332 C1e 11.102E01

TECHNICAL SPECIFICATION

NOTE The update status of your copy of the Technical Specification is ascertained by the reference at the bottom of this page. The pictures, front faces of the control boxes, panels and displays included in this document are only given for information.

BP 107- 10 , AVENUE MARCEL CACHIN 93123 LA COURNEUV E CEDEX, FRANCE TEL : 00.01.49.34.45.00 - TELEX : EUROCOPTER 231 268 F TELEFAX: 00.01.49.34.45.30

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This document is the property of EUROCOPTER ; no part of it shall be reproduced or transmitted without the express prior written authorization of EUROCOPTER and its contents shall not be disclosed. © EUROCOPTER 2013

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Technical Specification

CONTENTS

1.

Forewor For ewor d ....................... .................................. ....................... ....................... ....................... ....................... ....................... ....................... ..................... .......... 1-1 1.1.

2.

Foreword .......................................................................................................................... 1-1

Character Charac ter isti is ti cs ....................... .................................. ...................... ....................... ....................... ...................... ....................... ....................... ............. 2-1 2.1.

Standard lay-out ........................................................................................................... 2.1-1

2.2.

Weight Weigh t b reakd own ow n of o f AS A S 332 C1e stand ard airc ai rc raft raf t .............. ..................... ............... ............... ............... ............. ..... 2.2-1

2.3.

Stand ard dimens di mens io ns .............. ...................... ............... ............... ............... ............... ................ ............... ............... ............... ............... ............... ......... .. 2.3-1

2.4.

Stand ard airc raft defi niti ni ti on .............. ...................... ............... ............... ............... ............... ............... .............. ............... ............... ............... .......... 2.4-1

3.

Descr ip ti on ...................... ................................. ...................... ....................... ....................... ...................... ....................... ....................... ................... ........ 3-1 3.1.

Fuselage ........................................................................................................................ 3.1-1

3.1.1 General General .................................................................... .................................................................................................................................. .............................................................. 3.1-1 3.1.2 Canopy....................................................................................................................... ................................................................................................................................... ............ 3.1-2 3.1.3 Central structure .................................................................................................................... .................................................................................................................... 3.1-2 3.1.4 Polyurethane Polyurethane white paint anti-corrosive anti-corrosive treatment ................................................................ ................................................................ 3.1-5 3.1.5 Intermediate Intermediate structure and tail boom .......................................................................... ..................................................................................... ........... 3.1-6 3.1.5.1 Intermediate structure ....................................................................................... ........................................................................................................ ................. 3.1-6 3.1.5.2 Tail boom ........................................................................................................................... ........................................................................................................................... 3.1-7 3.1.6 Cowlings and inspection doors .............................................................................................. .............................................................................................. 3.1-8 3.1.6.1 Cowlings ............................................................................................................................ 3.1-8 3.1.6.2 Access to the transmission transmission deck deck ....................................................................................... ....................................................................................... 3.1-9 3.1.6.3 Inspection doors ................................................................................... ................................................................................................................ ............................. 3.1-9

3.2.

Landing gear ................................................................................................................. 3.2-1

3.2.1 3.2.2

3.3.

Nose landing gear.......................................................................... .................................................................................................................. ........................................ 3.2-1 Main landing gear .................................................................................................................. 3.2-2

Cock pit pi t and cabi n ............... ....................... ............... ............... ................ ............... ............... ............... ............... ............... ............... ............... .............. ....... 3.3-1 3.3-1

3.3.1 Cockpit .............................................................................................................. ................................................................................................................................... ..................... 3.3-1 3.3.1.1 General .............................................................................................................................. .............................................................................................................................. 3.3-1 3.3.1.2 Cockpit lay-out ................................................................................................................... ................................................................................................................... 3.3-2 3.3.1.3 Detailed description ........................................................................................................... ........................................................................................................... 3.3-3 3.3.1.4 Access to the cockpit cockpit ..................................................................... ......................................................................................................... .................................... 3.3-7 3.3.1.5 Cockpit lighting........................................................................................... lighting................................................................................................................... ........................ 3.3-8 3.3.1.6 Cockpit heating and ventilation.......................................................................................... 3.3-8 3.3.1.7 Cockpit panes .................................................................................................................... .................................................................................................................... 3.3-9 3.3.1.8 Seats .................................................................................................................................. .................................................................................................................................. 3.3-9 3.3.1.9 Rear panels.......................................................................... ...................................................................................................................... ............................................ 3.3-10 3.3.2 Cabin............................................................................................................................... .................................................................................................................................... ..... 3.3-11 3.3.2.1 General ............................................................................................................................ ............................................................................................................................ 3.3-11 3.3.2.2 Cabin main dimensions dimensions .............................................................................. ................................................................................................... ..................... 3.3-12 3.3.2.3 Cabin floor............................................................................. ........................................................................................................................ ........................................... 3.3-12 3.3.2.4 Access to the cabin cabin ....................................................................... .......................................................................................................... ................................... 3.3-13 3.3.2.5 Emergency Emergency exits .............................................................................................................. 3.3-14 3.3.2.6 Sound proofing upholstery............................................................................................. ............................................................................................... .. 3.3-14 3.3.2.7 Fire safety ........................................................................................................................ ........................................................................................................................ 3.3-15 3.3.2.8 Cabin lighting ................................................................................................................... ................................................................................................................... 3.3-15 3.3.2.9 Ventilation/heating Ventilation/heating of the cabin........................................................................................ 3.3-15 3.3.2.10 Electrical utility equipment ........................................................................................... ........................................................................................... 3.3-16

3.4.

Flight Fli ght co ntro nt ro ls ................ ....................... ............... ............... ............... ............... ............... ............... ............... ............... ............... ............... ............... ............. ..... 3.4-1 3.4-1

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3.5.

Au to mat ic fl ig ht co nt ro l s ys tem .................................................................................. 3.5-1

3.5.1 Advanced Helicopter Cockpit & Avionics System (AHCAS TM) ........................................... 3.5-1 3.5.1.1 Description ......................................................................................................................... 3.5-3 3.5.1.2 FDS functions .................................................................................................................... 3.5-4 3.5.1.3 Display formats .................................................................................................................. 3.5-5 3.5.1.4 Complementary display equipment ................................................................................... 3.5-7 3.5.2 Vehicle Monitoring System .................................................................................................... 3.5-8 3.5.2.1 General description............................................................................................................ 3.5-8 3.5.2.2 VMS functions.................................................................................................................... 3.5-9 3.5.2.3 VMS architecture ............................................................................................................. 3.5-12 3.5.3 Automatic Flight Control Subsystem ................................................................................... 3.5-13 3.5.3.1 AFCS functions ................................................................................................................ 3.5-14 3.5.3.2 AFCS Control and Reconfiguration ................................................................................. 3.5-16 3.5.4 Primary reference sensors................................................................................................... 3.5-16 3.5.5 Customized sensors and peripherals (Options) .................................................................. 3.5-17 3.5.5.1 Radio navigation sensors ................................................................................................ 3.5-17 3.5.5.2 Weather Radar................................................................................................................. 3.5-17 3.5.5.3 Flight management Subsystem ....................................................................................... 3.5-17

3.6.

Power plant ................................................................................................................... 3.6-1

3.6.1 General .................................................................................................................................. 3.6-1 3.6.2 Turbine engines ..................................................................................................................... 3.6-2 3.6.2.1 Description ......................................................................................................................... 3.6-2 3.6.2.2 Modular conception ........................................................................................................... 3.6-5 3.6.2.3 Engine/main gearbox coupling .......................................................................................... 3.6-5 3.6.2.4 Engine controls .................................................................................................................. 3.6-6 3.6.2.5 Engine ratings .................................................................................................................... 3.6-6 3.6.2.6 Fire-extinguishing system .................................................................................................. 3.6-7 3.6.2.7 Engine lubricants ............................................................................................................... 3.6-9 3.6.2.8 Engine washing facility without cowlings opening ............................................................. 3.6-9

3.7.

Fuel system ................................................................................................................... 3.7-1

3.7.1 3.7.2 3.7.3

3.8.

General description................................................................................................................ 3.7-1 Usable fuels ........................................................................................................................... 3.7-3 Optional additional fuel tanks................................................................................................. 3.7-4

Trans mi ssion sy st em ................................................................................................... 3.8-1

3.8.1 General description................................................................................................................ 3.8-1 3.8.2 Main gearbox ......................................................................................................................... 3.8-2 3.8.2.1 Description ......................................................................................................................... 3.8-2 3.8.2.2 Main gearbox cooling system ............................................................................................ 3.8-3 3.8.2.3 Fire detection system in the MGB compartment ............................................................... 3.8-4 3.8.3 Tail drive ................................................................................................................................ 3.8-5 3.8.4 Gearbox lubricants................................................................................................................. 3.8-7 3.8.4.1 Oil capacities of gearboxes................................................................................................ 3.8-7 3.8.4.2 Oil for main gearboxes plus main rotor head hinges ......................................................... 3.8-7

3.9.

Rotors ............................................................................................................................ 3.9-1

3.9.1 3.9.2 3.9.3 3.9.4

3.10. 3.10.1 3.10.2 3.10.3 3.10.4 3.10.5

3.11.

Main rotor head...................................................................................................................... 3.9-1 Main rotor blades ................................................................................................................... 3.9-2 Tail rotor................................................................................................................................. 3.9-3 Tail rotor blades ..................................................................................................................... 3.9-3

Hydr auli c sy st em .................................................................................................... 3.10-1 General description.............................................................................................................. 3.10-1 Hydraulic generation diagram.............................................................................................. 3.10-3 Servo units system .............................................................................................................. 3.10-3 Routing of hydraulic circuits ................................................................................................. 3.10-5 Hydraulic fluid specifications ................................................................................................ 3.10-5

Elect ri cal syst em ..................................................................................................... 3.11-1

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3.11.1 3.11.2

General description.............................................................................................................. 3.11-1 Electrical system block diagram .......................................................................................... 3.11-5

3.12.

Exterio r li ghti ng ...................................................................................................... 3.12-1

3.13.

Ai r d ata s ys tem ............................................................................... ........................ 3.13-1

3.13.1 General ................................................................................................................................ 3.13-1 3.13.2 Pilot circuit............................................................................................................................ 3.13-2 3.13.3 Copilot circuit ....................................................................................................................... 3.13-2 3.13.4 Emergency circuit ................................................................................................................ 3.13-2 3.13.5 Pilot heating ......................................................................................................................... 3.13-3 3.13.5.1 Pilot system.................................................................................................................. 3.13-3 3.13.5.2 Copilot system ............................................................................................................. 3.13-3 3.13.5.3 Emergency system ...................................................................................................... 3.13-3 3.13.5.4 Failure of LCD indication ............................................................................................. 3.13-3

3.14.

Pneumatic syst em .................................................................................................. 3.14-1

3.15.

Grou nd hand ling and pi cketin g ............................................................................. 3.15-1

3.16.

Struc tur al pro vis ion s for opt ion al equip ment ...................................................... 3.16-1

3.17.

Ai rb or ne k it .............................. ........................................................... ..................... 3.17-1

3.18.

Ac co mp anyin g l it erat ur e ....................................................................................... . 3.18-1

4.

Perfor mance ........................................................................................................... 4-1 4.1.

Main performance ......................................................................................................... 4.1-1

4.1.1 4.1.2 4.1.3 4.1.4

Performance on 2 engines..................................................................................................... 4.1-1 Performance on 1 engine ...................................................................................................... 4.1-3 Performance in external load carrying mission...................................................................... 4.1-3 Performance on 1 engine ...................................................................................................... 4.1-4

4.2.

Ab br evi ati on s ........................................................ ........................................................ 4.2-1

4.3.

Perfo rm ance ch arts ...................................................................................................... 4.3-1

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TRADE MARKS & ABBREVIATIONS Trade Marks EUROCOPTER, its logo, SUPER PUMA, SPHERIFLEX, M'ARMS, are trade marks of the EUROCOPTER group.

