MANUAL DE VUELO EC 135

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EC 135Training ManualIntro

Intro 1July 2002For training and information only

EC 135 -- Training Manual

EUROCOPTER DEUTSCHLAND GmbHHelicopter Training CenterP.O. Box 1353D--86603 DonauwörthPhone: (0049) 906 71--4481Fax: (0049) 90671--4499





Foreword

Welcome to the EUROCOPTER EC 135 Training Course. This coursewas designed to instruct pilots and maintenance personnel on theEC 135 helicopter.

The training manual is comprised of 9 modules and takes intoconsideration, to a certain extent, ATA 104 specifications. It correlates

to the sequence of the factory training you will receive.Annotation to the Training Manual

This training manual is not a subject for revision service. It is themanufacturer’s practice to improve continously its products andtherefore the right is reserved to make without notice alterations indesign or manufacture which may deemed necessary.

All rights reserved.

Reproduction or translation in whole or in part of the contents of thispublication without permission of EUROCOPTER is not authorized.

1. edition December 2000

1. revision June 2002

2. revision July 2002

Modules

00 General Information

01 Lifting System

02 Fuselage

03 Tail Unit

04 Flight Control05 Landing Gear

06 Power Plant

07 Standard Equipment (not applicable for thismanual)

08 Optional Equipment (not applicable for this

manual)09 Electrical System

10 Inspections





Abbreviations

A

A Ampere

a/c, acft Aircraft

AC Alternating current AEO All Engines Operative

Ah Ampere hours

AR Auto

ARIS Anti resonance rotor isolation system

ATA Air Line Transport Association

B

B.A. Bleed air

BAT Battery

BIT Built in test

B.L. Buttock line

CCAD Caution and advisory display

CAS Calibrated airspeed

Cat. Category

CCW Counter clock wise

CDS Cockpit Display System

CG Center of gravity

CPDS Central panel display system

CSAS Control stability augmention system

CT Continuous test

CTR Center

CW Clockwise

D

DC Direct current

DCU Data control unit

DG Directional Gyro

DISCH Discharge

E

EEC Electronic engine control (P&W)

EECU Electronic engine control unit (TM)EFIS Electronic flight instrument system

e.g. For example

EGT Exchaust gas temperature

EHA Electronic hydraulic actuator





EMER Emergency

ENG Engine

EPU External power unit

EXT External; extinguisher

F

FADEC Full Authority Digital Engine Control

FCDM Flight control display module

FCDS Flight control display system

Fh Flight hours

FLIR Forward looking infra red

FLI First limit indication

FLI Flight manual

FMM Fuel metering module

FMS Flight manual supplement

FMU Fuel metering unit

FRP Fibre reinforced plastic

F.S. Fuselage station

ft Foot (feet)

G

GA Go around

GAL; gal Gallon

GEN Generator

GRP Glassfibre reinforced plastic

GS, gs Ground Speed

GSE Ground service equipment

H

h; hr Hours of time

hPa Hectopascal

HTG Heating

HTR sw Heater switch

HUMS Health and Usage Monitoring System

HV Height velocity

HY, HYD, HYDR Hydraulic

I

IAC--AR Interstate Aviation Commitee--Aviation Register

IAS Indicated airspeed

IC IntercommunicationICP Instrument control panel

ICS Intercommunication system

i.e. That is (id est)

IFR Instrument flight rules

IFCO In Flight Change Over





IGE In ground effect

IMC Instrumental meteorolocical conditions

Imp. Imperial

in. Inch

IND Indicator

INV InverterISA International Standard Atmosphere

J

JAR Joint Airworthiness Requirements

K

KCAS Knots calibrated airspeedkg Kilogram

KIAS Knots indicated airspeed

km Kilometer

kt Knot

KTAS Knots true airspeed

kW Kilowatt

L

L, l, LTR, ltr Liter

lb Pound

LBA Luftfahrt Bundesamt

LDG Landing

LDP Landing decicion point

LEP List of effective pages

LH Left hand

LOAP List of applicable publications

LRM Line replaceable moduleLRU Line replaceable unit

LVDT Linear voltage differential transducer

M

m Meter

MAN Manual mode of operation

max Maximum

MC, mc Maximum continuous

MCP Maximum continuous power

MEL Minimum equipment list

MFD Multi function display

MGT Measured gas temperatureMHS Mechano--hydraulic servo actuator

MIL Military standard, military specification

min. Minimum

MISC Miscellaneous

MM Mast moment





mm Millimeter

MMC Metal matrix compose

MMEL Master minimum equipment list

MOD Modification

MSL Mean sea level

MTBF Mean time between failureMTOW Maximum take-off weight

N

NACA

N1, n1, Ng, ng Gas generator speed

N2, n2, Np, np Power turbine speed

NAV Navigation (radio)

ND Navigation display

NMS Navigation management system

No., no. Number

NORM Normal mode of operation

NR, NRO Rotor speedNVG Night vision goggles

O

OAT Outside air temperature

OEI One engine inoperative

OGE Out of ground effect

OPT Optional equipment

OVHT Overheat

P

PA Pascal

PA Pressure altitude

PAX Passanger

pb Push button

PEC Position error correction

PFD Primary flight display

PLA Power lever angle

P/N Part number

POR Point of regulation

R

RA Radio altimeter

RAI Registro Aeronautico Italiano

R/C Rate of climb

RCU Reconfiguration control unit

R/D Rate of decent

RD Reference datum

Rev. Revision





RH Right hand

RPM, rpm Revolutions per minute

S

s, sec. Seconds of time

SAR Search and rescue

SAS Stability augmention system

SB Service bulletin

SEL Selector

SEMA Smart electro-mecanical actuator

SGL Single

SHED Shedding

SHP Shaft horse power

SL Sea level

SMD Smart multifunction display

S/N Serial number

SOV Shut-off valve

SPAS Stick position augmention systemSPIFR Single pilot IFR

sq Square

SRU Shop repalcement unit

STA. Station

STBY Stand-by

std Standard

SW, sw ’Switch

SYS System

V

VH Maximum horizontal speed

VHF Very high frequency

VMC Visual meteorolocical conditions

VMO, VMO Maximum operating speed

VNE, VNE Never exceed speed

VOR VHF omnidirectional radio ranging

VRM Video and radar module

VTOSS Take-of safety speed

V Y Best rate-of -climb speed

W

W.L. Waterline

WXR Weather radar

X

XMSN Transmission

XPDR/XTR Transponder



EC 135Training ManualGeneral

00 -- 1July 2002For training and information only

General Description





Handling of the EC 135 124. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Lifting 124. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Jacking of the EC 135 126. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Shoring 126. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Weighing 128. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Leveling 130. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Towing and Pushing 132. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Parking and Mooring 134. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .





The Development of the EC 135

History

The first EUROCOPTER (ex. MBB, ex. BÖLKOW) helicopter withglass fiber rotor blades was the single blade helicopter BO 102, acaptive trainer, operating for the first time in 1957. In 1961 the single

seater BO 103 followed, the only helicopter to fly with one rotor blade.In 1962/63, a new hingeless rotor system was created, andsuccessfully tried on an Alouette II, in Marignane, France.

From 1960 to 64 the high speed helicopter BO 46 was designed withthe Derschmitt rotor system.

In 1964 these helicopters were followed by the multi purpose 2 1/2 tontwin engine helicopter BO 105.

TosubstitutetheBO105after20yearsinduty,theBO108wascreatedand flown on Okt. 15th, 1988 for the first time. Consultations withpotential customers -- operators of EUROCOPTER products and ofcompeting types -- showed that cabin volume should be increased andvisibility improved and that greater emphasis would have to be put onmission flexibility (the cabin floor, for instance should be flat andunobstructed to allow easy conversion from passengers to cargoroles). In late 1992, the design was modified to provideaccommodation for max. six passengers, instead of the BO 108’sthree, and two crew. The Aerospatiale developed Fenestron AntiTorque system was adapted, and the EC 135 as it is today took shape.

In the middle of 1996, the certification by the German (LBA) and the American Airworthiness Authorities (FAA) was completed.

