly(Figure 1-21), drive link, and pitch links. Theswashplate transfers cyclic control motionsfrom the non-rotating to the rotating controlsystem. The swashplate and support encirclethe mast directly above the transmission. Theswashplate mounts on a universal support(pivot sleeve and uniball) that permits it tomove in any direction. Movement of the cyclicresul ts in a correspond ing t i l t o f theswashplate and the main rotor.
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-78———Rev. 1—23 OCT 2008
Figure 1-20. Tail Rotor Controls
206L4_MD_01_0039
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18.Adjustable control tubes
Pedals
Pedal adjuster
Control tube
Control tube
Control tube
Counter weight
Bellcrank
Control tube
Bellcrank
Control tube
Walking beam
Control tube
Walking beam
Control tube
Bellcrank
Rod assembly
Tail rotor hub and blade
Pitch change mechanism
6
ALTERNATE
PEDALS
MANUFACTURER’S DATA BHT-206L4-MD-1
23 OCT 2008—Rev. 1———1-79
Figure 1-21. Swashplate and Support206L4_MD_01_0040
NutWasherSupportCollective leverSpherical bearingInner ringOuter ringBootCollar set
BHT-206L4-MD-1 MANUFACTURER’S DATA
1-80———Rev. 1—23 OCT 2008
The collective lever and link assembly ismounted to the swashplate support assemblyand transfers collective inputs to the lowerswashplate. Movement of the collective pitchlever actuates the collective sleeve assemblythat in turn raises or lowers the swashplateand transmits collective input to the mainrotor.
The swashplate drive assembly consists of acollar set, idler drive link, and idler lever. Thecollar set is attached to the mast and the idlerdrive link is attached to the outer ring of theswashpla te . Th is connects the upperswashplate to the mast, causing it to rotatewith the mast.
The pitch link assemblies connect the pitchhorns on the blade grips to the swashplatewhich transmits control input from both thecollective and cyclic controls.
1-34-G. DUAL CONTROLS
Installation of dual controls provides acollective, cyclic, and a tail rotor control pedalassembly for the copilot. The copilot controlsare connected to the pilot controls by meansof the jackshaft, control tubes, and electricalwiring. Quick disconnects are provided forthe collective and cyclic. The copilot controlsdo not provide electrical cargo release, flightidle stop, throttle bezel marking, starterswitch, or landing light controls.
1-35. HYDRAULIC SYSTEM
The hydraulic system (Figure 1-22) providespressurized fluid to operate the cyclic andcollective flight control servo actuators.Operation of the system is electrical lycontrolled by means of the HYDRAULICSYSTEM switch on the miscellaneous controlpanel. When the hydraulic system solenoidvalve is de-energized (switch ON), the
pressurized hydraulic fluid flows to the threeservo actuators (Figure 1-23). When it isenergized (switch OFF), the pressurizedhydraul ic f luid is d irected back to thereservoir and bypasses the three servoactuators. In case of a total electrical failure,the system is fail-safe ON (no electrical power= hydraulics ON).
The cyclic and collective servo actuatorsupport is installed on the cabin roof forwardof the transmission. It serves as a mount forthe three servo actuators and associatedbellcranks. The collective servo actuator ismounted in the center position and the twocyclic servo actuators are mounted on theoutboard positions.
The hydraul ic reservoir and cover areconstructed from magnesium alloy. Thereservoir is mounted on a brace and supportforward of the main transmission.
The hydraulic system has no specific coolingsystem. Heat is dissipated from the hoses andlines as air circulates beneath the forwardcowling (radiation cooling to the atmosphere).
1-35-A. HYDRAULIC SYSTEMCOMPONENTS
The hydrau l ic system components(Figure 1-22) include the hydraulic pump,hydraulic solenoid valve, hydraulic filters, andpressure relief valve.
1-35-A-1. HYDRAULIC PUMP
The hydraulic pump (Figure 1-22) is mountedon the transmission lower case and driven bythe transmission accessory drive. The pumpis a constant pressure, variable delivery,self-lubricated type designed to operatecontinuously and provide a rated dischargepressure of 1000 ±25 PSI.
The hydraulic solenoid valve (Figure 1-22) ismounted on the cabin roof forward of thehydraulic pump and reservoir. The valve iselectrically closed and is controlled by theHYDRAULIC SYSTEM swi tch on themiscel laneous panel (F igure 1-4) . Theelectrical circuit is protected by a 5-ampHYDR SYSTEM c i rcu i t breaker on theoverhead console. Because the valve musthave electrical power to close, the system isreferred to as fail-safe ON. When the switch isturned off, hydraulic fluid bypasses theservos and returns to the hydraulic reservoir.
1-35-A-3. HYDRAULIC FILTERS
The hydraulic filters (Figure 1-22) are installedon a bracket on the forward right side of thecabin roof near the main transmission. Eachfilter assembly contains a filter indicator thatindicates an impending clogged filter. Theindicator consists of a red button mounted onthe f i l ter assembly housing. When the
differential pressure across the filter is 70 ±10PSI, the red button will rise. To preventinaccurate indications of clogging, theindicator will not work when the hydraulicfluid temperature is less than 2°C (35°F). Thereturn line filter incorporates a bypass valve.If the filter element in the return line becomesclogged, the bypass valve will open and allowfluid to bypass the filter and return to thereservoir.
1-35-A-4. PRESSURE RELIEF VALVE
The pressure relief valve (Figure 1-22) isinstalled forward of the transmission and onthe left side of the cyclic and collective servoactuator support. During normal operations,the relief valve is in the closed position. Ifsystem pressure should increase to 1075 to1375 PSI, the valve will open to preventdamage to the system by returning excesspressure into the return side of the solenoidvalve. The valve will close when pressurereturns to normal.
MANUFACTURER’S DATA BHT-206L4-MD-1
Figure 1-23. Hydraulic System Schematic206L4_MD_01_0042_c1
RETURN
SEQUENCEVALVE CHECK
VALVES
INPUT FROMFLIGHT CONTROLS
CYLINDEROUTPUT
SERVO ACTUATOR - TYPICAL
SEE DETAIL B
PRESSURERETURNSUCTION
PRESSURERETURN
TEST PORT
DIFFERENTIAL RELIEF VALVE
CYCLIC CYCLIC COLLECTIVE
SEE DETAIL B
SEE DETAIL B
SEE
DETAIL B
SEE DETAIL A
SERVOACTUATOR
PRESSURE
RETURN
ENERGIZED-SYSTEM OFF
DE-ENERGIZED-SYSTEM ON
SOLENOID VALVE SCHEMATIC
DETAIL A
SERVOACTUATOR
PRESSURE
SYSTEM CAPACITY 65 CUBIC INCHESRESERVOIR CAPACITY 40.0 CUBIC INCHESRESERVE CAPACITY 24.3 CUBIC INCHESOPERATING PRESSURE 1000 PSIOPERATING TEMPERATURE -65 TO 160°F (-54 TO 71°C)HYDRAULIC FLUID MIL-H-5606