Abb reviation s AC

Alternative Current

MGB

Main Gearbox

AFCS

Automatic Flight Control System

MGW

Maximum Gross Weight

AP

Autopilot

MTOW

Maximum Take-Off Weight

AUW

All-Up Weight

NATO

North Atlantic Treaty Organisation

DC

Direct Current

Nf

Free turbine speed

EASA

European Aviation Safety Agency

Ng

Engine Generator Speed

EEW

Equipped Empty Weight

ΔNg

Ng difference

ELT

Emergency Locator Transmitter

NR

Rotor speed

EPC

Engine Power Check

OAT

Outside Air Temperature

FAA

Federal Aviation Administration

OGE

Out of Ground Effect

FADEC Full Authority Digital Engine Control

PA

Pressure Altitude

FAR

Federal Aviation Regulations

RH

Right Hand side

FLI

First Limitation Indication

SL

Sea Level

GHW

Ground Handling Wheel

T4

Power turbine inlet temperature

GPS

Global Positioning System

TAS

True Airspeed

GS

Glide Slope

TBO

Time Between Overhaul

ICS

Intercommunication System

TC

Transport Canada

IGE

In Ground Effect

TGB

Tail Gearbox

ISA

International Standard Atmosphere

TOP

Take-Off Power

JAA

Joint Aviation Authorities

TQ or

Engine Torque

JAR

Joint Aviation Requirements

UK

United Kingdom

LCD

Liquid Crystal Display

USA

United States of America

LG

Landing Gear

VHF

Very High Frequency

LH

Left Hand side

VIP

Very Important Person

LOC

Localizer

VNE

Never Exceed Speed

LRU

Line Replaceable Unit

VOR

VHF Omnidirectional Radio Range

MCP

Maximum Continuous Power

(RPM of the main rotor)

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Units °C

degree Celsius

kW

kilowatt

°F

degree Fahrenheit

lb

pound

A

ampere

lb/h

pound per hour

A/h

ampere per hour

m

meter

bar

bar

m² 3

square meter

ch

cheval-vapeur

m

cSt

centistokes

mm

millimeter

daN

dekanewton

m/sec

meter per second

ft

foot

nm

nautical mile

ft/min

foot per minute

psi

pound-force per square inch

ft²

square foot

rpm

revolutions per minute

ft

cubic foot

shp

shaft horsepower

hr:min

hours:minutes

UK gal

gallon UK

kg

kilogram

UK gal/h

gallon UK per hour

kg/h

kilogram per hour

US gal

gallon US

kg/m²

kilogram per square meter

US gal/h

gallon US per hour

km

kilometer

V

volt

km/h

kilometer per hour

VA

voltampere

kt

knot

W

watt

3

cubic meter

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1. Foreword

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1.1. Foreword

The twin engined AS332 C1e Super Puma is the most recent evolution of the Super Puma MK1 helicopter family which beneficiate from the most advanced avionics developped for the EC225 aircraft, which main features are :    

Reduced pilots workload for better mission effectiveness Outstanding autopilot precision for enhanced safety Full compatibility with the new generation mission equipment EGPWS …) New state-of-the-art technology to solve obsolescence issues.

(Flir, D-map,

The SUPER PUMA AS332 C1e incorporates technological features introduced by EUROCOPTER in the field of maintenance and operational capabilities : 

An Advanced Helicopter Cockpit & Avionics System (AHCAS™) including a Flight Display System (4 AMLCD) and a Vehicle Management system (2 EID)



Automatic Flight Control system which is a 4-axis dual duplex digital auto-pilot, fully integrated with the avionics



Composite material for the rotor blades



Modular design for the major transmission components



All-weather IFR capability, even under icing conditions with a special optional equipment



.On condition" maintenance for most of the equipment items and several transmission assemblies



Improvement of the survivability in case of crash.

The twin engine concept, combined with an extensive power reserve, makes the SUPER PUMA an aircraft particularly suited to various missions, and means safety and retention of operational capabilities over a wide altitude and temperature envelope. It results in excellent performance on one engine.

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It is a medium weight Helicopter with a large cabin particularly suited for passenger transport. Fitted with the appropriate equipment, the Super Puma is capable of the following missions further to the basic passenger transport mission : o o o o

Note :

Internal load or external load carrying Casualty evacuation and MEDEVAC missions Search and rescue VIP transpor

The pictures, front faces of the control boxes, panels and displays included in this document are given for information, and could be modified without prior notice.

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2. Characteristics

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2.1. Standard lay-out



Minimum crew (D.G.A.C. Category A and B certification) -



Passenger transport (in addition to the crew) -



VFR : 1 pilot (with at least one lane of each autopilot channel engaged) IFR : 2 pilots

Up to 17 "comfort" seats

Casualty-evacuation (in addition to the crew) : -

up to 6 stretcher-patients + 6 seated places

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2.2. Weight breakdown of AS 332 C1e standard aircraft

kg

Ib



Standard aircraft empty weight (including unusable fuel and engine oil)

4,450

9,810



Maximum gross -weight at take-off

8,600

18,960

4,150

9,150

160

350

. Payload + fuel

3,990

8,800

Maximum operating weight wit h external load

9,350

20,615

4,900

10,805

4,500

9,920



Corresponding useful load

. Crew of 2







Corresponding useful load

Maximum sl ing capacity

Note The empty weight of the standard aircraft covered by this Type Specification includes the Flight Manual, the lubricants and unusable fuel. Tolerance on the weight breakdown : ± 2 %.

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2.3. Standard dimensions

Ai rc raf t o ver all di mensi on s : 

Length, rotor rotating

18.70 m

61.35 ft



Main rotor diameter

15.60 m

51.18 ft



Height at top of tail rotor*

4.92 m

16.14 ft



Tail rotor diameter

3.05 m

10.00 ft



Length with blades folded

15.53 m

50.95 ft



Width with blades folded

3.79 m

12.43 ft



Height at rotor head*

4.60 m

15.09 ft



Ground clearance under cabin*

0.47 m

1.54 ft



Ground clearance under tail rotor*

1.88 m

6.16 ft



Fuselage width

3.38 m

11.08 ft

Cabin dimensions : 

Length at floor level

4.68 m

15.35 ft



Maximum length

6.05 m

19.85 ft



Maximum width

1.80 m

5.90 ft



Maximum height

1.55 m

5.08 ft

7.80 m2

84.00 sq.ft

11.40 m3

402.60 cu.ft



Area available



Volume available

*

At 4,460 kg (9,830 lb)

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Ac ces s t o t he c abi n : 

Lateral sliding doors :

. Width

1.30 m

4.26 ft

. Height

1.35 m

4.43 ft

1.75 m2

18.83 sq.ft

. Maximum width

0.98 m

3.21 ft

. Minimum width

0.70 m

2.29 ft

. Length

1.90 m

6.23 ft

. Area



Rear panel :

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2.4. Standard aircraft definition

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3. Description

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3.1. Fuselage 3.1.1

General Of semi-monocoque construction, the fuselage includes from front to rear : the canopy, the central structure, the intermediate structure and the tail boom. Paint finish is applied in accordance with the customer's paint scheme.

1 2 3 4 5 6 7

air intake cowling fire wall engine cowlings transmission deck engine sliding cowling upper structure tail rotor drive shaft fixed cowling 8 tail rotor drive shaft opening fairings 9 tail gearbox fairing 10 pylon fairing

11 12 13 14 15 16 17 18 19 20 21

horizontal stabilizer tail skid (steel) lower fin tail boom intermediate structure loaching hatch cabin door (RH door opposite hand) cabin floor landing gear fairing footstep hydraulic line protective channel

22 23

bottom structure fuel tank compartment trimming cockpit floor console radome copilot's door (pilot's door on opposite side) canopy forward fixed fairing (overhead panel)

24 25 26 27 28 29

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3.1.2

Canopy The canopy is made up of a welded light alloy framework supporting transparent panels which ensure maximum visibility for the crew. The center windshield panel is in clear plexiglass. The lateral win dshield , in front of the pilot and the copilot, are made of triplex glass and are electricaly de-iced. All other transparent panels are in plexiglass. The canopy is covered by a metal box-type structure secured by screws to the cockpit bulkhead and canopy framework. This fairing supports the control quadrant and the overhead panel. Two large jettisonable doors give access to the cockpit. Each of them is fitted with fixed parts for an armour panel. The nose of the aircraft consists in a radome made of fiberglass and resin. It is articulated and can be swinged upwards to give access to the front avionics compartment.

3.1.3

Central structure It results from the assembly of the lower and upper structures, at the level of the main frames. This assembly builds a rigid frame-work capable of protecting the passengers from being crushed by the upper mechanical assemblies (main gearbox, main rotor, engines) in case of a crash landing. The engine bay floor bears the supporting brackets of the engines. The transmission deck bears the MGB suspension mount ("barbecue"), and the 3 suspension bar attachment brackets which insure the transmission of the lift strains, the torque and the bending moments of the main rotor mast to the central structure. Some hard points are located on the lower structure, just in front of the sliding plug doors, on which can be fixed such equipment as : FLIR, search lights, etc... 3 hard points have been integrated in the fuselage just above the RH sliding plug door to receive the fittings for an optional rescue hoist. The lower structure bears also the hard points meant to receive the floatation gear fittings (front module and lateral modules).

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Finally, the rear part of the lower structure embodies the hard points on which are fixed the main landing gear assemblies and the optional sponsons.

Upper structure

Lower structure

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Hard points

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3.1.4

Polyurethane whit e paint anti-corr osi ve treatment The polyurethane white paint is applied as shown on the drawing here under:

1 : inside sliding cowling 2 : mechanical floor 3 : structure around cabin windows 4 : structure behind sponsons 5 : nose wheel compartment

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3.1.5

Intermediate str uct ure and tail boom

3.1.5.1

Intermediate str uct ure The intermediate structure links the central part of the fuselage to the tail boom. It consists of two half-shells rivetted together at the top and closed, at the bottom, by a hatch. A frame ends at the rear the intermediate structure and provides the punction with the tail boom assembly. The intermediate structures constitutes the rear section of the cabin. It includes two LH and RH stowage compartments for the airborne kit and luggage. A storage place is also provided for a maintenance ladder.

Intermediate struc ture, view tow ard the rear

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3.1.5.2

Tail boo m It is made up of 5 main assemblies. The horizontal structure : It extends the outside contour of the forward structure, its upper datum line remaining in the transmission platform plane.



It is constitued by frames linked by stringers. This structure is covered by light alloy skin panels. The junction of this structure with the intermediate structure is made by bolts and localing spigots making for an easy and quick assembly. This structure supports the tail rotor drive shaft bearings, the intermediate gearbox attachment fitting and some tail rotor controls. 

The oblique structure : It consists of two longitudinal and transverse ribs. The skin panels are in light alloy and stiffened. It is joined to the horizontal structure by rivetted finger butt-straps. It supports the tail gear box and the stabilizer.



The ventral fin : It consists in a light alloy skin rivetted on 2 frames integral with the tail boom



The tail skid : It is made of a steel tube flattened and curved at its free end. At the other end, it is hinged in the vertical plane and held in its rest position by a strut incorporating a shock absorber. It protects the tailboom when landing approches are made in an excessively nose-up attitude.



The stabilizer : It improves the flight qualities of the aircraft. The composite stabilizer consists of 5 metal sandwich ribs onto which the upper and lower light alloy/nomex honeycomb sandwich skin panel are bonded. It is made up of a grafite fiber spar tube which passes through the oblique structure and rests on two bearings. An attachment fitting on the leading edge of the profile insures a correct angular positioning of the stabilizer.

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3.1.6

Cowlings and ins pection door s

3.1.6.1

Cowlings The transmission and ancillary components are protected by light alloy or laminated cowlings. These removable, hinged or sliding cowlings include the following : 

The engine air intake stub frame which can slide forward on rails to facilitate maintenance. The engine air intakes are protected by grids which perform a dual function as follows : . .

protection from ingestion of foreign matter protection from ice.

For operation in severe icing conditions at temperatures close to 0°C, the engine antiicing protection by grids is completed by heating mats, the purpose of which is to heat the air intake duct. This anti-icing device includes two heating mats per air intake, made of elastomeric material, which are bonded inside the duct. They are fed with 3-phase 115 V a.c. current. Heating power is 865 Watt for each air intake. The air intake bracket is provided with a boss for housing the fan of the heating/ventilation system. 

Two engine cowlings, acting as a work platform when open



The sliding cowling



The tail rotor drive hinged cowlings



The landing gear fairings



The hydraulic and electric line fairing



The TGB and tail servo-control fairing

1 air intake stub frame 2 engine cowlings 3 sliding cowling 4 tail drive hinged cowlings 5 landing gear fairings 6 hydraulic and electric line fairings 7 TGB and tail servo control fairing

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To facilitate pre-flight inspections, ports have been drilled in cowlings 2, 3, 4 and 7 for visual inspection of engine, IGB, TGB oil levels as well as inspection of the hydraulic fluid level in reservoirs from the ground. The main gearbox oil level can be checked through a small door located on the sliding cowling, accessible from the cabin floor.

3.1.6.2

Access to the transm issi on deck Located at the rear of the RH sliding door, foot-steps built into the structure and a handle provide access to the transmission deck.

3.1.6.3

Inspectio n door s

3.1.6.3.1.

Internal inspection doors

1 2 3 4 5 6 7

pilot station overhead Panel starting unit cover electrical equipment panel electrical cabinet doors and panels flying control cabinet panels inspection door on pilot's bulkhead inspection door on pilot's floor

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3.1.6.3.2.