Engine Versions

The following engine versions are possible:

-- EC 135 P1equipped with Pratt & Whitney PW 206 B engines.

-- EC 135 P2equipped with Pratt & Whitney PW 206 B2 engines.

-- EC135 T1equipped with Turbomeca ARRIUS 2B1, 2B1A, 2B1A_1

-- EC135 T2equipped with Turbomeca ARRIUS 2B2 engines.

Both engine types are in the 450 KW class. The maximum take-off

weight for both standard versions is 2720 kg (upgrade to 2835 kgMTOW is possible), with external load 2900 kg.

Cockpit Versions

Two major cockpit versions are possible:

-- CPDS (Central Panel Display System with multifunctionscreens) together with analog flight instruments. As an

option, the CPDS can be combined with FCDS (FlightControl Display System).-- CDS (Cockpit Display System) with analog flight

instruments or EFIS (Electronic Flight InstrumentationSystem)

NOTE CDS Standard cockpit has been replaced by CPDS

cockpit.





EC 135 Variants

EC135 P1Pratt&Whitney Engine

206 B

EC135 T1TURBOMECA Engine

ARRIUS 2B1, 2B1A, 2B1A_1

CDSCPDS CPDS+FCDS CDS CPDS CPDS+FCDS

EC135 T2TURBOMECA Engine

ARRIUS 2B2

CPDS CPDS+FCDS

EC135 P2Pratt&Whitney Engine

206 B2

CPDS+FCDS CPDS

AnalogInstruments

EFIS AnalogInstruments

EFIS





General Description of the EC 135

General

The EC 135 is a light multi purpose twin engine helicopter in the 2.5tclass. There are five seats in the basic version, they can be extendedup to eight seats.

Engines

The EC 135 T is powered by two engines Turbomeca ARRIUS 2B, theEC 135 P is powered by two engines Pratt & Whittney PW206 B. Theyare equipped with a digital engine control system.

Transmission

The main transmission is a two-stage flat gearbox (produced by

Zahnradfabrik Friedrichshafen ZF), which is mounted by ananti-resonance rotor isolation system (ARIS) on the transmissiondeck.

Main Rotor

The helicopter is equipped with a four-bladed hingeless andbearingless main rotor (BMR). The inboard flexbeam enablesmovement of the blades in all axes. Blade pitch angles are controlled

through integrated glass/carbon fibre control cuffs.The main rotor control linkage system is of conventional design. Thehydraulic system for the main rotor controls is designed as a duplexsystem with tandem piston (both systems are active). In case of afailure of one system, the remaining system has sufficient power toensure safe flight operation and a safe landing.

Tail Rotor System

The helicopter is equipped with a “Fenestron” tailrotor system. Thereare 10 blades rotating in a housing integrated in the tail boom.

The Fenestron is controlled via a “Flexball” type cable, routed fromthe

pedals to the input control rod of the Fenestron.Tail Boom

The tail boom can be separated from the fuselage, and consists of tailboom cone, the horizontal tail plane with end-plates, vertical fin withintegrated tail rotor, tail rotor gearbox and fairing.





Dimensions

10.20m

12.16m

10.20m

5.87m

3.51m

2.65m

3.20m

3.35m

2.00m 1.56m

5°

0.66m3°





Maintenance Concept

General

“Maintenance” covers all scheduled and unscheduled maintenanceactivities. It also applies to the on condition maintenance. It is basedon condition monitoring by visual checks/inspections and diagnosticfeatures such as chip detectors, filter bypass indicators, boroscopeaccess, failure code indications, built-in tests, warning lights etc.

Maintenance Levels

EC 135 maintenance is split into three maintenance levels:

-- Organizational Level (O)-- Intermediate Level (I)

-- Depot Level (D)Organizational Level

The organizational level covers tasks of the daily servicing,maintenance checks, inspections for condition, exchange ofcomponents (LRU’s) and quick, simple repairs as specified in theaircraft maintenance manual (AMM).

The work generally takes place at the operators site.After a “on the job

training” these checks can be carried out by pilots, mechanics andoperators.

Intermediate Level

The intermediate level covers repairs on/off helicopter, extendedperiodical inspections as specified in the aircraft maintenance manual.To fulfill these tasks, maintenance facility, qualified personel, testequipment and special tools are required.

NOTE The maintenance manual covers all tasks of

organizational level and intermediate level.

Depot Level (D)

Depot level covers major repair or overhaul at the manufacturer or atauthorized service stations under industrial premises.

More extensive tools/test equipment and specialized personnel arenecessary.

NOTE Documentation and spares for depot level tasks

will be delivered to authorized customers only.

NOTE Information about inspections and intervals are to

be found in chapter 10 of this training manual.





Maintenance Concept

Maintenance

ScheduledUnscheduledOn Condition

OrganizationalLevel (O)

IntermediateLevel (I)

Depot Level (D)

Daily servicing, maintenance checks

inspections for condition, exchange of LRU‘s. acc. to AMM -- Can be carried

out by a mechanic or by the pilot (i.e. main transmission servicing).

Repair on/off the helicopter extended

periodical inspections acc. to AMM -- maintenance facility, qualified personnel,

test equipment and special tools are required (i.e. main transmission change).

Major repair or overhaul at the manufacturer or at authorized service stations

acc. to special documentation. Tools/test equipment and specialized personnel

are neccessary (i.e. main transmission overhaul).

Manufacturer/authorized customers only





Documentation of the EC 135

General

The documentation of the EC 135 consists of two main groups:

-- EC 135 helicopter documentation written byEUROCOPTER

-- Other manufacturers documentationsLayout

The whole documentation library is prepared in general compliancewith Air Transport Association Specification 100 and ATA 2100. Thecustomized documentation is available for certain H/C serial numbersor a group of H/C serial numbers. A part of the documentation libraryis delivered on CD ROM.

Revision Reissue

Changes in the helicopter equipment, maintenance practices,procedures etc. update and replace the manual content. To ensurethat the EC 135 manuals continue to show the latest information, twicea year the CD ROM is replaced by a reissue. The preceding issue thenbecomes obsolete and must be discarded.

ATA Numbering

The numbering system provides a procedure for dividing material intochapter section subject and page. The number is composed of threeelements, which have two numbers each. The chapter and sectionelement are established by ATA 2100. Subject and unit elementnumbers are assigned by ECD.

Page Number Blocks

Page number blocks are used for the different sections of themaintenance manual to logically place the activities in sequence asfollows: Procedures have either a brief subtopic or a combination ofsubtopics i.e. Removal/Installation, Inspection/Test. If subtopics arebrief, then they are combined in one topic under MaintenancePractices. If the subtopics become lengthy so that a combinationwould require numerous pages, the topics are broken out into pagenumber blocks.

-- Pageblock 1--99 System Description-- Pageblock 101--199 Troubleshooting-- Pageblock 201--299 Maintenance Procedures-- Pageblock 301--399 Servicing-- Pageblock 401--499 Removal/Installation-- Pageblock 501--599 Adjustment/Test-- Pageblock 601--699 Inspection-- Pageblock 701--799 Cleaning/Painting-- Pageblock 801--899 Repair

-- Pageblock 901--999 Storage

NOTE Element 1, element 2 and the pageblocks are set

by the ATA 2100 schematic. The following

elements can be defined by the aircraft

manufacturer as required.





ATA Numbering

28 -- 10 -- 00Chapter Section Subject/Unit

1. Element 2. Element 3. Element





Mechanic’s Documentation

The mechanic has available (CD or hardcopy):

-- Systems Description Section (SDS)-- Aircraft Maintenance Manual (AMM)-- Master Servicing Manual (MSM)-- Wiring Diagram Manual (WDM)

-- Illustrated Parts Catalog inclusive Tools Catalog (IPC)

Pilot’s Documentation

The pilot has four documents available (hardcopy):

-- The Flight Manual (FLM) according Helicopter AssociationInternational, HAI

-- Log Book

-- Pilot’s Checklist (PCL)-- Master Minimum Equipment List (MMEL)

NOTE The Flight Manual, the Pilot’s Checklist and the

Log Book are hardcopies and must always be

present in the helicopter.