External inspection doors These doors give access to miscellaneous equipment items of the aircraft.

1

Radome

6

Door to rear transverse tank content gauge

2

Nose landing gear door

7

Door to rear transverse tank water drain

3

Door to transverse fuel tank content

4

Door to transverse fuel tank water drains

and jet pump 8

and door 5

Longitudinal fuel tank connector base door blanking panel

3.1.6.3.3.

th

Central 5 fuel tank unit content gauge and drain access door

9 10

Pylon door Central compartment blanking panel

Cabin floor panels The various floor panels give access to the following :



Electrical equipment Fuel tanks Fuel transfer pump



Landing gear hydraulic power generator and lower flying control bellcranks.

 

1 2 3 4 5 6 9

Electrical equipment } } } Fuel tanks } } }

7 8

Fuel transfer pump Landing gear hydraulic power generator and lower flying control bellcranks

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3.2. Landing gear The Super Puma is equipped with tricycle landing gear . The nose landing gear leg retracts into the fuselage while the main landing gear retracts into the side fairings.

3.2.1

Nose landi ng gear

The nose landing gear consists of a vertical oleo-pneumatic shock absorber retracted rearwards by an actuating cylinder. This cylinder is provided with internal mechanical systems locking the landing gear in "UP" and "DOWN" positions. The nose landing gear is equipped with two 6-inches aluminum alloy wheels and two tubeless tires inflated to 7 bars. The twin-wheel assembly can rotate through 360° and is provided with a friction type anti-shimmy device and a self-centering system effective over ±120°. At the end of the shock absorber extension stroke, wheel rolling direction automatically comes back aligned with the center line of the helicopter. A manual control enables the pilot to lock the gear in the centered position. When unlocked, the wheel assembly swivels freely in accordance with the maneuvers of the aircraft on the ground. The total travel of the shock absorber is 0.35 m. The twin-wheel axle is equipped with spools required for towing or steering using a tow bar defined by EC.

332 C1e 11.102.01 E

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3.2.2

Main landi ng gear Each main landing gear consists of a trailing arm pivoting over a transverse pin secured to the fuselage and a nearly vertical oleo-pneumatic shock absorber strut securing the trailing arm to the fuselage. This shock absorber strut acts both as a shock-absorber and retraction cylinder, and the landing gear is locked hydraulically in extension and retraction. For a normal landing, only the low pressure section is in operation. If the limit landing conditions are exceeded, the high pressure section absorbs the additional energy without failure. Landing gear retraction in flight is obtained by compressing the low-pressure section and by transferring pressurized oil to the helicopter hydraulic system.

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Each main landing gear is equipped with a 10-inches wheel mounted on a swinging arm. This wheel is fitted with a tubeless tire inflated to 9 bars (130 PSI) The normal shock absorber travel is 0.32 m while the total travel of 0.51 m is used in case of a crash-landing and for retracting the main gear in flight.

The landing gear system is supplied from the LH hydraulic system which incorporates 2 pumps. In the event of an electric or hydraulic failure, the landing gear can be extended using an emergency electric pump connected to the battery, the supply of which is provided from a reservoir independent from the main system. The main axles are capable of accommodating mooring rings for lateral tying down. Each wh eel is equipped with a single disc wheel brake controlled hydraulically. The hydraulic brake unit is actuated either by the pilot's or copilot's feet control under pressure from the hydraulic power supply system or by a hand control under accumulator pressure (parking). In static conditions, the brake system is capable of holding the helicopter on 10° slopes. The landing gear operation is controlled from the cockpit through an "up and down" toggle switch, a control panel and a nose wheel locking control. Two cockpit lights provide a warning to the crew when the landing gear is not extended at aircraft speed below 110 km/hr (60 kts)

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1 2 3 4 5

-

Landing gear operating and landing gear "extended locked" indicators Landing gear operating in flight "Retraction - extension" control switch Emergency electric pump operating Landing light

The following data are detected and displayed in the cockpit: 

Nose landing gear extended and locked, wheel centered



Nose landing gear retracted and locked



Landing gear in motion



Main landing gear extended



Main landing gear retracted and locked



Emergency electric pump operating



Landing gear not extended and air speed < 60 kts

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3.3. Cockpit and cabin 3.3.1

Cockpit

3.3.1.1

General It has been designed for a crew with 2 pilots and possibly a 3rd man. The pilot seat is located at the RH place, the copilot's one at the LH place (and the 3rd man in the corridor linking the cockpit and the cabin). The ergonomy has been thoroughly studied in order to optimize the exterior visibility , the flight instrument lisibility and the accessibility of all controls from the crew stations.

1 - Overhead panel 2 - Pilot seat 3 - Instrument panel

4 - Radio console 5 - Copilot seat 6 - Third man seat Cockpit ergonomy

332 C1e 11.102.01 E

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3.3.1.2

Cockp it lay-out Slightly raised relative to the cabin floor the cockpit floor carries the pilot and copilot seats as well as the radio console, topped by the instrument panel, the dual controls, the emergency hydraulic system control and monitoring equipment, the nose landing gear self-centering device and the mechanical pitch indicator. This floor can withstand the crash loads applied by the pilot and copilot seats.

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27

Jack connector Copilot's general cut-out handle Map reading light Overhead light Control lever quadrant Overhead panel Pilot's general cut-out handle Ventilation outlets Standby compass Pilot cyclic stick Sub-panel Pilot door jettison handle Pilot's rudder pedals Pedal adjustment Collective pitch friction lock Circuit breaker panel Ash-tray Nose wheel centering device Pilot collective stick Lower radio console Upper radio console Copilot collective stick Copilot rudder pedals Copilot door jettisonable handle Copilot cyclic pitch stick Instrument panel Ventilation control

The cockpit has been designed to accommodate all the flight instruments as well as optional radio-communication or radio-navigation equipment enabling the aircraft to be flown in IFR conditions.

332 C1e 11.102.01 E

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3.3.1.3

Detailed descri pti on

3.3.1.3.1.

Instrument panel The instrument panel is anti-vibration mounted and can tilt backward for maintenance operations. It incorporates a vizor to provide anti-glare facility. The instrument panel layout presented here below is only indicative. It can be modified according to the customization required by the contractual definition.

1, 4 - Piloting, Navigation and Mission Multifunction displays 2 - Ice detector (optional) NR/ILS button indicator Master warning light (red 'WARN' and amber 'CAUT' lights) Landing gear warning light 3 - NR/Nf indicator 5 - Stop watch 6 - Warning panel 7 - VMS displays 8 - Automatic Flight Control Panel (AFCP) 9 - ISI (Integrated Stand-by Instrument) and dedicated lighting control 10 - Hydraulics command panel (LH–, RH and Aux)

The instrument panel lay-out here above can be completed with to the customer installation.

NB: pictures are given for information only (not contractual) 332 C1e 11.102.01 E

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3.3.1.3.2. Sub-panel

1 Standby battery monitoring and control push-button 2 LH, RH and central windscreens de-icing control switches (central deicing in option) 3 Rear I.C.S. cut-off switch (optional) 4 Engine monitoring panels and ventilation control switches 5 Over-speed test selector switch and associated warning lights 6 Power test switch 7 Aural warning switch 8 Pitots de-icing control switches rd 9 Windscreens wiper control switch (3 wiper in option)

3.3.1.3.3.

Radio consoles

1 Landing gear control and monitoring 2 Auto-pilot hydraulic unit electrical preheating 3 Landing gear emergency control and parking brake control 4 - Fuel circuits controls and monitoring -Auto pilot control - Optional equipment 5 VNE –PITCH display plate

The console panel lay-out here above is only indicative. It can be modified according to the customer requirements.

332 C1e 11.102.01 E

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3.3.1.3.4.

The "cap" fairing and the overhead panel An extension of the transmission platform provides the cockpit ceiling. It accommodates the control and monitoring equipment for the various installations, chiefly electrical, the fuel flow and fuel shut-off levers as well as the rotor brake control and the master cut-off handles.

(example of installation) The master cut-off handles, one above the pilot and the other above the copilot are available to the crew, including the 3rd crew member. They cut off the electrical power generators and operate the fuel shut-off cocks when they are actuated in the event of a crash. Then, only the equipment connected to the direct battery bus bar remain energized. These controls or warning lights are normally fitted in the locations described below. However, they may be relocated depending on the supplementary equipment fitted to the aircraft. Other controls, associated with optional installations, are also located on the overhead panel.

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1 2 3 4 5 6 7 8 9

Mission selector switch Electrical generation Overhead panel lighting control Console and sub-panel lighting control Instrument panel lighting control, Pilot's side Light-up signs, lighting control Map reading light, RH side Landing lights Emergency exits, lighting control

10 11 12 13 14 15 16 17 18

Cabin lighting control External lighting control Fire extinguisher control panel unit Electrical generation monitoring panel Fire detection circuit test panel Air conditioning control panel (*) Ventilation control switch Map reading light, LH side Instrument panel lighting control Copilot's side 19 Master Switch

(*) Optional

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3.3.1.4

Access to the cock pit

Access is provided : From the outside, on the RH and LH sides by footsteps built into the structure and two jettisonable doors (pilot and copilot) The travel of the cockpit doors is controlled by pneumatic door stops The pilot and copilot doors are fitted with a map case. From the cabin, by a central corridor located between two cabinets. A partitioning curtain is fitted between the cockpit and the cabin.

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3.3.1.5

Cockpit lig hti ng There are four white lighting systems on the cockpit overhead panel :   

a bi-mode general lighting by dome light an extension light used in an emergency and supplied from the battery two white map spot lights (pilot, copilot)

The instrument panel and sub-panel (pilot, copilot), the control pedestal and the overhead panel are supplied by two circuits (normal and emergency circuit). These circuits are separately supplied and the lighting brightness is adjustable through 4 potentiometers : -

-

one for the instrument panel pilot's side, one for the instrument panel copilot's side, one for the console and sub panel, one for the overhead panel.

NVG Compatibility can be proposed as an option.

3.3.1.6

Cockpit heating and ventil ation Ventilation and heating of the cockpit are ensured by : 

one front aerator two panel demisting dif fusers, located behind the instrument panel two diff users at floor level two adjust able aerators on the ceiling

A control located behind the pilot's seat makes it possible to direct the cold air or the hot air either to the upper demisting ramps in the cockpit or to the lower ones or to both of them simultaneously. The ventilation and heating system of the cockpit is part of the general ventilation heating system described after.

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3.3.1.7

Cockp it panes The pilot and copilot front panes are made of triplex glass. All the other panels are made of perspex. The pilot and copilot front panes are de-iced. The de-icing controls (one for each pane) are located on the instrument sub-panel. The pilot and copilot front panes are fitted with windshield wipers. Their controls are fitted on the instrument sub-panel. The lateral windshield wipers three positions control swit ch makes it possible to select the following operating conditions: manual (the wiper is operated by a push button located on the collective lever), slow speed, fast speed.

  

3.3.1.8

Seats The pilot's and copilot's seats are adjustable in height and fore-and-aft. They are equipped with belts and extensible should er harnesses.

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A "third man" seat is located in the corridor connecting the cockpit to the cabin. This seat can be folded up and locked when not in use. It has a back rest and an extensible belt.

3.3.1.9

Rear panel s The cockpit is separated from the cabin by two cabinets (flying controls cabinet on the RH side, and electrical cabinet on the LH side). The bulkhead behind the copilot's seat carries a battery of circuit breakers. In the corridor, the electrical cabinet panel carries the remaining circuit breakers.

1 2 3 4 5 6

Distribution panel Utility light Distribution panel Distribution panel Emergency total and static pressure selector cocks Fire extinguisher

Cockpit bu lkhead and access

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3.3.2

Cabin

3.3.2.1

General

The cabin, of a straightforward shape without projections, is largely centered under the main rotor axis. It offers ample loading space over a wide C.G. range. The cabin paint and trimmings are in light beige colour.

View of the cabin toward the front With utility seats installation (option)

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3.3.2.2

Cabin main dim ensions

3.3.2.3

Cabin floor The cabin floor, fitted with 13 tie-down rings, can withstand a distributed load of 310 lb/sq.ft (1500 kg/m 2). It is made up of panels screwed to the bottom structure. These panels are also used for access to the various items of equipment located under the floor (tanks, fuel system etc.).

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3.3.2.4

Access to the cabin Access is gained through 2 footsteps and 2 lateral sliding plug doors . These doors can be opened and locked in the open position in flight. The maximum speed for opening these doors is 100 km/hr (55 kts) and the maximum speed with one or both doors open is 278 km/hr (150 kts). A warning light on the instrument panel comes on when a door is not locked. The rear of the cabin includes a removable panel . It can be used to introduce long loads into the cabin. Flying is possible with this panel removed, at any speed.