Operator’s Technical Control Documentation

The following documents are kept by the operator’s technical control:

-- Historical Record-- LOAP (List of applicable publications, hardcopy)-- Service Bulletins / Alert Bulletins, (hardcopy)-- Service Informations / Alert Service Informations,

(hardcopy)Other Manufacturer’s Documentation

The other manufacturers (engines, Avionics and optional equipment)deliver their own documentation:

-- Engine Maintenance Manual-- Engine Illustrated Parts Catalog

-- Engine Service Bulletins / Service Letters-- Avionics Manuals-- Special optional equipment (e.g. external hoist system)

NOTE The valid manuals incl. the revision status are

published in the LOAP (list of applicable

publications).





FLMT1, CDS/CPDST2 CPDSP1, CDS/CPDSP2 CPDS

Log Book

MMEL

SDS

WDM

IPC + Tools

AMM

HistoricalRecord

SB/ASB

SI/ASI

LOAP

ECD Helicopter Documentation EC 135

Mechanic PilotOperator

CD--ROMPCLT1, CDS/CPDS

T2 CPDSP1, CDS/CPDSP2 CPDS

MSM





Reference Planes

General

The frame coordinates of the EC 135 are defined in accorance withLN 65619 (Luftfahrtnorm). All dimensions are given in the metricsystem (mm).

The reference planes are used to determine locations on and withinthe helicopter.

Definitions

Locations on and within the helicopter can be determined in relationto fuselage stations, buttock lines and waterlines, measured inmillimeters (mm) from known reference points. Fuselage stations,buttock lines, and waterlines are planes perpendicular to each other.

Reference plane is the plane at the longitudinal centerline of thehelicopter perpendicular to the cabin floor.

Fuselage Stations

Fuselage stations (FS) are vertical planes perpendicular to, andmeasured along, the longitudinal axis of the helicopter.

Station 0 is an imaginary vertical plane forward of the nose of the

helicopter, from which all horizontal distances are measured forbalance purposes (see also “reference datum”).

Buttock Lines (+/ -- Y Coordinates, Lateral)

Buttock lines (BL) are vertical planes perpendicular to, and measuredto the left and right along the lateral axis of the helicopter.

Buttock line (0) is the plane at the longitudinal centerline of the

helicopter.Waterline (+ Z Coordinates, Vertical)

Waterlines (WL) are horizontalplanes perpendicular to, and measuredalong, the vertical axis of the helicopter.

Waterplane (0) is a plane 1505 mm below the cabin floor at fuselagestation 2160 mm.

Reference Datum (+ X Coordinates Longitudinal)

The reference datum (RD) is an imaginary vertical plane foreward ofthe helicopter nose. The station is located is located 4000 mm in frontof the leveling point (center of double frame #4).

NOTE The standard helicopter is well clear to the

reference planes in order to avoid negative

coordinates (X; Z) after exterior optional equipmentis mounted.





Reference Planes

X 1099.32

Z 1505X 4000

X 2160 Y+

Y--

Z

X





Cockpit Arrangement

General

The EC 135 is provided with several units for monitoring, warning andcontrol purposes. These units are installed to certain control panels.

Control Panels

The control panels installed in the EC 135 are subdivided into:

-- Overhead Panel-- Instrument Panel-- Slanted Console-- Center Console





NDPFDNAVSOURCEEXT

DH

CRS

TST

BARO

POS

STDNDPFDNAVSOURCEEXT

DH

CRS

TST

BARO

POS

STD

Overhead Panel

Cockpit Arrangement (CPDS; FCDS)

Instrument Panel

Slant Console

Center Console

Center Post

Cyclic Stick

Collective Pitch





Instrument Panel with CPDS

General

The instrument panel contains most of the displays and instrumentsand some of the control units installed in the helicopter. Theconfiguration of the instrument panel varies according to operatorsneeds and the associated equipment.

System Components

The instrument panel consists of:

-- Center Console-- RH section-- LH section

Center Console

The center console of the instrument panel contains the CDS (CockpitDisplay System) in earlier versions or the CPDS (Central Panel DisplaySystem) with analog back up instruments and the warning unit todisplay system/engine conditions. A chronograph is also included. Anumber of switches for engine and electrical system operation arelocated on the center console, too.

RH Section

The RH section of the instrument panel contains theinstruments/displays for flight control and navigation. A number ofswitches may be provided for controlling the radio/navigation system.

A nozzle is provided for regulating fresh air supply.

LH Section

The LH section of the instrument panel is specified for the copilot. Theconfiguration of the LH section varies according to helicopterequipment.





Instrument Panel (CPDS, Analog Flight Instruments)

c

Switch Unit

Antiglare Device Warning Unit

VEMD

LH SECTION CENTER SECTION RH SECTION

CAD

Nozzle

Altitude Indicator

Vertical SpeedIndicator

Airspeed Indicator Analog Clock

D--HUMS

Triple Speed IndicationN2 /NRO

Artificial Horizon

HSI





INTENTIONALLY LEFT BLANK





ND

PFDNAVSOURCE

EXT

DH

CRS

TST

BARO

POS

STD

Instrument Panel (CPDS, FCDS)

Switch Unit

Antiglare Device Warning Unit


CPDS

Navigation Display

Nozzle

Analog Instruments

Primary Flight Display

ND

PFDNAVSOURCE

EXT

DH

CRS

TST

BARO

POS

STD





Instrument Panel with CDS

General

All the instruments and indications to monitor the helicopter systemsare installed in the center section of the instrument panel.

Configuration

-- The following instruments, indicators and switches areinstalled in the center section of the instrument panel:

-- Warning unit-- Triple rotor RPM indicator (incl. N2 indication for eng. 1/2)-- Torque indicator-- Dual TOT indicator-- Dual nN1 indicator (T1 engine only)

-- Dual N1 indicator (P1 engine only)-- Chronograph-- Switch unit-- Oil temperature and pressure indicator for engines and

main transmission (Different limit markers with the differentengines)

-- Cockpit Display System (CDS)





Instrument Panel (CDS, P1, EFIS Cockpit)

SCROLLDISPLAY

SELECT

SYSTEM II

MM

XFERLOWFREELOW

BAT AMPSGEN AMPSDC VOLTS

LMT

SYSTEM I

PAGE

CAUTIONPAGE

VOLT AMP

FUEL

HOOKLOAD

VneGROSSMASS

CDSFAIL

BRIGHTNESSWEIGHT(Vne)

2

OPT1

PF

M

T

CABLELENGTH

RAD ALT

LBKGLBKG AUXSPLY2

LBKG LBKGMAINSPLY1

OAT

TQ

S YS TEM I MI SC SY STE M II

MASTER

CAUTION

c

Switch Unit

Antiglare DeviceWarning Unit

Analog Instruments:System/Engine


CDS

Switches

Nozzle

Altitude Indicator

Vertical Speed

Indicator

Airspeed Indicator

EFIS

PITCH

DAMPER

Oil Temperature/Pressure Indicators





Engine Monitoring Instruments TM

Triple Rotor RPM Indicator

TOT Indicator

Torque Indicator

N1 Indicator





Dual nN1 Indicator T1

General

The Dual n N1 indicator is part of the speed sensing system. It is a2--pointer instrument and indicates the RPM of the following:

-- n gas producer RPM between the max. allowed (computedby the FADEC) RPM and the present RPM for engine 1

and engine 2.It is a 2--pointer instrument. The pointers are labelled “1” and “2”. Theindication range is from -- 8 % to + 4 %.

Dual N1 Indicator P1

General

The Dual N1 indicator is part of the speed sensing system. It is a2--pointer instrument and indicates the RPM of the following:

-- Gas producer RPM for engine 1 and engine 2

It is a 2--pointer instrument. The pointers are labelled “1” and “2”. Theindication range is from 0 % to + 120 %.

NOTE The limit values might be different according to theengine version installed.





Engine Monitoring Instruments P1

N1 Indicator

Torque Indicator

TOT Indicator

Triple Rotor RPM Indicator





Oil Temperature and Pressure Indicator

General

The oil temperature and pressure indicator is an instrument clusterindicating oil temperature and oil pressure for each engine and for themain transmission on six individual indicators.