Rear panel remov ed

Rear panel ins talled

Lateral sliding plug door

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3.3.2.5

Emergency exit s Emergency exits include the two cabin sliding doors and the pilot and copilot doors (these four components can be jettisoned from inside and outside) plus all the cabin windows and the bubble window of the rear panel.

Main emergency exit door s

Window jettison

3.3.2.6

Sound proo fin g upho lst ery This consists of removable padded cloth panels so as to facilitate maintenance operations. It limits the noise in the cabin to a comfortable level.

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3.3.2.7

Fire safety An extinguisher and an axe are fitted to the door of the flight-control cabinet at the front of the cabin. An anti-smoke device (oxygen bottle + mask) is located in the corridor between the cockpit and the cabin.

3.3.2.8

Cabin lig hti ng The cabin lighting system includes two fluorescent light strips, controllable from the cockpit. Lighting of the panels bearing emergency markings close to the sliding doors, and of both "emergency exit" panels is ensured by the standby battery which also supplies the three dimmer lamps in the fluorescent strips in the case of failure of the main system.

3.3.2.9

Venti lation /heatin g of the cabin The ventilation and heating are provided by a flow of cold air or a mixture of cold and hot air into the cabin. 

cold air is taken from the outside by a fan located between the two engine air intakes



hot air is bled from the engines

The hot air flow is adjusted by the control located behind the pilot. The installation is capable of increasing the cabin temperature by 35°C relative to the outside air temperature. The fan control is located on the overhead panel. Heating is controlled from a needle valve mounted on the RH side of the cockpit overhead panel. A three-way valve ensures the following functions : 332 C1e 11.102.01 E

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  

hot air supply to the cabin and cockpit hot air supply in the cockpit and autopilot hydraulic unit pre-heater complete shut off of the heating.

Upper diffuser unit

Lower diffuser unit

Air is distributed through 10 inlets in the upper part of the cabin and through 4 diffusers in the lower part of the cabin. Stack air is evacuated through 2 openings installed rearwards of the cabin.

3.3.2.10

Electrical uti lit y equipm ent 6 power receptacles (28 volts, 15 A) are located in the cabin for feeding heating blankets or extension lights. On the rear upright of the RH door a utility connector (28 volts, 25 A) enables ground utility equipment to be plugged in.

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3.4. Flight controls The SUPER PUMA cockpit is provided with dual controls (pilot on the RH side, copilot on the LH side). The controls, of a conventional type, include : 

Two floor hinged cyclic control sticks controlling the variation of the swash plate tilt angle



Two collective pitch levers which control the vertical movements of the swash plate. By high wind, when the aircraft is on the ground, the cyclic stick can be secured by two rods and the collective stick by one rod. Collective pitch lever handles 1

Hoist cable cutter*

2

Landing light on/off and automatic retraction

3

Landing light elevation and azimuth control

4

Emergency floatation gear inflation control*

5

Collective pitch lever disengagement

6

Engine bleed valve control

7

Windshield wiper control

8

Go around engagement/disengagement

9

Collective lever trim

10

AP hydraulic unit cut-off

11

Free

*

Operational only if the relevant optional items of

equipment have been ordered

332 C1e 11.102.01 E

3.4-1

This document is the property of EUROCOPTER; no part of it shall be reproduced or transmitted without the express prior written authorization of EUROCOPTER and its contents shall not be disclosed. © EUROCOPTER 2013


Cyclic stick handles 1 2 3 4 5 6 7 8 9 10 11

Free Free Ground speed Cyclic trip Free Cyclic stick disengagement Disengagement of AP higher modes Free Cargo sling load release* Autopilot disengagement I.C.S. PTT*

*

Operational only if the relevant optional items of equipment have been ordered.



Two adjustable pedal units acting on the tail rotor pitch setting, thereby controlling the aircraft in yaw. A mechanical coupling is provided to vary automatically the tail rotor pitch setting for any change in collective pitch setting



Four hydraulic servo units , three of them acting upon the swash plate (cyclic and collective pitch channels), and the fourth one on the spider of the tail rotor control system Each of the structurally independent chambers of the servo units is equipped with a distributor and is fed by a separate hydraulic system



The hydraulic unit of the autopilot (auxiliary servo controls).

332 C1e 11.102.01 E

3.4-2



Flight control block d iagram

332 C1e 11.102.01 E

3.4-3



Main and tail rotor co ntrols

332 C1e 11.102.01 E

3.4-4



3.5. Automatic flight control system

The helicopter is fitted with a new basic avionics system, the Advanced Helicopter Cockpit & Avionics System (AHCASTM) The AHCAS is designed to assist the crew in performing Flight management, Navigation and Communication management, Vehicle management and Mission management through glass cockpit displays, digital computers and associated centralized cockpit controls. Depending on customer's requirements, the AHCAS is completed with dedicated :

3.5.1



Flight Management Subsystem



Radio-navigation subsystem



Radio-communication subsystem

Advanced Helico pter Cockpit & Avio nics System (AHCAS TM) The AHCAS is mainly composed of the following subsystems, each being redundant and re-configurable in case of failure: 

Flight Display Subsystem (FDS) - quadruplex



Vehicle Management Subsystem (VMS) – duplex with backup



Automatic Flight Control Subsystem (AFCS) – dual/duplex

These subsystems are tightly integrated so as to provide the best crew functionality and availability, to save space and load, while reaching the required rules and safety level for civil and military certification.

The general system architecture is presented on the following page.

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(*) optional

System architecture

Flight management system, radio navigation, radio communication and mission systems are optionals.

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Flight Display Subsystem The Flight Display Subsystem (FDS) is a full glass cockpit display system built from up to date Active Matrix Liquid Crystal Display (AMLCD) technology. Symbology and control modes are enhanced from the proven MK2 IFDS, with for instance a Flight and Navigation Display (FND) mode gathering piloting and short term navigation information on a single format, thus allowing use of the other displays for mission purpose (FLIR/DMAP/NAV PLAN….) All the sensors data are concentrated in every display, resulting in a quadruplex system which allows great Minimum Master Equipment List (MMEL) enhancement.

3.5.1.1

Description The FDS comprises four 6" by 8" landscape AMLCD displays hereafter called MultiFunction Display (MFD) suitable to display every available format (FND, Video, DMAP, FLIR…).

6” x 8” AMLCD Multi Function Display The four physically identical MFDs are installed on the instrument panel, two in front of each pi lot . These MFDs are controlled by mean of bezel software keys. Each MFD is a smart display, composed of a Processing Unit (PU) and a Keyboard & Display Unit (KDU). The Processing Unit includes the following parts:  Symbol generator  Processing unit  I/O interfaces including Arinc 429, Arinc 453, discrete and video  DC 28V Power supply The Keyboard & Display Unit includes the following parts:  Flat screen Soft keys on the screen bezel  332 C1e 11.102.01 E

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3.5.1.2

FDS funct ions Each MFD acquires data from Primary reference sensors, radio-navigation sensors, and computed data from AFCS, FMS, VMS and generates the relevant formats to be displayed using its own symbol generator . Therefore the display system works in a quadruplex mode, so ensuring the MMEL requirement.

Each MFD performs checking of data to be displayed at several level of processing from acquisition up to LCD physical screen. The MFD will raise an appropriate flag if : 

invalid, failed or out of range data to be displayed is received



discrepancy detected between data coming from duplex sensors



manual reconfiguration on remaining sensor

The MFDs act as data concentrators and re-emitters for cross monitoring purpose, for control acknowledgment and to provide necessary data to any subsystems. The data exchange between MFDs and the system is performed according to Arinc 429 standard and using few discrete signals (for pin-programming and safety redundancy). The MFD's soft-keys are designed such a manner they control the current display format function. As examples : 

when Navigation Display format is selected, soft-keys permit to select navigation source or declutter of navigation data

The MFDs receive video directly from available video sensors (e.g. FLIR, DMAP..).

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3.5.1.3

Display form ats The basic display formats on the MFDs are the following :

3.5.1.3.1.

Flight and Navigation Display (FND) This format gathers every information allowing piloting and control (on the upper part of the format) as well as navigation short term trajectory (on the lower part). The FND is declined in several sub-modes which differs mainly by changing the lower part of the screen (horizontal situation) according to the flight phase. This sub-modes are:    

HSI SCT

: :

Horizontal Situation Indicator 120° sector flight plan

WRX HOV

: :

weather radar mode to be super imposed or not on SCT format hovering situation

FND format with HSI sub mode

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3.5.1.3.2.

Navigation Display (NAVD) This format, associated to navigation management, presents the long term flight plan in several sub-modes : 

ROSE

:

360° heading rose



SCT

:

120° sector horizontal situation (with or without weather radar)



PLAN

:

360° horizontal situation with or without radar

NAVD format w ith SCT sub m ode

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3.5.1.4

Complementary disp lay equipm ent

The complementary display equipment set comprises the following equipment: 

ISI (Integrated Stand by Indicator) that delivers back-up attitude and Air data parameters displayed in a PFD format

Integrated Standby Indicator     

Warning and Caution Master warning (red): one in front of each crew member Master caution (amber): one in front of each crew member NR/Nf: indicating to the pilot and copilot the rotor and free turbine RPM one clock

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3.5.2

Vehic le Monit oring System

3.5.2.1

General descr ipti on The Vehicle Monitoring System (VMS) provides engine and vehicle data concentration for display with high safety levels. Its main function is to process and display the following parameters: 

Gas generator speed (N1)



Total Output Temperature (TOT)



Torque (TRQ)



Pressure and Temperature of Engines



Pressure and Temperature of Gear Boxes



Hydraulic pressure



Fuel Pressure



Outside Air Temperature





Weight System maintenance function

The System is composed of a duplex computer (AMC: Aircraft Management Computer) and two 3.9’’x5.2” displays (EID: Electronic Instrument Display) located one over the other on the central part of the instrument panel. A set of sensors is directly linked to the computer or through an ancillary board (SIU: System Interface Unit).The AMC computes these analog data to deliver on a digital link the value of the parameters to be displayed on the two EIDs. This Link also delivers to Flight Management System (FMS) the following informations: 

Fuel Flow, fuel quantity and gross weight



Pressure and temperature coming from the ADC

A dedicated maintenance mode provides basic maintenance data of analogs sensors connected to VMS, FMS and AFCS status.

VMS, ADC,

A maintenance PC can be linked to the system for up and down load.

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3.5.2.2

VMS funct ions The VMS performs the functions listed hereafter.

3.5.2.2.1.

Engine data monitoring The engines are monitored on the following format ( ENG for mat): Engine 1 & 2 N1 value Engine 1 & 2 TOT value  Torque (TRQ) value  OAT  Engines states (Start / Training /  Bleed valve/Bleed valve offset) System messages  Attention getters for OEI duration  AMC reconfiguration yellow  rectangle 

OFS

STA

TRQ % 70.0

N1 % 98.0

25.0

45.0

T TOT°C 750

N1 % 65.0

AMC 2 TOT°C 650 START

OAT 24.5 °C MESSAGE LINE

ON

+ -

OFF

3.5.2.2.2.

Vehicle data monitoring The vehicle data formats are displayed on the lower EID in a normal mode of operation a)

The vehicle parameters are monitored on the VEH for mat:

     

ENG 1 OIL P T

Engine 1 & 2 oil temperature Engine 1 & 2 oil pressure Main gear box oil temperature Main gear box oil pressures Hydraulic pressures Fuel pressure

3.0

MGB

128 °C BAR STBY T

67 °C FUELP FUEL P

P

6.0 BAR P

3.5 BAR HYP P

1

170 BAR

1.1 BAR

OIL

ENG 2 OIL P 3.0 BAR T

67 °C 2 FUEL FUEL P FUEL P P

175 BAR

A UX

1.1 BAR NUM

120 BA R ON

STATUS

ENG

+ -

OF F

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b)

The STATUS format displays miscellaneous information relevant to subsystem parameters:

-

ADC parameters Fuel flows NR value IGB & TGB temperature values Failure Message (at LRU level)

1 ------120

SYSTEM STATUS 2 OAT 17 ALT 900 TAS 120 F.F 120 IGB TGB NR

125 110 280

°C FEET KNOTS L/H °C °C RPM

LRU FAILED: ON

MAINT

VEH

+ -

OFF

Illustrations for information only, symbology not representative

c)

The WEIGHT format displays performance versus weight/flight envelope:

332 C1e 11.102.01 E

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3.5.2.2.3.

System ancillaries and maintenance In addition to the information available in flight, the following pages can be displayed on the ground only: a)

Display of maintenance data (MAINT for mat)

List of tests involved in failure detection and LRU/LRM concerned including Sensors, Ancillary Units, EID AMC, ADC, FMS and AFCS status

b)

Configuration management (CONFIG format)

Unit selection, selection.

c)

Configuration

Uploading and Downloading on PC (PC format) allows exchange of data recorded during flight.

All the controls are under PC management.

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d)

Test and maintenance page (Test f ormat) This mode selected by the crew includes:

-

3.5.2.3

A process to test displays softkeys. A display pattern to test symbol generator. The reference part number of EID and AMC. A display of active pin program of EID and AMC. The reference of the maintenance table.

VMS archit ecture Inside the computer signal conditioning is simplex but independent for each signal so that a single component failure results in the loss of at most a single parameter After conditioning, electrical signals are received and processed by the two processing modules and the results, exchanged through the cross-talk links, are compared by the two processors. In case of discrepancy a failure message is displayed. In case of one module failure, all the functions remain available on the other module.

VMS architectur e

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3.5.3

Auto matic Flight Contro l Subsyst em The 4-axes AFCS of the AS332C1e is based on the Avionique Nouvelle AFCS family concept, designed for the new Eurocopter range from light single engine to heavy twin engine helicopters. The AFCS comprises two Autopilot Modules which are linked to the MFDs and controlled by several stick buttons, an AFCAU (Automatic Flight Control Auxiliary Unit) for APM and Trim engagement/disengagement, and two FCP (Flight Control Panel) located within close reach of each crew member for upper mode engagement or reference modification.

AFCS Flig ht Contr ol Panel

AFCS Flig ht Contr ol Panel wi th SAR mod es (i n o pt io n)

AFCAU fr on t p anel Each APM includes 2 separate processing units which self-monitor themselves (dual fail-passive architecture). In case of failure of one APM, the other takes full control over the actuators and maintains full operationality (fail-operative feature).

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3.5.3.1

AFCS fun ctions The digital computer performs:

3.5.3.1.1.



Flight control (basic stabilization and upper modes)



Sensor monitoring and pre-flight test



Flight envelope processing



Integrated maintenance

Flight Control Function Upon AFCS engagement, the default mode of operation (called basic stabilization) includes: 

On the pitch and roll axis : attitude hold



On the yaw axis : heading hold with turn coordination

It also includes automatic piloting-help adaptation to hand-on/feat-on flight and axes decoupling. On the basic version of the digital AFCS, the following modes are available: 

IAS

:

Airspeed acquisition and hold



ALT

:

Pressure altitude hold



ALT/A

:

Selected pressure altitude acquisition



V/S

:

Vertical speed acquisition and hold

CR/HT

:

Capture and retention of radio height



HDG

:

Selected heading acquisition and hold



LOC

:

Acquisition and hold of a localizer beam



G/S

:

Glide slope capture and hold



NAV

:

Coupling to the navigation computer



VOR

:

VOR navigation



GO-AROUND :



Missed approach mode corresponding to the acquisition and hold of a self climbing flight path with collective and pitch axes

Moreover, for SAR operations, specific optional modes are available: 

F/TDN

:

Automatic transition to the hover



TUP

:

Automatic transition from the hover



HT/HOV

:

3D Hover hold



GSPD

:

Ground speed acquisition and hold

CRHT

:

Hold of radio altitude on collective axis



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3.5.3.1.2.

Monitoring function The digital computer of each half-system monitors the data received from the following sensors, and raises an appropriate alert to the crew in case of discrepancy between dual sensor data or detected failure: 

The 2 AHRS's and ADC's for attitudes and air data



The aircraft / engine sensors used to compute the flight envelope



The 2 radio altimeters for radio height (applicable only with the optional second radio altimeter)

In addition, the AFCS uses ISI sensors as a third data source to improve safety and operational availability according to MMEL (Master Minimum Equipment List) requirement in case of AHRS failure.

The RCU (Reconfiguration Control Unit) permits to select a precise unit in case of failure of the nominal one:

Front panel of the basic RCU

3.5.3.1.3.

Flight envelope function The purpose of this function is to provide the crew with: 

The VNE



The VTOSS



The best climb speed

These information are calculated with the aircraft weight which is initialized at take-off by the crew, and then updated during the flight, according to the f uel consumption.

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3.5. 3.5.3. 3.2 2

AFCS Contro l and Reconf Reconf igur ation AFCS modes engagement engagement is operated from two AFCS Control Panel located on the instrument panel (on pilot side and on copilot side), and from cyclic and collective sticks grips and from AFCAU. The AFCS sub-systems can be reconfigured whether automatically upon failure detection or upon crew selection

3.5.4 3.5.4

Primary referenc e sensor sens ors s The AHCAS is fitted with: 

Two digital ADC ADC which compute compute standard and corrected altitude, indicated indicated airspeed, airspeed, true airspeed, static air temperature, vertical speed

Two digital AHRS, using FOG technology, which compute magnetic heading, attitudes, attitude rates, body acceleration, acceleration, baro-inertial baro-inertial vertical speed



The so called primary reference sensors provide data directely to each MFD and each APM for safety reasons. reasons.

Air Data Computer Computer

Attitude and Heading Heading Reference Reference System

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3.5. 3.5.5 5

Customi zed zed sensor s and peripherals (Opti (Opti ons ) Thanks to open architecture and use of standard interface, the AHCAS can accommodate various type of optional subsystem or equipment, according to particular mission needs and customer requirements. The customizable system parts associated with AHCAS are the following.

3.5. 3.5.5. 5.1 1

Radio Radio navigatio n sensors The AHCAS provides display of radio-navigation data issued from a customized radionavigation set. Each MFD of the FDS receives radio-navigation data in A429 form to display on FND or NAVD formats, from various sensors, for example: VOR/ILS  DME  DF or ADF  TCAS 

Some of the radio-navigation data are forwarded to the AFCS to operate the relevant upper modes. The customized set of radio-navigation equipment (optional) is described in a separate chapter.

3.5.5.2 3.5.5.2

Weather Radar The AHCAS can process and provide display of weather radar data received from a standard A453 equipment or in video form.

3.5. 3.5.5. 5.3 3

Flight management management Subsyst em The AHCAS can interface a customized FMS, according A429 IFDS or A702 standard, with its own sensors (Doppler, INS or GPS). This subsystem can be more or less powerful and complex, depending on customer requirements. Each MFD of the FDS provides basic data to the FMS through multipurpose concentrated data A429 lines and receives navigation data (track, ground speed, distance-to-go, cross track, route waypoints, etc…) from the FMS in A429 form to display on FND or NAVD formats.

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Blank

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3.6. Power plant 3.6.1

General The power unit consists of two Turbomeca Makila 1 A 1 free turbine engines mounted side by side above the fuselage. The engines are installed ahead of the main gearbox in two totally independent compartments. The engines are identical and interchangeable after turning the tail pipe. Each one transmits power to the main gearbox directly through a diaphragm type coupling.

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3.6.2

Turbine engin es

3.6.2.1

Description

The Makila turbine engine comprises from the front to the rear 

an air intake casing



a 3-stage axial compressor, with titanium wheels



a centrifugal compressor, with titanium wheel



a combustion chamber, of the annular type. Centrifugal fuel injection through an injection wheel insures, in a simple way, correct pulverization at all engines speeds under a moderate injection pump pressure.



a 2-stage high pressure turbine (which drives the compressor)



a 2-stage high pressure turbine which drives the output power shaft.

the power take-off located at the rear which rotates at a nominal speed of 22,850 rpm.





an exhaust pipe which may be directed to the left or the right according to the engine location (RH or LH).

Each engine assembly constitutes a self contained unit comprising all the systems, equipment and accessories required for its operation, mainly :

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

1 fuel system Each engine is fitted with a fuel filter including a clogging indicator, besides the filters provided on the aircraft on the fuel supply line. The fuel inlet is borated at the platform level and the connections to the engines are made through a fire resistant hose equipped with a quick-disconnect self sealing coupling. The fuel is heated through a heat exchanger by cooling the oil of the engine.



1 governing system This system does not require any action from the crew on the fuel flow (no throttle twist grip on the collective pitch lever). One of the functions of this governor system is to maintain a constant free turbine r.p.m. between maximum and minimum demands. Through its design, it ensures an even load distribution between the two engines without requiring the addition of a matching system. In flight the pilot controls the power through rotor pitch only. Other functions ensured by the governor system : .

Limitation of the gas generator r.p.m. to its maximum value in case of free turbine overloading

.

Optimum acceleration of gas generator when pitch is increased ; from idle r.p.m. maximum power is reached in less than 3 seconds.

.

Insensibility to a booster pump pressure drop

.

Fuel metering through throttle lever, in case of governor system failure. Thus, engine power is not lost in the event of governor failure.

.

Automatic starting of gas generator and governor idle r.p.m.

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These functions are possible through : .

An electronic control box, located in the electrical cabinet, for the free turbine governing and automatic starting sequence.

.

A hydro-mechanical governor unit, located in the engine accessories area. This unit uses only fuel as a servo-mechanism operating fluid.

The electronic control box serves the free turbine r.p.m. and gives a signal, corrected for the collective pitch level position, which is then used in the gas generator r.p.m. sharing channel. A warning light, located in the cockpit, warms the pilot of a governor system defect, but the system being redundant, flight may be continued. The hydromechanical governor system meters the fuel flow according to the gas generator r.p.m. and the reference value set by the electronic governing system. An acceleration control system allows engine acceleration in good conditions. A deceleration control system prevents engine flame-out when power is decreased suddenly. 

1 lubricating system

The Makila 1 A 1 engine is provided with a self-contained lubrication system with integral oil tank and cooling system. The main components are : .

One oil tank located in the engine air intake casing (8 liters), with sight gauge

.

One gear type oil pump

.

One filter, complete with by-pass valve

.

Four scavenge pumps, each sucking oil, from the different engine modules. Each pump includes a strainer and a magnetic plug.

.

One heat by changer (oil is cooled by the fuel supplied to the engine)

In flight monitoring is ensured by : . . 

temperature and pressure indications filter clogging indication

1 engine anti-icing system making use of engine lubricating oil flow through an oil/fuel heat exchanger

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3.6.2.2

Modular conc eption The Makila turbine engine is of modular design. It is composed of 5 modules plus the exhaust pipe.

The two Makila engines are installed side-by-side at the top of the fuselage, forward of the main gear box, in two separate compartments. Titanium fire-walls are provided between the engine and main gearbox compartment. The platform supporting the engines is also titanium. The compartments are enclosed by hinged cowlings. The engines are attached : 



3.6.2.3

on one side, to the main gear box through the gimbal joint tube which absorbs the axial loads and the torque on the other side, to the structure through struts, with spherical bearings, which absorbs the transverse and vertical loads.

Engine/main gearbo x coup lin g The engine power is transmitted to the main gearbox through a shaft having a flexible coupling at both ends. On the main gearbox side, the drive flange slides on the input gear shaft. The coupling shaft is housed in a coupling tube equipped with a gimbal joint located at mid-distance between the flexible couplings to allow for slight engine/input shaft axis misalignment.

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3.6.2.4

Engine cont rol s The fuel flow levers and the shut-off cocks are grouped together at the front of the overhead panel.

1 2 3 4

3.6.2.5

LH engine fuel flow control RH engine fuel flow control RH engine fuel shut-off cock control LH engine fuel shut-off cock control

Engine rati ngs Power per engine, in standard atmosphere, at sea level :

-

Maximum emergency power Take-off power Intermediate emergency power Maximum continuous power

kW 1,400 1,357 1,330 1,185

ch 1,902 1,845 1,807 1,610

332 C1e 11.102.01 E

shp 1,877 1,819 1,783 1,588

3.6-6 ®

®



3.6.2.6

Fire-exting uis hing syst em

3.6.2.6.1.

Detection system The engine compartments are protected by a dual fire detection system comprising a serie of detectors fitted to each engine. A logic circuitry and a warning system on the instrument panel, complete the installation.

3.6.2.6.2.

Extinguishing system The extinguishing system includes two interconnected freon-filled fire bottles, each equipped with dual percussion heads and supplying two discharge lines which are routed into the engine compartments. Push-buttons allow the crew to fire and discharge the selected system bottle. An emergency button allows the opposite fire bottle to be operated, to provide a second shot. Ventilation of the engine compartment is normally provided by ram air via opening into the engine intake cowling.

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In the event of fire, this ventilation is automatically "cut-off" by activation of blanking flaps which are operated either from the fuel "shut-off levers" or from the emergency electrical "cut-off" handle.

1 2 3 4 5 6 7 8

3.6.2.6.3.