-- The temperature indicators are calibrated in °C

-- The pressure indicators are calibrated in bar

According to the engine type installed (TM or PW) the indicators havedifferent scaling and different limit markers.

The indicator illumination is adjusted with the aid of instrumentillumination potentiometer INSTR in the overhead panel.

More detailed description is given in the associated chapters.





Oil Temperature-- and Pressure Indicator

Turbomeca

Pratt&Whitney





Cockpit Display System (CDS)

Introduction

The Cockpit Display System (CDS) provides indication of aircraftstatus information such as caution and advisory messages to the crewaccomplished by operating data and indication of special operationmodes. It consistsof a self contained unit installed in the center section

of the instrument panel.Various switches facilitate operation of the device and allow control ofthe indications. The brightness is automatically controlled with the aidof a sensor.

An indication light flashes as soon as the CDS discovers an internalmalfunction.

The CDS is capable of identifying the type of engine installedaccording to the wiring of the connectors.

The casing of the CDS is cooled by the cabin ventilation system or theair-conditioning system, if installed.

Associated Controls and Indicators

In order to provide proper function and handling, the following controlsand indicators beside the CDS are available:

-- MASTER CAUTION indication light

The MASTER CAUTION indication light is installed in the center partof the instrument panel RH of the warning unit.

-- Switch CDS/AUDIO RESET

The switch CDS/AUDIO RES is installed in the grip of the cyclic controlstick and enables the pilot and copilot (if dual pilot controls areinstalled) to acknowledge the CAUTION indications.

-- Test Switch TEST/CDS

The test switch TEST/CDS is installed inthe overhead panel. It triggersthe testing of the CDS indications.

-- CDS OVTP indication light

The CDS OVTP indication light is installed in the center part of theinstrument panel below the CDS on the left side. The light comes onif the internal temperature is higher than 63 °C.

Power Supply

In order to guarantee continuous operation even in the event of failureof one of the essential busbars, the CDS is supplied by bothESSENTIAL busbars via the circuit breakers located in the overheadpanel.

-- CDS/SYS 1-- CDS/SYS 2

Data Storage An CDS integrated memory has two functions which are as follows:

-- Storage of all of the CAUTION indications having occurredwithin the penultimate minute

-- Storage of the failures reported to the CDS by the enginecontrol units along with their respective failure codes.

EC





CDS -- General Arrangement

Master

Caution

ESSENTIAL I

CB CDS/SYS1 CB CDS/SYS2

Aircraft DataEngine Data

ESSENTIAL II

CDS AUDIO

RES

CDS

OVTP

EC 135





ConfigurationThe CDS provides the crew with information while at the same timeindicating the present state of various systems of the helicopter.The CDS performs the following tasks:

-- CAUTION indication-- Advisory indication-- Indication of engine parameter (engine cycle counter),

FADEC--MEM--codes and malfunction indications-- Indication of helicopter’s power supply voltage and current-- Outside air temperature indication-- Mast moment bargraph with limit warning light*-- Fuel system indication-- Calculation and indication of Vne velocity **

-- Radar altimeter indication-- Indication of length of rescue winch cable*-- Indication of load attached to external cargo hook*-- Engine operating hours counter

* Only available when the resp. systems are installed in the helicopter.

** The key Vne is installed in early CDS versions only. Current versionsare provided with a key FUNCTION.

The CDS is divided into several panels to enhance overall view. Eachof these panels serve assigned functions.

The basic brightness of the indications is controlled through the keysBRIGHTNESS.

Colors of Indications-- Amber

The upper display which is theprimary display is split into four sections.In the upper part CAUTIONS are displayed separately for SYST I/IIand MISC. The color of the cautions is amber.

-- Green

The lower part of the upper display shows the ADVISORIES The colorof the advisories is green.

-- White

The color in the lower display which is the secondary display in generalis white.

Exceptions are made with the mast moment indication which is green-- yellow -- red and fuel low indications in the fuel display which are red.

EC 135





CDS -- Displays and Controls

Caution DisplaySYST I, MISC, SYST II

Advisory Display

Page Light

Mast Moment Indication

VOLT/AMP Key

Scroll Buttons

Display Select Switch

CDS FAIL Indication

BRIGHTNESS KeysWEIGHT Key

Opt 1/2 Display

FUEL SYSTEM Display

OAT Indication

Electrical System Display

TORQUE Display

Engine Parameter DisplayDefault Values:N1 for TM, TOT for P&W

CAUTION PAGE Button

Brightness Sensor

EC 135





CDS OperationPower Supply and Self Test

The CDS is activated by setting the battery master switch BAT MSTRin ON position. This causes the CDS self test to be carried out. TheCDS checks also the presence of the following engine cautions forSYS I and SYS II:

ENG FAIL ENG FAILENG OIL P ENG OIL PFUEL PRESS FUEL PRESS

HYD PRESS HYD PRESSXMSN OIL P XMSN OIL P

GEN DISCON GEN DISCON

If the cautions have been successfully detected INP PASSED comeson on the advisory display below the message CDS PASSED andengine configuration (early CDS versions only). If a caution is missing,INP FAIL appears in the center column of the caution display, followedby the missing caution to the left/right.

The pilot has to acknowledge the messages by pushing theCDS/AUDIO RES button on the stick grip. Subsequent to theacknowledgement the CDS starts normal operation. If the self test wasnot successful CDS FAIL will appear on the display.

The indication light CDS FAIL comes on only when the CDS self testis faulty.

Mast Moment Failure

If there is a failure of the mast moment system detected, the cautionMMFAILED comes up in the MISC field (depends on the part number).

Continuity Test

Continuity tests of the connecting cables between some sensors and

the CDS are made during CDS power -- ON self test. A failure isindicated by displaying the respective detector name with anadditional ...CT on the caution panel. If a ...CT -- caution is indicated,the monitoring circuit of the corresponding system must be assumedto be unable to activate the real system caution in case of systemfailure.

CDS Test Switch

The CDS test switch, located on the test switch panel of the overheadconsole provides test function of the display screens and lamps of theCDS. Activation of the test switch causes the screens and lamps of theCDS and the indication CDS OVTP to illuminate.

EC 135





CDS Self Test

ENG FAILENG OIL PFUEL PRESSHYD PRESS

XMSN OIL PGEN DISCON

ENG FAILENG OIL PFUEL PRESSHYD PRESS

XMSN OIL PGEN DISCON

CDS PASSED(Engine config.)*

INP PASSED

SYSTEM I MISC SYSTEM II

CDS PASSED(Engine config.)


INP FAIL

CDS FAIL


HYD PRESS

All Parameters Available, Self Test Passed Signal HYD PRESS Missing

Self Test Not Passed.

* Early CDS versions only

EC 135





CDS Caution DisplayGeneral

The cautions are displayed in the CAUTION display, separately forsystem 1, system 2 and miscellaneous.

New cautions emerging on the screen are accompanied by flashinglines above and below the caution. Cautions, displayed before, areextinguished from the display but stored in the background.

Each new caution indication causes the MASTER CAUTION light tocome on (The master caution light is located right beside the warningpanel).

The cautions must be acknowledged by pressing the CDS/AUDIORES button which is located on the cyclic stick.

After pressing the CDS/AUDIO RES button the master caution light

goes off and the CDS changes to the prioritized display mode. Thatmeans, that all active cautions are displayed in sequence of priority.

If there are more acknowledged cautions than can be displayed on thescreen simultaneously, the PAGE light illuminates and the additionalcautions can be called up from the second page by pressing theCAUTION PAGE button. If the CAUTION PAGE button has not beenpressed for 10 seconds, the top priority cautions are displayed.

NOTE The following two listings show all possible

cautions/advisories at the time this manual hasbeen printed.

The caution configuration in the individual

helicopter depends on the helicopter serial

number, CDS configuration and optional

equipment installed.The cautions will be explained in the respective

chapters.