"Engine 1 fire" indicator lights "Engine 2 fire" indicator lights Engine 1 "normal" percussion push-button Engine 1 "emergency" percussion push-button Engine 2 "normal" percussion push-button Engine 2 "emergency" percussion push-button Fire extinguisher bottle n°1 Double check valve

9 10 11 12 13 14 15 16

"Emergency" percussion head Fire extinguisher bottle n°2 Pressure gauge "Normal" percussion head Freon distribution line Freon diffuser "Fire extinguisher bottle n°1" discharged "Fire extinguisher bottle n°2" discharged

Operation of the fire extinguisher system In case of fire in one of the engine compartments, the fire detectors cause the relevant red "engine fire" indicator light to come on. The pilot has then to press the corresponding "normal" (1st shot) percussion push button which causes the fire extinguisher bottle n°1 to discharge its freon load in the engine compartment. The amber "fire extinguisher bottle n°1 discharged" indicator light comes on. Should the fire extinction not be achieved with the fire extinguisher bottle n°1, the pilot has the possibility of a second shot by pressing the "emergency" percussion push button which causes the fire extinguisher bottle n°2 to discharge. The amber "fire extinguisher bottle n°2 discharged" light indicates the achievement of the second bottle dischargment.

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3.6.2.7

Engine lub ric ants

Normal lubricant French U.K. U.S. Specifications Specifications Specifications -

DERD 2499

MIL L 23699

NATO Symbols

Remarks

0.156

Synthetic Oil

Al ter nat iv e lu br ic ant s French U.K. U.S. Specifications Specifications Specifications

NATO Symbols

-

0.148

-

MIL L 7808

Remarks

Synthetic AIR 3514

-

3.6.2.8

DERD 2497

0.150 -

Oil

0.160

Engine washing facility without cowling s opening Following a flight in sand or salt laden atmosphere, it is necessary to rinse the engine compressors. To ease this task, a built-in engine wash system is provided to permit washing of the engine compressor sections with engines at idling speed, rotor turning, but without the need to open engine cowls. The installation comprises a spray ring integral to each engine air intake and equipped with a self-sealing coupling. A plumbing links this self-sealing coupling on each engine to an externally mounted self sealing connection located on the LH side of the helicopter fuselage, where a ground power wash rig may be connected.

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The surplus liquid is evacuated through the engine exhaust dump valves.

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3.7. Fuel system 3.7.1

General descrip tion The fuel system of the Super Puma comprises 5 fuel tanks housed in the fuselage bottom structure. The fuel is contained in two independent groups of flexible cells. The total fuel capacity is 1,537 litres (407 US gal.), 1,517 litres (402 US gal.) of which are usable. Each fuel tank group feeds the engine located on the same side. The RH and LH longitudinal tanks include a fuel gauge incorporating high level and low level switches, 2 booster pumps, a drain cock and a water bleed. Group 2 RH longitudinal tank Forward transverse tank -

Group 1 - LH longitudinal fuel tank Rear transverse tank Rear tank The forward and rear transverse fuel tanks include a fuel gauge with a high level switch and a jet pump, which boosts the fuel into the longitudinal tanks. The rear fuel tank is fitted with a fuel gauge with a high level switch ; the fuel contained in this tank flows into the rear transverse tank by gravity. The bodies of the fuel gauges are embrittled at the bottom the prevent the fuel tank walls from being perforated in the case of a crash landing. Each fuel tank is provided with its own air vents composed of hoses capable of absorbing structural distortions during a crash landing. A roll-over valve is installed on each hose. 332 C1e 11.102.01 E

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Fuel can be transferred from one tank group to the other through a cross feed system including a two-way pump which is controlled from the instrument panel.

Fuelling is performed through two ports located on the RH side of the aircraft, connected to the central and rear transverse tanks. A socket near each port enables the aircraft to be earthed during refuelling. The controls and monitoring instruments of the fuel system are operated from the cockpit (instrument panel-pedestal). Two filter clogging mechanical indicators are located in the cabin and can be inspected through a sight fitted on each filter case cover.

1. 2. 3. 4. 5. 6. 7.

Fuel shut off cocks Low pressure switch Pressure transmitters Check valves Filter by-pass Filters Clogging differential pressure switch 8. Gauges 9. Jet pumps 10 .Check valves 11 .Transfer pump 12. Check valve 13 .Booster pumps Fuel s ystem diagram

Fuel s ystem diagram

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1 2 3 4 5 6

3.7.2

Filling points RH longitudinal tank Forward transverse tank LH longitudinal tank Rear transverse tank Rear tank

Usable fuels Désignation

French Specifications

U.K Specifications

U.S. Specifications

NATO Symbols

KEROSENE* F34 KEROSENE F35 JET A1 KEROSENE JET A WIDE CUT* JP4 - TR4 WIDE CUT JP4 (AVTAG) WIDE CUT

AIR 3405 F34 AIR 3405 F35

D ENG RD 2453

MIL-T-83133 JP8 ASTM-D-1655 Jet A1

F34

ASTM-D-1655 JET A MIL-T-5624 GRADE JP4 MIL-T-5624

-

HIGH FLASH POINT TR5-F43 HIGH FLASH* POINT TR5

AIR 3404 F43 AIR 3404 F44

AIR 3407

D ENG RD 2494

D ENG RD 2454 D ENG RD 2486

F35

F40

ASTM-D-1655 JET B D ENG RD 2498 AVCAT D ENG RD 2452 AVCAT

F43 MIL-T-5624 GRADE JP5

F44

Note : Use suffixes and amendments in force. * Fuels including anti-icing additive.

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3.7.3

Optional additio nal fuel tanks The Super Puma can accommodate, as optional equipment :    

2 external fuel tanks 1 central auxiliary fuel tank 1 cabin compartment fuel tank 2 groups of ferrying fuel tanks installed in the cabin

CAPACITY OF THE OPTIONAL FUEL TANKS External fuel t anks

2 x 318 litres 2 x 83 US gal. 2 x 251 kg 2 x 553 lb

Central f uel tank

1 x 321 litres 1 x 84 US gal. 1 x 254 kg 1 x 560 lb

Ferryin g fuel tanks 5 x 475 litres 5 x 126 US gal. 5 x 375 kg 5 x 826 lb

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3.8. Transmission system

3.8.1

General descrip tion The transmission components conveying torque to the rotors essentially comprise shafts and gearboxes (main, intermediate and tail gearboxes).

Engine to main rotor transmission is provided by : 

Two coupling shafts conveying the motion of the engine free turbines to the main gearbox



The main gearbox



The main rotor head with integral hub.

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3.8.2

Main gearbox

3.8.2.1

Description The main gearbox is attached to the transmission deck through 3 suspension bars and a flexible mounting plate. The main gearbox includes 5 stages reducing the engine output shaft speed (22,841 r.p.m.) to : -

265 r.p.m. for the main rotor 4 888 r.p.m. for the tail rotor drive

The MGB is provided with two free wheels whose drive shafts are fitted with torquemeters. The MGB drives all the accessories required for the ancillary systems (2 alternators, 2 hydraulic pumps, oil cooler fan and rotor brake). It is equipped with the following monitoring equipment : 

Dual system oil filter



Magnetic plug



Oil sight gauge



Oil sampling port for analysis purposes



Oil pressure taps and contactor



Ports for endoscopic inspection of the power system pinions (A, B, C)



Oil temperature probe connectors

The main gearbox includes a rotor brake enabling the rotor to come to a complete stop within about 12 to 15 seconds.

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The main gearbox transmits the following power : -

With 2 engines running : .

-

2,235 kW

3,040 ch

3,000 shp

1,550 kW

2,105 ch

2,080 shp

With 1 engine running : .

3.8.2.2

Max. power

Max. power

Main gearbo x cool ing syst em The MGB includes two oil systems consisting of a main oil system and an emergency oil system. In the main system, the oil is cooled by way of a pump which sucks the oil from the bottom of the MGB casing and forces it to the oil cooler. The oil is then ducted into the MGB through a filter. The fan on the cooling unit is driven by a shaft which is directly connected to a main gearbox power take-off. In the emergency system, the oil is filtered but not cooled. A pump delivers the oil via the filter to the jets.

1

2 3 4 5 6 7 8 9

rotor shaft lubrication line strainer Oil pressure warning filter by-pass MGB filter main oil pressure indicator oil cooling malfunction fan } cooling oil cooler } unit

10 11

12 13 14 15 16 17

oil cooler relief valve emergency system oil pressure indicator magnetic plug MGB chip warning main oil pump emergency oil pump oil temperature indicator oil temperature warning

MGB oil syst em diagram

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The cooling unit is located under the sliding cowling. It maintains the temperature of the MGB oil below the maximum authorized value for ambient air temperatures of up to 50°C.

1 2 3 4 5

3.8. 3.8.2. 2.3 3

Fan drive shaft Fan protective grid Fan duct Heat exchanger Fan exhaust

Fire detection syst em in the MGB MGB comp artment The MGB compartment is provided with a fire detection system composed of 14-fire detectors. detectors . Fire warning is given on the warning panel ; a test light is installed on the overhead panel. Al l th e vi tal co mp on ent s such as servo control levers, horizontal transmission shaft, tail rotor control cable and hydraulic and electric circuitry have been designed to resist fire. fire.

In the MGB compartment, the fire detectors are arranged in pairs : one detector for MGB system 1 and one for MGB system 2. Both systems monitor the same sensitive points in the MGB compartment.

Critical points monitored monitored by the fire detector in the MGB compartment : 1 2 3 4

detector for ventilation air outlet detectors for hydraulic pumps & alternators detectors for RH & LH hydraulic hydraulic units detectors for servo controls

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3.8.3 3.8.3

Tail Tail driv dr ive e Power transmission to the anti-torque rotor is achieved from the main gearbox rear reduction gear through the following : 

the horizontal horizontal drive shaft assembly assembly divided into seven seven separate sections. The front and rear shafts have no bearing blocks. The other shafts are each supported by a self-lubricated bearing. The shafts are connected by flexible couplings. They are interchangeable and factory balanced.



the intermediate intermediate gearbox gearbox is is situated at the tail boom/pylon boom/pylon junction junction and and changes changes the drive angle by 40°. The reduction obtained is down from 4,888 r.p.m. to 3,751 r.p.m.



the oblique shaft



the tail gearbox, gearbox, changing changing the drive angle angle by 92°. This reduces reduces the speed speed from 3,751 r.p.m. to 1,279 r.p.m. and drives drives the tail rotor rotor shaft.

1 2 3 4 5 6 7 8

Front shaft Dual bearing support Central shafts Bearing supports Rear shaft Intermediate gearbox Oblique shaft Tail gearbox

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Intermediate gearbox

Tail gearbox

The intermediate and tail gearboxes are splash-lubricated. Each gearbox is fitted with a level sight, a filler plug, a drain valve with a magnetic plug, a temperature probe connector and a port for an endoscopic check.

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3.8.4

Gearbox lubr icants

3.8.4.1

Oil capacities of gearboxes Gearbox oil capacity

3.8.4.2

MGB

24 litres

6.3 US gal (cooling unit included)

IGB

0.75 litre

0.2 US gal

TGB

1.5 litre

0.40 US gal

Oil for main gearboxes plus main rot or head hing es 



Normal lubricant French Specifications

U.K Specifications

U.S. Specifications

NATO Symbols

Remarks

-

DERD 2499

MIL L 23699

0.156

Synthetic oil

Alternative lubricants French Specifications

U.K Specifications

U.S. Specifications

NATO Symbols

Remarks

-

DERD 2497

-

0.160

Synthetic oil

AIR 3514

-

-

0.150

Synthetic oil

-

-

MIL.L.7808

0.148

Synthetic oil

AIR 3525

DTD 581 C

MIL.L.6086

0.155

Mineral oil

Note : Use suffixes and amendments in force.

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Blank

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3.9. Rotors The main rotor consists of a rotor head and 4 blades rotating clockwise (viewed from above) at a rated speed of 265 r.p.m.

3.9.1

Main rotor head It consists of a four-arm hub integral with a vertical shaft. Each arm includes flapping/drag and feathering hinges (interchangeable modules). The drag hinge is linked to a viscoelastic frequency adaptor. The rotor head is equipped with droop restrainers to limit the rotor blade flapping downwards and with coning restrainers making easier rotor spinning in high wind conditions. The rotor head is also fitted with a self-contained lubricating system including four oil tanks sight gauges. At the bottom of the rotor shaft, a strainer prevents the main gearbox from being contaminated by metal particles. In association with this strainer, a magnetic plug is used to detect particles. Each blade pitch change and flapping/drag hinge is fitted with a magnetic plug at its lowest point. A dome shaped fairing on the main rotor hub is used to reduce the rotor head drag.

The main rotor head hinges are lubricated either with NATO O.156 (MIL-L 23699) or NATO O.155 (MIL-L 6086) synthetic oil. Alternative oil NATO O.148 (MIL-L 7808).