EC 135



Training ManualGeneral


CDS Advisory DisplayGeneral

The section below the caution display contains the advisory displaywhich keeps the pilot informed about operating conditions of additionalequipment which is not essential for the flight.

Matrix of Advisory Combination

Basic Opt.Equipm.DPIFR 1

Opt.Equipm.DPIFR 2

2.Priority*

BLEED AIRBleed air heating active

X X X

LDG LIGHTStandard and/oroptional landing light on

X X X

P/S--HTR--PHeating of the pitot pilotside is active

X X

P/S--HTR--CPHeating of the pitotcopiltot side is active

X

LDG L RETR

Search and landinglight retracts at rest

X

LDG L EXTDSearch and landinglight extended

X

HOOK UNLDLoad is < 5 kg

X X

AIR COND Air condition systemactive

X X

AUX XFER Aux. fuel valve is in-open position

X X

CA CUT ARMCable cut circuit test ispassed

X

IRInfra red light is active

X X

IFCOThe IR filter is active

X X X

* 2. priority means: If all advisories are ON, the advisories of the

2. priority will not be displayed.

Display Select Switch / Scroll Button

General

The display select switch has six selectable positions which provideinformation and date about several engine parameters, failure codes,

operation parameters etc.The informations can be displayed by selecting a certain switchposition and pressing the scroll buttons to scroll in the menu.

Selectable Parameters

The following table describes the possible parameters in dependencyon the chosen display select switch position.

EC 135T i i M l





Display Select and Scroll Switch

SCROLL

DISPLAY

SELECT

2

OPT1

PF

M

T

Display Select Switch

Scroll Buttons

1

EC 135Training Manual





Posi--tion

Parameter Description

P PARAMS(Normal flight position)

RealtimeFADECparameterscanbesequentiallyselectedbymeansofthescrollbuttons.Theyaredis-playedontheengineparameterdisplay.ThedisplaydefaultuponpowerisN1(TM)andTOT(PW)Thepossibleparametersarelistedbelow.

N1 Gas generator turbine RPM [%]

N2 Power turbine RPM [%]

TQT Torque trim values of both engines [%] (TM only)

QMAT Torque trim values of both engines [%] (P&W only)

EGT Exhaust gas temperature [ûC] (TM only)

TOT Turbine outlet Temperature [ûC] (P&W only)

T1 Air temperature measured at the compressor air inlet and provided to the engine control unit. [ûC]

CLP Collective pitch resp. Linear--Voltage--Differential--Transducer--Position (LVDT) [%]

P0 Air pressure measured in both FADEC boxes [hPa]N2T Power turbine reference speed trim value [%]

N1C N1 Cycle counter

N2C N2 Cycle counter

MEM CODES Numerical Failure Codes

F FAIL MSG

The Fail Message provides abbreviated messages for active failure codes. They are displayed on the

advisory display. When viewing the FADEC failure messages and no fail code exists, a blank is displayedcontinuously. The indication scrolls automatically for 3 seconds each when more than one exists. All ofthe malfunction codes are stored. They are deleted with the next engine start when the N1 RPM exceeds20 %.






Posi--tion

Parameter Description

M MEM CODE

Stored failure codes can be selected by means of the scroll buttons and are displayed on the engineparameter display by means of numerical failure codes. These codes correspond to the abbreviatedmessages under FAIL MSG and are described in the respective maintenance manual. Mast momentexceedance MMEXC is displayed in the advisory display.

T OTh

The Operating Time counter provides automatic timer function to continuously keep and indicate the

engine operating time. The time is displayed on the engine parameter display.The counter starts when the resp. ENG FAIL CAUTION disappears and the collective lever position isabove 10%.It stops when the collective lever position is below 10% and ENG FAIL CAUTION is active.

OPT1 Enables the operator to select between VNE and RAD ALT indication on the upper option line bymeans of the scroll buttons.

VNE* The VNE depends on gross mass, pressure altitude and OAT. The present VNE is calculated and per-manently updated by the CDS. By pushing the WEIGHT button the pilot can choose between thesymbols “>” or “<”.“>” means that the gross mass is equal or greater than 2300 kg (standard presetting).“<” means that the gross mass is lower than 2300 kg.

RAD ALT * A four digit display indicates the radar altitude from 0 to 2500 ft and a RAD ALT light will come on ifthe radar altimeter is active.

OPT2 Enables the operator to select between HOOK LOAD and CABLE LENGTH indication on the loweroption line by means of the scroll buttons.

HOOK LOAD *

A four digit display indicates the loading of the external cargo hook. In addition the illuminated HOOK LOAD sign comes on.

CABLELENGTH *

The display indicates the length of the lowered cable. If the rescue hoist is active the illuminatedCABLE LENGTH sign comes on and the moving mode of the hoist (lowering or retracting) isindicated.

*Only when the respective system is installed.






Torque IndicationThe torque is permanently displayed on the torque display in %.

Electrical Power Indication

The aircraft‘s electrical voltages and currents can be shown on theelectrical system display. The VOLT/AMP button enables the crew to

select between DC VOLTS, GEN AMPS or BAT AMPS. The defaultsetting is DC VOLTS.

Outside Air Indication

Outside air temperature is permanently displayed in ûC. The value isalso internally used for VNE calculation.


The EC 135 is equipped with a hingeless and bearingless rotorsystemand therefore high bending moments occur at the rotormast,particularly during close ground operation. The bending of therotorshaft is monitored by a mast moment measuring system.

The mast moment indication consists of a bargraph and a limit light.

The bargraph is a three-color indication, indicating the mast momentlinear from 0 to 100% in green, yellow and finally red.

The LIMIT light remains on until a “cold start” of the CDS occurs.

If the input signal from the mast moment measuring system is out ofspecified values, the caution MM FAIL will be displayed.

Fuel Quantity IndicationTheCDSdisplaysthefuelmassesandfuelsystemstatusofthesupplytank 1, supply tank 2, main tank and (if installed) auxiliary tank. Eachof the tank displays contain a bargraph display and a numeric textdisplay.

The supply 1 and supply 2 displays contain a LOW indication whichilluminates when the resp. tank‘s content is below a specified value.

The FREE advisory indication comes on when the free volume of themain tank is greater than the current volume of the auxiliary tank.

The XFER advisory indication comes on when fuel is being transferredinto the main tank (transfer valve OPEN).






Outside Air Temperature, Mast Moment and Fuel System Indication

Outside AirTemperature Display

Mast Moment Displayincl.Limit Light

Fuel Display incl.Low Level Warningincl.FREE and XFER

Advisory




gGeneral


Central Panel Display System (CPDS)General

The Central Panel Display System is an electronic indicating systemand presents various parameters of the onboard systems on threescreens.

CAD (Caution and Advisory Display)

The CAD displays cautions, advisory messages and fuel systemindications. If the VEMD fails, the CAD can take over and displayselected parameters from it.

The display instrument of the CAD consists of a color screen,integrated in the left-hand side of the center section of the instrumentpanel.

VEMD (Vehicle and Engine Monitoring Display)

The VEMD displays engine and dynamic system parameters. Inaddition, it can present data relating to onboard systems (e.g. aircraftelectrical system, autopilot) and to optional equipment (e.g. cargohook).

If the CAD fails, the VEMD displays selected CAUTIONs. The duplexconfiguration of the VEMD provides redundancy so that twoprocessing modules are each individually capable of taking over all

tasks.Both the VEMD screens are installed in the in the right-hand side of thecenter section of the instrument panel.

Test Switch

The test switch triggers the CPDS to display the test page with thecomplete color spectrum and the software version.

Circuit Breakers

The CAD and the VEMD are supplied with voltage, each via two circuitbreakers, from the ESSENTIAL busbars 1 and 2. The circuit breakersare arranged in the overhead panel.

CDS/AUDIO RES Switch

The CDS/AUDIO RES switch is used by the pilot and copilot toacknowledge displayed cautions. It has the same function as theSELECT key on the CAD.

The switch is installed in the grip of the pilot’s cyclic stick and, if dualcontrols are installed, one is also installed in the copilot’s cyclic stickgrip.