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3.9.2

Main rotor blades The rotor blades, with evolving profile, mainly consist of a fiber glass roving spar, carbon and glass fabric laminated skins, with moltoprene and honeycomb fillers. Leading edge protection against erosion is afforded by stainless steel capping. The trailing edge is reinforced by a carbon-fibre strip. Each blade ends in a removable tip cap enclosing the balance weights. The blades are each secured by 2 pins to the rotor head. 

length :

7.00 m (22.96 ft)



chord :

0.60 m (1.97 ft)



airfoil section :

cambered with tapeting thickness ratio



twist :

from 0 to -8.74° over 5.65 m (18.53 ft)



paint finish :

dark green on upper and lower surfaces

The main rotor blades are individually interchangeable.

332 C1e 11.102.01 E

3.9-2 ®

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3.9.3

Tail rotor The five-blades tail rotor is attached to the tip of the pylon, on the right hand side. It rotates clockwise (viewed from the left of the aircraft) in a vertical plane at a rated speed of 1,279 r.p.m. The tail rotor hub is integral with the rotor shaft. The rotor, articulated in the flapping plane, is controlled in collective pitch setting by a spider.

3.9.4

Tail rotor blades The tail rotor blades are of composite construction and consist of the following :     

a glass roving spar a glass fibre skin a titanium leading edge a foam filler a tip balance weight support

The rotor blades are individually interchangeable. 

length :

1.22 m (4.00 ft)



chord :

0.20 m (0.66 ft)



airfoil section :

curved and tapered



paint :

dark green

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332 C1e 11.102.01 E

3.9-4 ®

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3.10.

3.10.1

Hydraulic system

General descrip tion The hydraulic installation includes two fully independent main hydraulic generation systems, and a third independent power source used for landing gear emergency operation only. 

The main RH power generation system (one main gearbox driven pump), supplies the lower section of the main servo controls body, and the RH section of the tail servo-control body.



The main LH power generation system (one main gearbox driven pump) supplies the upper section of the main servo controls body, the LH section of the tail servo-control body, and the ancillary hydraulic units which incorporate the following items : the autopilot, the landing gear, the rotor brake, the wheel brakes and the hoist *. In addition, the main hydraulic system incorporates a D.C. supplied auxiliary electrical pump. It is used by the pilot or maintenance crew to operate the flight controls prior to rotor engagement. It is also used as a back up for some consumers of the LH hydraulic system in case of failure of the main LH pump and for landing gear extension in order to reduce relevant time. The auxiliary pump also supplies hydraulic power to permit rigging of the flying controls and related adjustments of the autopilot. It is also used to charge up the parking brake accumulator.



A landing gear emergency power source is provided. This system includes an independent hydraulic fluid reservoir and an electrical pump which powers the landing gear struts extension chambers to provide landing gear emergency extension operation.

The main (LH) and RH systems have their operating and monitoring equipment grouped into independent self contained units. The LH group incorporates the auxiliary electrical pump and its own monitoring system. The operating pressure of each system is 175 bars (2500 psi). This optimum pressure is supplied by the two self-governing main gearbox hydraulic pumps. The LH pump provides a flow of 27 litres/min (7.13 US gal/min) and the RH pump 12 litres/min (3.17 US gal/min).

332 C1e 11.102.01 E

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Each system includes a reservoir with a useful capacity of 8 litres (2.11 US gal.) including a sight level, a filter and a low level sensor. * Optional equipment

The pipes are protected by stainless steel sheaths on the RH system and by stainless steel sheaths plus silicone sheaths on the LH system. The ancillary system is connected to the LH system through rigid pipes routed outside the cabin and protected by a gulley. Pressure gauges on the instrument panel and warning lights on the warning panel are provided for the hydraulic generation system monitoring. For ground testing, two external self-sealing connectors (power receptacle) are provided on the LH system, and two internal ones on the RH system.

1 2 3 4 5 6

filler cap strainer (100 microns) low level sensor vent hydraulic fluid reservoir fluid level sight

7 8 9 10

self-sealing union for unit pressure line self-sealing union for unit suction line drain plug hydraulic accumulator

332 C1e 11.102.01 E

ground

power

ground

power

3.10-2 ®

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3.10.2

Hydraulic generatio n diagram

3.10.3

Servo uni ts syst em The aircraft is controlled by 4 hydraulic servo-units, three active on the main rotor and the fourth on the tail rotor control shaft. Each servo unit is of the dual tandem cylinder type. Each cylinder, supplied from a separate hydraulic system, is controlled through two distributor valves with seizure warning. The RH hydraulic system supplies the lower cylinders of the main servo units and the RH cylinder of the tail servo unit. The LH hydraulic system supplies the upper cylinders of the main servo units and the LH cylinder of the tail servo unit. The RH system is "simple" because used only for the servo-units. It is less vulnerable than the LH system which, further to the servo units, is used to supply other hydraulic systems. The servo-unit system is supplied by 2 pumps. In the event of a single pump failure, there is no consequence on the operation of the system since the remaining one is sufficient to supply one cylinder of each servo-unit.

332 C1e 11.102.01 E

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In the event of leakage on the hydraulic system itself, the servo-units system is automatically automatically isolated.

332 C1e 11.102.01 E

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3.10 3.10.4 .4

Routing of hydr aulic cir cuit s

3.10 3.10.5 .5

Hydraulic flu id specifi cations Normal fluid French Specifications

U.K Specifications.

-

U.S. Specifications

NATO Symbols

Remarks

MIL.H.83282

H.537

Synthetic oil

Remarks

Al ter nat iv e fl ui d French Specifications

U.K Specifications.

U.S. Specifications

NATO Symbols

AIR 3520

DEF STAN 91 48/1

MIL.H.5606

H.515

Mineral oil

GRADE OM 15 Note : Use suffixes and amendments in force.

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332 C1e 11.102.01 E

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3.11. 3.11.1

Electrical system

General descrip tion In addition to GPU receptacles (115VAC 400Hz and 28VDC), the electrical installation mainly includes tw o 20/30 KVA, 400 Hz alternators driven by the main gearbox . It comprises : 



Fed by alternator no.1 .

A thr ee-phase A.C. generation netwo rk delivering 115/200 volts, 400 Hz

.

A single phase A.C. generation network delivering 115 volts, 400 Hz. This single-phase systems is derived from the above 3-phase system.

Fed by alternator no.2 .

A thr ee-phase A.C. generation netwo rk delivering 115/200 volts, 400 Hz

.

A 28-vol t D.C. netwo rk through 2 transformer-rectifiers. This network charges the aircraft battery (43 A.hr) and the standby battery.

Power distibution scheme 332 C1e 11.102.01 E

3.11-1 ®

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

Fed by a st andby battery : .

Emergency supply of the standby horizon, emergency exits and ordinance lights.

Each network is fitted with the various accessories required for the protection and governing of the circuits. In case of failure of one of the alternators, an automatic switchover system caters for the power supply to the A.C. and D.C. networks from the other alternator. Safe, reliable operation of the vital components is ensured by several protective devices : Shorting of the bus bars : automatic switching device isolating the shorted bus bar on the D.C. power system and supplying the vital equipment items through the bus bar remaining available on the A.C. power system. 

Battery network : disconnection of the battery from the structure (grounding suppressed) through differential protection of the electrical control in case of crash landing. This network is specially routed and the cables are fire-proof. 



Ground po wer receptacles :

. .

D.C. voltage surge and drop A.C. voltage surge and drop and phase and frequency inversion

All simple cases of failure are covered by circuit redundancy. All the circuits are composed of high temperature-resistant cables coated with kapton insulator. The control protection and monitoring boxes of all circuits are grouped in the LH cabinet separating the cockpit from the cabin. The battery is located in the aircraft nose. The standby battery is housed in the aircraft aft section.

332 C1e 11.102.01 E

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The A.C. and D.C. networks are monitored through 2 voltmeters and 2 ammeters on the cockpit overhead panel.

2

selector-switches,

AC p ow er s ys tem co nt ro ls and in di cat or s

332 C1e 11.102.01 E

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DC power system contr ols and indicators

1 2 3 4 5 6 7 8

-

Copilot's "general cut-out handle" Pilot's "general cut-out handle" Control switches Decoupling indicator lights Battery temperature indicator Dual voltmeter – Ammeter Ammeter selector switch Volmeter selector switch

332 C1e 11.102.01 E

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3.11.2

Electr ical syst em bloc k diagram

332 C1e 11.102.01 E

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3.12.

Exterior lighting

The aircraft is provided with 3 position lights (2 side position lights on the sponsons (red on LH side, green on RH side), one white position light at the top of the tail pylon), a red anti-collision light at the top of this pylon, together with a white 600 W landing light, under the nose of the aircraft. The landing light is retractable, and adjustable in azimuth over 360° and 120° in elevation.

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332 C1e 11.102.01 E

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3.13. 3.13.1

Air data system

General The Air Data System is composed of 3 independent circuits. A pilot circuit, A copilot circuit, An emergency circuit. These circuits are connected to two dual pitot tubes including each:  

2 total pressure ports (T1 and T2) 3 static pressure ports (S1, S2 and S3)

1 - ADC 1 2 - Cockpit static pressure switches 3 - RH pitot tube (pressure port) 4 - LH pitot tube (pressure port) 5 - Shut-off valves 6 - ADC 2

Ai r Dat a System lay-ou t

332 C1e 11.102.01 E

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3.13.2

Pilot cir cuit The pilot circuit includes: 

The RH total pressure port T1



The RH and LH static pressure ports S2

The total and static pressure systems are connected to the transducers of the ADC No.2 which transmits pilot system air data to the EFIS screens (4)

3.13.3

Copilot cir cuit The copilot circuit includes: 

The LH total pressure port T1




The total and static pressure systems are connected to the transducers of the ADC No.1 which transmits copilot system air data to the 4 EFIS screens

3.13.4

Emergency cir cui t The emergency circuit includes: 

The RH total pressure port T2




These ports are only connected to the Integrated Standby Instruments (ISIS). In addition, 2 shut-off valves make it possible to switch the six static pressure ports (2S1, 2S2, 2S3) over the cabin stand-by alternate pressure ports.

1 - Pilot shut-off valve 2 - Copilot shut-off valve 3 - Integrated Standby Indicator

Ai r Dat a Sys tem

332 C1e 11.102.01 E

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3.13.5

Pilot heating

3.13.5.1

Pilot sys tem It comprises: 

A pitot heater supplied by the secondary DC system,



A pitot tube support arm heating supplied by the AC system,







3.13.5.2

A relay, allowing the pitot and its tubing heating systems to be energized, which is controlled by an ON/OFF switch mounted on the sub-panel in the cockpit, A sub-panel mounted "AMBER" failure warning light is controlled by a relay which monitors pitot heating system continuity, A "weight on wheel" protection circuit reduces by half the power required for pitot heating by operating a relay whilst the aircraft is on the ground.

Copilot system Similar to the pilot system.

3.13.5.3

Emergency sys tem Identical in design to the two systems above. The only difference consists in the pitot heating being supplied through the essential DC system. This maintains pitot heating during flight with total loss of alternators (power supply is through battery only). Note

3.13.5.4

: Each pitot failure warning light is repeated on the warning panel by a single amber light.

Failur e of LCD ind ication In the event of total LCD indication failure, airspeed and altitude data remains available on the Integrated Standby Instrument mounted on the instrument panel.

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3.14.

Pneumatic system

Compressed air (P2) is bled from the engine compressors. The whole system is intended to operate the heating system and the following optional equipment :



the sand filters the de-icing system on the horizontal stabilizer.

1 2 3 4

Heating system Inflatable seal of the engine air intakes Air intake bullet inflatable seal (option) Inflatable leading edge of the de-iced horizontal stabilizer (option)



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3.15.

Ground handling and picketing

Various points are provided on the aircraft fuselage for ground handling and for maintenance tooling :



4 mooring rings A

:

- 2 to be fitted on frame 1715, on each side - 2 to be fitted in main landing gear axles



2 rough weather mooring fittings B :

fitted on frame 6815



Gripping points C

:

To tie down the main blades 1 on each main landing gear wheel axle C



Jacking points D

:

3 under the aircraft



De-booging poin ts C

:

1 on each main landing gear unit

Nota : 3 MGB fixing points can be used for lifting with a sling or crane (with special tooling, to be contracted separately) The main maintenance tooling is fitted using : 

9 fixtures fo r the cranes (3 installation possibilities)



6 fixtur es for th e maintenance ladder (2 on each side of the fuselage and 2 on the pylon, starboard side)



2 fittings on the nose landing gear for towing bar (towing bag in option : to be contracted separately)

332 C1e 11.102.01 E

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D : Jacking points

332 C1e 11.102.01 E

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3.16.

Structural prov isions for optional equipment

(Drilled holes, doublers, system tappings, recesses etc.)