Voltage Adjusting Element

An adjusting element for each voltage indication of the VEMD isintegrated in the sensor units, mounted to frame 1 in the forward partof the helicopter. Hereby the voltage drop in the VEMD indication canbe corrected.

Maintenance Connector

Two maintenance connectors are mounted to the rear part of the

slanted console.CPDS OVHT Caution

The CPDS overheat caution is triggered by a temperature switch in theinstrument panel between 51 and 55 °C.




General


CPDS -- Locations

ENGI

ENGII

OFF

M AX

CDS AUDIO

RES

CAD

MaintenanceConnectors

SwitchCDS AUDIO RES

Stick Grip

Circuit Breakers

Test Switch

VEMD

Voltage Adjusting Elementintegrated into Sensor Unitsat Frame 1

EC 135Training ManualG l



General


Color Code Ranges and their MeaningThe range of colors used for displays on the screens of the CPDScovers five different colors in addition to black and white. Eachindividual color has a specific significance.

Black Background, Text on colored backgroundWhite Scales, Display arrows (pointers), numbers, etc.

Yellow Limits, Defect symbols

Red Limits, Defect symbols Amber CautionsGreen Advisories

Cyan Tech. Units, Selections, Demarcations etc.Blue Fuel quantity level

CAD Operation

The CAD is operated by the following keys in the front panel:

Key FunctionOFF Switches CAD on/offSCROLL Selects different screen pages (e.g. second page with

cautions)

SELECT Acknowledges new cautionsBRT + Increases brightness of screenBRT -- Decreases brightness of screen

VEMD OperationThe VEMD is operated by the following keys located on the front panelof the display monitor:

Key FunctionOFF 1 Switches upper screen and processing module 1 on/

offOFF 2 Switches lower screen and processing module 2 on/off

SCROLL Cycles to next page, depending on operating modeand status

RESET Initiates return to normal screen display or topreviously displayed page (depending on theoperating mode)

SELECT Selects a particular data field+ / -- Input of values to data field

ENTER Acknowledges selection of a data field or a data entryto a data fieldBRT + Increases brightness of screen by continuous

adjustmentBRT -- Decreases brightness of screen by continuous

adjustment




General


CPDS

0

2

4

6

810

12

14

16

CAD VEMD




General


Function of the CPDSOverall System

The CAD and VEMD are each powered by two independent powersupplies and their respective circuits are each protected by two circuitbreakers.AsbothareconnectedtoESSENTIALbusbars1and2,theiroperational integrity is ensured if one of the busbars should fail.

Status of the CPDS

With the CPDS GROUND and FLIGHT status are distinguished on thebases of the following parameters:

GROUND Status:

-- n1 RPM engine 1 and engine 2 < 50 %-- XMSN oil pressure < 1bar

FLIGHT Status:

-- n1 RPM engine 1 or engine 2 > 50 %-- XMSN oil pressure > 1bar-- Angle of collective lever CLP

> 28.5% (Turbomeca) or > 17% (Pratt&Whitney).

Switch-on Sequence (Power up)

The CPDS is activated as soon as the aircraft electrical system isenergized on the ground. An internal self-test and an external self-testare run to establish the functional integrity of the CPDS:

While the internal self-test is running, the message TEST INPROGRESS will be displayed on the CAD/VEMD and the soft-- andhardware is checked.

After the internal self test is passed, the external self test is performed.While the presence of the followingparameters isverified the messageEXTERNAL SELF TEST IN PROGRESS will be displayed on theCAD/VEMD.

SYS I MISC SYS II

ENG CHIP XMSN CHIP ENG CHIP

ENG FUEL FILT TRGB CHIP ENG FUEL FILTFUEL OIL FILT XMSN OIL TEMP FUEL OIL FILT

XMSN OIL FILT

During the external test, the wiring of certain sensors is checked witha continuity test (CT). If a failure occurs, the respective sensor isdisplayed on the CAD as a caution with CT as a supplement.




General


Functional Schematic CPDS

VEMDSYS II

CADSYS II

VEMD

P1P1ESS BUS I

VEMDSYS I

CADSYS I

ESS BUS II

AUDIOGONG

Master

Caution

TEMP

Air Cond.

ARINC 429

WARNINGUNIT

TEST CDS/ WARN UNIT

CDS/ AUDIO RES

CPDS OVHT

MAINT.CONN

ARINC 429

FADEC

2 1FCDMSensors

APM

HUMS

PelicanRack

CAD

VOLTAGE ADJUSTMENT

VOLTAGE ADJUSTMENT




General


After the external self-test the functional integrity of the peripheralassemblies is tested. After the test has run, the following cautions willbe displayed on the CAD:

SYS I MISC SYS II

ENG FAIL+ F PUMP AFT** ENG FAIL+

ENG OIL P+ F PUMP FWD** ENG OIL P+FADEC FAIL* EPU DOOR FADEC FAIL*

FUEL PRESS+ BAT DISCON FUEL PRESSHYD PRESS+ EXT POWER HYD PRESSXMSN OIL P+ XMSN OIL P

GEN DISCON+ GEN DISCONINVERTER*** INVERTER***

PITOT HTR PITOT HTR

* only when the FADEC is switched off ** only when the fuel pumps are off or running dry *** only if the respective system is installed+ only these cautions trigger the INP FAIL, if they are not active duringthe test.

If an error occurs during the test, INP FAIL will appear at the bottomedgeofcolumnMISCandayellowbaraboveandbelowtherespective

caution will flash. The corresponding caution will appear on the CAD. After 8 seconds, the ACK NEEDED prompt is displayed on the upperVEMD screen.

In case of a malfunction the respective caution will flash with a yellowbow, above and below. This message has to be acknowledged by theCDS/Audio Reset or the select button.

Test Pattern

If the switch TEST CDS/WARN UNIT or TEST CDS/WU is set toposition CDS, a test pattern appears with Cyclic Redundant Code(CRC), part number and configuration file number.

Cyclic Redundant Code

Check sum for the configuration file deviations (manufacturer only).

Part NumberLast two digits of the part number identify the software version.

Example:

B19030GB05 corresponds to software version V2001A

Configuration File

All software versions are delivered with a basic configuration file.Necessary changes (e.g. after installation of optional equipment)might require the upload of a customized configuration file deliveredby EUROCOPTER.

Example:

Software version V2001A,Basic configuration file L316M30S0001Customized configuration files L316M30SXXXX

NOTE The CPDS description shows the latest standards.

Major changes with part numbers and serial

numbers are shown in an overview page at the endof the CPDS description.





Test Pattern (Example Software Version V2001A)

(1)

(8)

(1)

(7)

(8) (1) (8) (1)

(1)

(4) (3) (5)

(2) (6) (1) (8)(8)

B19030GB05(8)

4E2F60A6(8)

(1) (1)L316M30S0001

(8)

Cyclic Redundant CodePart Number

Configuration File Number

1 white2 yellow

3 cyan4 green5 magenta6 red7 blue8 black





CPDS--Architcture for N1( n

N1), TOT, TQ

ENGINE 1 ENGINE 2

N1Sensor

N1Sensor

FADEC 1 FADEC 2

CAD

VEMD

Module 1

VEMD

Module 2

CROSSTALK

AnalogSignals

N1 Analog N1 Analog

N1 /TOT/TQDigitalRS 422

N1 /TOT/TQDigitalRS 422

N1 Duplex

TQ

nN1 (only TM)

TOT (only PW)

AnalogSignals

TQ

nN1 (only TM)

TOT (only PW)

TOT MatchingResistor (only TM)

TOT MatchingResistor (only TM)

(UpperScreen) (LowerScreen)





CPDS ModesGeneral

The following modes are available:

Flight Mode

-- CAU/Fuel (Caution and Fuel Page)-- FLI (First Limit Indicator)

-- ELEC/VEH (Engine and Electrical Parameters)-- Flight Report-- System Status-- Caution Fuel Fail-- CAU Backup

Ground Mode (Engines Shut Down)

In addition

-- Maintenance Menu-- Configuration (AC Config Page)

CAUTION / FUEL -- PageThe CAUTION / FUEL page is displayed automatically on the CAD.The fuel quantity parameters are displayed only on the CAD and areno longer available if the CAD fails. The units of measurement on thispage can be changed in the configuration mode (A/C CONFIG page).

The cautions inform the crew of defects in onboard systems. Theyappear in yellow characters in the three columns of the upper half of

the CAD. The columns are divided as follows:-- Left column: messages relating to eng. 1 and system 1-- Center column: messages relating to non-redundant

systems-- Right column: messages relating to eng. 2 and system 2

Cautions are listed in the order of their appearance (i.e. oldest cautionat the top). If there is not enough room on the page to display all the

cautions, e.g., “1 of 2” will appear at the top of the center column toindicate the presence of a second page with cautions. This page canbe accessed with the SCROLL key, but there will be an automaticreturn to page 1 after 15 seconds.

When a new caution appears, all the acknowledged cautions ondisplay will disappear, and a yellow bar will flash above and below thenew caution. At the same time, the MASTER CAUTION caption nextto the warning unit will illuminate.

The crew have to acknowledge the caution(s) by operating theCDS/AUDIO RES switch on the cyclic stick or the SELECT key on theCAD. If the CAD has failed, the SELECT key on the VEMD must bepressed. This leads to all cautions being displayed normally in theorder of their appearance. Also, the MASTER CAUTION caption willextinguish and is free for the next error message (caution).




NOTE Cautions with the letters CT at the end indicate

negative continuity test of the respective cautioncircuit only.

NOTE If the CAD and one VEMD screen fail only a

degraded Caution list is available on the remaining

screen (see respective FLM).

Advisories

The advisories appear in green characters below the cautions in theMISC column and provide the crew with information about theoperational status and optional equipment.

In certain cases, instead of being displayed on the first page, theadvisories may be displayed on the pages following pages. If a newcaution appears, the advisories will disappear until the caution hasbeen acknowledged. The green advisories appear initially in the lowerpart of the display fields and then form a column, one after another,under the cautions.

The following advisories are possible (depending on optionalequipment):

BLEED AIR Bleed air supply has been activated AIR COND Air conditioning system is active

HOOK UNLD No load on load hookS/L LIGHT Search and landing light is active

S/L LT EXT Search and landing light is fullyretracted

IFCO IFCO filter is activeIR ON The IR--screen of the SX 16 is activeSAND FILT Sand filter is active

AUX XFER Auxiliary tank fuel valve openTRAIN ARM Training mode is active (T2, P2 only)





First Limit Page (FLI) P1/T1

The FLI page is displayed on the upper VEMD screen. It contains thefollowing data:

-- FLI zone for TOT, N1 (nN1 with T1), TRQ-- Mast moment indication-- Message zone

Mast Moment Indicator

The mast moment indicator indicates the bending moment of the mainrotor. When entering the yellow range (50% MM) a yellow line appearsunder the letters MM. When entering the red range (66% MM) the linereverts to red, the LIMIT symbol and the warning GONG come on.

The time of exceedance and the maximum value (last flight andaccumulation) can be displayed in the maintenance mode.

NOTE A logbook entry and maintenance action is

required if the red region has been entered.

Periodical maintenance action is required if a

helicopter is operated without or with a defectivemast moment system.

Message Zone

The message zone displays messages concerning failures anddetected overlimits that are either not visible on the current displaypage or require action by the crew e.g. to switch off a screen.

The following list shows the messages in the order of their priority:

-- LANE 1 FAILED PRESS OFF1

-- LANE 2 FAILED PRESS OFF2-- CAD FAILED PRESS OFF-- CAUTION DETECTED-- VEH PARAM OVER LIMIT-- GEN PARAM OVER LIMIT (normal during engine starting)-- BAT PARAM OVER LIMIT

-- DC VOLT PARAM OVER LIMIT-- CROSST TALK FAILED PRESS OFF2-- VEMD BRIGHTNESS CONTROL FAILED-- CAD BRIGHTNESS CONTROL FAILED





First Limit Page P1/T1 (Example TM 2B1)

FLI DEGR

FLI FAILFLI DEGRmay appear asCAUTION on both sides

Solid white rectanglemarks the parameter represented

by the pointer

ENG FAILFADEC FAILENG MANUIDLETRAIN

TRAIN IDLEmay appear as CAUTION on both sides


Message Zone

LIMIT WarningLIMIT Counter

T

ENGFAIL





FLI ZONE P1/T1

The engine 1 and 2 parameters are generated by the two FADECsystems and are displayed on the screen as numerical values with thecorresponding measurement units.

In addition, the parameter that is nearest to its limit is displayed as ananalog pointer on a scale (i.e. First Limit Indication) and the numericalvalue of the parameter indicated by the pointer is marked by a whiterectangle.

If a parameter fails, it is displayed in yellow characters without itsassociated numeric value.

Limit Light/Counter

AEO above MCP

Five seconds before the 5 min power (AEO) time limit is reached thered box, the limit light and the counter appear and the box flashes.

When the time limit is expired, the red box is fixed.

OEI above MCP

When entering the 2.5 min power (OEI) the counter appears

immedeately. The limit light and the red box come on 5 sec before thetimelimitisreached.Theboxflashesandbecomesfixedwhenthetimelimit is expired.

When the pilot leaves the limited range the limit box and the audio tonestay for another 5 sec.





FLI -- Marking Symbology on Analog Display P1/T1 (Example TM 2B1)

Max. TOT starting (appears only during starting)

TOT starting transient (appears only during starting)TM max 5 sec, PW max. 2 sec.

AEO Max. Takeoff Power

OEI Max. Continuous Power

OEI 2.5 min Power

OEI Transient, max. 20 sec

AEO Take-off Power Range, max. 5 min

T Training Mode activated

T





First Limit Page (FLI) P2/T2

The FLI page is displayed on the upper VEMD screen. It contains thefollowing data:

-- FLI zone for TOT, N1 (nN1 with T2), TRQ-- Mast moment indication-- Message zone

Mast Moment Indicator

The mast moment indicator indicates the bending moment of the mainrotor. When entering the yellow range (50% MM) a yellow line appearsunder the letters MM. When entering the red range (66% MM) the linereverts to red, the LIMIT symbol and the warning GONG come on.

The time of exceedance and the maximum value (last flight andaccumulation) can be displayed in the maintenance mode.

NOTE A logbook entry and maintenance action is

required if the red region has been entered.

Periodical maintenance action is required if a

helicopter is operated without or with a defectivemast moment system.

Message Zone

The message zone displays messages concerning failures anddetected overlimits that are either not visible on the current displaypage or require action by the crew e.g. to switch off a screen.

The following list shows the messages in the order of their priority:

-- LANE 1 FAILED PRESS OFF1

-- LANE 2 FAILED PRESS OFF2-- CAD FAILED PRESS OFF-- CAUTION DETECTED-- VEH PARAM OVER LIMIT-- GEN PARAM OVER LIMIT (normal during engine starting)-- BAT PARAM OVER LIMIT

-- DC VOLT PARAM OVER LIMIT-- CROSST TALK FAILED PRESS OFF2-- VEMD BRIGHTNESS CONTROL FAILED-- CAD BRIGHTNESS CONTROL FAILED





First Limit Page T2 (P2 highly similar)

FLI DEGR

FLI FAILFLI DEGRmay appear asCAUTION on both sides

Solid white rectanglemarks the parameter representedby the pointer

ENG FAILFADEC FAILENG MANUIDLETRAINTRAIN IDLE

may appear as CAUTION on both sides


Message Zone

LIMIT WarningLIMIT Counter

ENGFAIL

T




ENG EXCEED C ti i t d t ll i i f 3 ll ith f ll i l




ENG EXCEED Caution

EC135 T2

The ENG EXCEED caution appears on ground under the followingconditions:Exceedance of a single time excursion in a OEI power band (2’ or 30’’).

Significant exceedance of the 30’’ power band with reaching andmaintaining the following values for more than 5 seconds: 136% Tq,4.8%nN1 (only possible in case of topping function failure) or 1024 °CTOT.If due to the cumulated total time in on or both OEI power bands anyengine parameter does not allow a minimum of 3 pulls with full singleexcursion time, i.e. if the remaining total time is less than 90s and 360sfor the 30’’ and 2’ OEI power band respectively.