Structural provisions are provided in the aircraft for the following optional equipment : 

Fixed hoist



External and central auxiliary fuel tanks



Passenger seat installation



Casualty-carrying installation



External load-carrying (sling)



Emergency floatation gear



Ferrying tanks



Radio com./nav. antennae : VHF 1 and 2, VOR/ILS, radio-compass marker and radio altimeter



Utility seat installation



Orange screens for instrument-flying training

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3.17.

Airborne kit

The airborne kit delivered with each aircraft includes : 

6 static vent blanks



2 pitot head covers



1 engine air-intake grid protection cover



2 engine tail pipe blanks



4 mooring rings



2 rough-weather mooring fittings (included on the aircraft)



1 access ladder



1 data case



3 jacking ball joints



Main blade tie-down kit



Tail rotor blade lock



Fuel bleed line



1 stowing bag for the airborne kit

The weight of the airborne kit is not included in the empty weight of the aircraft.

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332 C1e 11.102.01 E

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3.18.

Accompanying literature

Each aircraft is delivered with the following relevant documents (in French or in English) : 

Flight Manual



Individual inspection log book, including : .

1 certificate of conformity

.

1 inspection log book complete with log cards raised for : -

those components (or the assemblies containing them) subject to a service life dependent on fatigue resistance

-

those components with unlimited service life when they are likely to be monitored in operation



Engine log books



Journey log book



Aircraft log book



Battery log book



Airworthiness Certificate (for those aircraft not registered in France, it is an "export" one testifying to the fact that the helicopter would have received its individual airworthiness certificate in France).

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4. Performance 4-1

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4-2

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4.1. 4.1. Main Main perfor perf ormanc mance e 4.1.1 4.1.1

Perf Perf ormance or mance on 2 engin es

Gross Weight 













Max. speed, VNE

Maximum cruise speed

Recommended cruise speed

Fuel consumption at recommended cruise speed

Fuel consumption at 70 kts

Rate-of-climb at 70 kts

lb

6,000 13,230

7,000 15,430

8,000 17,630

8,600 18,960

K m /h r

304

303

298

278

kts

164

164

161

150

k m /h r

283

281

278

262

kts

153

152

150

163

k m /h r

265

262

258

252

kts

143

141

139

136

k g /h r

481

486

497

502

l b /h r

1,060

1,071

1,096

k g /h r

335

353

374

389

l b /h r

739

778

825

858

m /s ec

14.7

12.1

9.8

8. 8 .2

f t /m i n

2,894

2,382

1,920

1,618

m

6,500

5,100

3,950

3,250

ft

21,325

16,700

12,959

m

5,850

4,450

3,200

2,300

ft

19,193

14,560

10,499

7,546

m

5,750

4,300

3,050

2,300

ft

18,865

14,108

10,007

7,546

m

5,200

3,550

2,100

1,400

ft

17,060

11 11,647

6, 6,890

4,593

1,107

Hover ceiling IGE (10ft) at take-off power



ISA

ISA + 20°C





kg

10,663

Hover ceiling OGE at take-off power





ISA

ISA + 20°C

4.1-1

332 C1e 11.102.01 E ®

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

Service ceiling Vz = 150 ft/min









ISA

ISA + 20°C

m

7,200

5,900

4,650

>2,895

ft

23,622

19,357

15,255

>9,500

m

6,500

5,100

3,800

2,895

ft

21,325

16,732

12,467

>9,500

km n.m.

670 362

668 361

661 357

639 345

km n.m

719 388

919 495

909 491

890 481

Maximum range (without fuel reserve, at recommended cruise speed) 

with standard fuel tanks



with external fuel tanks* *



with central auxiliary fuel tank* *

km n.m.

790 427

808 436

800 432

779 420



with external and central auxiliary fuel tanks* *

km n.m.

708 382

1054 569

1045 564

1030 556



with external, central and 4 ferrying tanks* *

km n.m.

657 355

1192 644

1721 929

1843 995

Maximum endurance (without fuel reserve, at 130 km/hr (70 kts)) 

with standard fuel tanks

hr : min

3:45

3:34

3:23

3:16



with external fuel tanks* *

hr : min

4:10

5:07

4:51

4:42



with central auxiliary fuel tanks* *

hr : min

4:28

4:21

4:07

3:59



with external and central auxiliary fuel tanks* *

hr : min

4:06

5:54

5:36

5:26



with external, central and 4 ferrying tanks* *

hr : min

3:48

6:43

9:30

9:57

* * Depending on the selected optional fuel tanks.

4.1-2

332 C1e 11.102.01 E ®

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4.1.2

Perf ormance on 1 engin e Performance in external load (maximum gross-weight of 9,350 kg – 20,615 lb) 



4.1.3

Rate of climb at 9,350 kg

6.4 m/sec 1,260 ft/min

Hover ceiling OGE at take-off power  ISA 

ISA + 10°C



ISA + 17°C

650 m 2,133 ft 300 m 984 ft Sea level

Perfo rmance in external load carryi ng miss ion



Rate of climb at 9,350 kg



Hover ceiling OGE at take-off power

6.4 m/sec 1,260 ft/min



ISA

650 m 2,133 ft



ISA + 10°C

300 m 984 ft



ISA + 17°C

Sea level

4.1-3

332 C1e 11.102.01 E ®

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4.1.4

Perf ormance on 1 engin e Take-off weight

kg Ib

Rate-of-climb at intermediate emergency power S.L. Service ceiling at intermediate emergency power S.L. (Vz = 0) ISA Service ceiling at intermediate emergency power S.L. (Vz = 0) ISA +20 Maximum temperature for take-off in Cat. A from clear heliport at S.L.









6,000 13,230

7,000 15,430

8,000 17,630

8,600 18,960

9.0 1,772

6.8 1,339

4.7 925

3.4 669

m ft

5,200 17,060

3,700 12,139

2,500 8,202

1,800 5,906

m ft

4,500 14,764

3,050 10,007

1,750 5,741

950 3,117

C°

> 50

> 50

> 50

40

m/sec. ft/min.

ISA

ISA + 20° C

7,500 16,534

6,900 15,211

Maximum take-off weight at max. emergency power kg

In hover IGE (10 ft)



Ib kg

In hover OGE



Ib

6,910 15,233

6,320 13,933

Operating limitations The aircraft is cleared to operate within the following altitude and temperature limitations : 

Maximum pressure altitude 

Flight :

-

M < 8,350 kg : 7,620 m - 25,000 ft M > 8,350 kg : 2,895 m - 9,500 ft



Take-off and landing :

4,572 m - 15,000 ft



Maximum temperature

ISA + 35°C, limited to 50°C



Minimum temperature

-30°C (basic) -45°C (with optional installation)

4.1-4

332 C1e 11.102.01 E ®

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4.2. Abbreviations AEO : AGL : DA : IGE : ISA : MCP : OEI : OGE : PA :

All Engines Operative Above Ground Level Density Altitude In Ground Effect International Standard Atmosphere Maximum Continuous Power One Engine Inoperative Out of Ground Effect Pressure Altitude

SL : TAS TOP : VNE : VTOL : Vtoss : Vy : Vz :

Sea Level True Air Speed Take-Off Power Never Exceed Speed Vertical Take-Off and Landing Take-off safety speed Optimum climbing speed Rate-of-climb

nautical miles knots feet/minute meters per seconds degrees Celsius

hr:min : kg : lb : km :

hours:minutes kilograms pounds kilometers

Units

nm : kts: ft/min : m/sec : °C:

332 C1e+ 11.102.01 E

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332 C1e+ 11.102.01 E

4.2-2



4.3. Performance charts The performance charts presented hereafter apply to an aircraft as per the standard definition. 

































Take-off weight in hover IGE, 10 ft, on 2 engines at take-off power (or max torque, no wind) Take-off weight in hover OGE, (on 2 engines at take-off power or max torque, no wind) Maximum cruise speed Pitch : 16.5 for weight  8,350 kg – 18,410 lb Pitch : 16° for weight > 8,350 kg – 18,410 lb ISA Maximum cruise speed Pitch : 16.5 for weight  8,350 kg – 18,410 lb Pitch : 16° for weight > 8,350 kg – 18,410 lb ISA + 20°C Recommended cruise speed (pitch 15°5) ISA

4-3.2

Recommended cruise speed (pitch 15°5) ISA + 20°C Rate of climb in oblique flight on 2 engines at best climb speed ISA

4-3.7

Rate of climb in oblique flight on 2 engines at best climb speed ISA + 20°C Rate of climb in oblique flight on 1 engine at intermediate emergency power ISA

4-3.9

4-3.3 4-3.4

4-3.5

4-3.6

4-3.8

4-3.10

Rate of climb in oblique flight on 1 engine at intermediate emergency power ISA + 20°C

4-3.11

Hourly fuel consumption at maximum cruise speed (pitch 16°5, M  8,350 kg) ISA

4-3.12

Hourly fuel consumption at maximum cruise speed (pitch 16°, M > 8,350 kg) ISA

4-3.13

Hourly fuel consumption at maximum cruise speed (pitch 16°5, M  8,350 kg) ISA + 20°C

4-3.14

Hourly fuel consumption at maximum cruise speed (pitch 16°, M > 8,350 kg) ISA + 20°C

4-3.15

Hourly fuel consumption at recommended cruise speed ISA

4-3.16

Hourly fuel consumption at recommended cruise speed ISA + 20°C

4-3.17

Take-off clear heliport Cat A

4-3.19

332 C1e 11.102.01 E

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TAKE-OFF WEIGHT IN HOVER IGE (HEIGHT = 10 FT) (10 ft, on 2 engines at t ake-off pow er or maximal t orque, no wi nd)

332 C1e 11.102.01 E

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TAKE-OFF WEIGHT IN HOVER OGE (on 2 engines at take-off power or maximal torqu e, no wi nd)

332 C1e 11.102.01 E

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MAXIMUM CRUISE SPEED Pitch : 16°5 for weigh t

8,350 kg - 18,410 lb

Pitch : 16° for weigh t > 8,350 kg - 18,410 lb ISA

332 C1e 11.102.01 E

4.3-4 ®

®



MAXIMUM CRUISE SPEED Pitch : 16°5 for weigh t

8,350 kg - 18,410 lb

Pitch : 16° for weigh t > 8,350 kg - 18,410 lb ISA + 20°C

332 C1e 11.102.01 E

4.3-5 ®

®



RECOMMENDED CRUISE SPEED Pitch : 15°5 ISA

332 C1e 11.102.01 E

4.3-6 ®

®



RECOMMENDED CRUISE SPEED Pitch : 15°5 ISA + 20°C

332 C1e 11.102.01 E

4.3-7 ®

®



RATE-OF-CLIMB IN OBLIQUE FLIGHT (on 2 engines, at best climb speed) ISA

332 C1e 11.102.01 E

4.3-8 ®

®



RATE-OF-CLIMB IN OBLIQUE FLIGHT (on 2 engines, at best climb speed) ISA + 20°C

332 C1e 11.102.01 E

4.3-9 ®

®



RATE-OF-CLIMB IN OBLIQUE FLIGHT (on 1 engine, at intermediate emergency power) ISA

332 C1e 11.102.01 E

4.3-10 ®

®



RATE-OF-CLIMB IN OBLIQUE FLIGHT (on 1 engine, at intermediate emergency power) ISA + 20°C

332 C1e 11.102.01 E

4.3-11 ®

®



HOURLY FUEL CONSUMPTION AT MA XIMUM CRUISE SPEED (pitc h 16°5) M

8350 kg ISA

332 C1e 11.102.01 E

4.3-12 ®

®



HOURLY FUEL CONSUMPTION AT MA XIMUM CRUISE SPEED (pitc h 16°) M > 8350 kg ISA

332 C1e 11.102.01 E

4.3-13 ®

®



HOURLY FUEL CONSUMPTION AT MA XIMUM CRUISE SPEED (pitc h 16°5) M

8350 kg

ISA + 20°C

332 C1e 11.102.01 E

4.3-14 ®

®



HOURLY FUEL CONSUMPTION AT MA XIMUM CRUISE SPEED (pitc h 16°) M > 8350 kg ISA + 20°C

332 C1e 11.102.01 E

4.3-15 ®

®



HOURLY FUEL CONSUMPTION AT RECOMMENDED CRUISE SPEED ISA

332 C1e 11.102.01 E

4.3-16 ®

®



HOURLY FUEL CONSUMPTION AT RECOMMENDED CRUISE SPEED ISA + 20°C

332 C1e 11.102.01 E

4.3-17 ®

®



Blank

332 C1e 11.102.01 E

4.3-18 ®

®



TAKE-OFF CLEAR HELIPORT Cat. A

332 C1e 11.102.01 E

4.3-19 ®

®


332 C1e 11.102E01

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