EC135 P2

The ENG EXCEED caution appears in flight under the followingconditions:Significant exceedance of the 30’’ power band with reaching andmaintaining the following values for more than 5 seconds: 133% Tq,104.3% N1 or 990 °C TOT (only possible in case of topping functionfailure).

Exceedance of a single time excursion in a OEI power band (2’ or 30’’).In the latest FADEC software version the caution disappears when the

respective power band is left.The total allowed time in a OEI power band is expired.

The ENG EXCEED caution appears on ground under the followingconditions:

If due to the cumulated total time in one or both OEI power bands any

engine parameter does not allow a minimum of 3 pulls with full single

excursion time i.e. if the remaining total time is less than 90s and 360sfor the 30’’ and 2’ OEI power band respectively.

NOTE The ENG EXCEED caution is stored in the FADEC

and appears at the next engine start up.

Warnings

LIMIT symbol with box and audio warning GONG

Two different limit conditions for the activation of the LIMIT light withbox and the audio GONG are possible:

-- A LIMIT symbol with box activation due to OEI/AEO

time limit exceedance.

As soon as only 5 s of the allowed time in either power band (5’, 2’ or30’’) are left, a LIMIT symbol with a blinking red box appears. Thisprovides the pilot with a precaution that the allowed time within the

power band is about to expire. If the allowed single time excursion isconsumed (counter reaches 0), the box stops blinking, turns intosteady state. The audio GONG is triggered.

-- A LIMIT symbol with box and activation due to limitingvalue exceedance.

Exceedance of one of the engine or H/C limiting parameters (30’’Power, 5’ take-off power, mastmoment) triggers the LIMIT symbol with

the box in the steady state together with the audio signal at once. NOTE Whenever red limit has been reached or an

exceedance is evident, a logbook entry andmaintenance action is required. Depending on time

and maximum value the lifetime of the major

components can be reduced or totally expired.


Digital Data Display




Digital Data Display

A value within thenormal operatingrange.

A solid white rectangle associated with a parameter indicates the parameter shown by the needle.

If operation in a yellow range is detected, a countdown timer is automatically switched on and thedigital data is yellow underlined.

If operation in the red range is detected, the red underlining of the digits flashes.

If a parameter is invalid,the numerical value disap-pears and a yellow failuresymbol appears.


Page for Electrical and Engine Parameters




Page for Electrical and Engine Parameters

(ELEC/VEH)

The page for the parameters of the engines and of the electricalsystem are displayed automatically on the lower VEMD screen. Theunits for the various parameters on this page can be selected in theconfiguration mode.

The following parameters ca be displayed:

-- Outside air temperature OAT-- Load on cargo hook/cable length external hoist (options)-- Voltage and current-- Oil pressure and oil temperature of the engines and of the

main transmission

The voltage and current indication automatically shows the voltage ofthe generators. This setting can be changed to generator current or

battery current (i.e. BAT display) by operation of the SELECT and +and -- keys. If a value is invalid, “XXX“ is displayedin yellowcharacters.

The oil pressure and temperature indication consists of a vertical barwith upper and lower limits for each parameter and a numeric displaywith associated unit of measurement.


Engine and Electrical System Parameter




Engine and Electrical System Parameter

kgHOOK

Generator Field:DC [V], GEN [AMPS], BAT [AMPS]

Outside Air Temperature

External Load: [kg, lb]

Bar Graph Markingsfor Pressure and Temperature[bar. psi. °C]




ELEC/VEH -- Bar Graph Display




p p y

Normal operation range Warning range, the numericvalue is yellow underlined

Maximum range, the numericvalue is red underlined(blinking) and the yellow and

red markings grow

If there is an unvalid para-meter, a yellow symbolappears


FLIGHT REPORT Page




g

TheELEC/VEHpagewill automaticallyswitch to the FLIGHT REPORTpage only if the engine N1 RPM drops below 50 % and the oil pressurein the main transmission is less than 1 bar.

The page contains the following data:

-- Flight number and flight duration-- Gas generator turbine cycles

-- Power turbine cycles-- Impeller cycles (Pratt&Whitney only)

The page is automatically cleared upon initiation of the next startphase.

Returning from this page to the nominal page is possible only byoperating the RESET key.


FLIGHT REPORT Page




Total Number of Cycles N1

Total Number of Cycles N2

Engine 1 Engine 2

Number of Impeller Cycles (PW)

Total Number of Impeller Cycles (PW)

Refers to Mast Moment

Number of Cycles N1

Number of Cycles N2

Duration of the last flight

OVER LIMIT DETECTEDFAILURE DETECTED


SYSTEM STATUS Page




The SYSTEM STATUS page is displayed on the lower VEMD screenand is called up by way of the SCROLL key. FADEC data from therespective engines are displayed.

The units for the various parameters on this page can be selected inthe configuration mode.

The MSG and FAIL lines display messages and error codes. These

lines can be accessed individually with the SELECT key. When a lineis selected, the + or -- key can be pressed to continuously cycle thecurrent messages and error codes for FADEC 1 and FADEC 2simultaneously in their respetitive order.

The values of the parameters of FADEC 1 and FADEC 2 are displayedbelow the MSG and FAIL lines and are continuously updated.


SYSTEM STATUS Page (TM)




EGTTRQtrim

Engine Inlet Air TemperatureCollective Pitch Position

N2 Ref. Speed Trim ValuePower Turbine RPMTorque TrimExhaust Gas TemperatureFADEC Ambient Air Pressure

--

+ or --

--SCROLL

XXXXXX MSG XXXXXX XXXXXX FAIL XXXXXX--

SELECT

--

+ or --

--SCROLL

--SELECT

activates “systemfailure” function

back to the previous page

MSG Indication: IDLE.....FADEC Failure Codes







SYSTEM STATUS Page (PW)




FADEC Ambient Air PressureCollective Pitch PositionTorque Gain TrimPower Turbine RPMN2 Ref. Speed Trim ValueEngine Inlet Air TemperatureTOT TrimN1 DerivatedTorque Match

MSG Indication: IDLE.....


CAUTION/BACKUP Page




The CAUTION/BACKUP page is displayed on the CAD only if theVEMD fails completelyor has beendeactivated. Thefollowing dataaredisplayed:

-- Cautions (degraded indication only)-- Advisories-- Numeric readout of fuel contents in main and supply tanks.

-- Engine 1 and 2 torque displays on analog scale withnumeric limiting values.

If a torque channel fails, the associated pointer and numerical readoutare faded out; the scale and TRQ parameter turn yellow.

As this page represents an emergency operating mode, no otherpages or data can be presented.


CAUTION/BACKUP Page




CAUTION/ADVISORY Half Page

BACKUP Page

Supply Tank 1 Supply Tank 2

Main Tank


CAUTION/FUEL FAIL Page

Th CAUTION/FUEL FAIL i di l d i ll h




The CAUTION/FUEL FAIL page is displayed automatically on thelower VEMD screen if the CAD has failed.

At the same time thenN1 information in the FLI (Turbo Meca Versionsonly) is lost and the FLI DEGR caution is triggered in the FLI and in thecaution couple page in the system I and system II column.

As the fuel information is only available in the CAD the caution couplepage shows an empty yellow box where normally the fuel quantity isdisplayed. Furthermore only a degraded caution list is available,indicated by CAU DEGR in the miscellaneous field.


CAUTION/FUEL Fail Page (Example TM)




FLI DEGR FLI DEGRCAU DEGR


CPDS Switch Over Functions

G l




General

Depending upon how many screens of the CPDS are available, thepages on the CAD and VEMD can be switched manually andautomatically.

Three operating modes of the CPDS are possible:

-- Nominal mode (3 screens available)

-- Derivative modes (2 screens available)-- Backup mode (1 screen available)

Normal Mode

In the normal mode all three screens are operative. All pages areavailable in a variety of combinations, except the CAUTION/COUPLE

page.The pages can be selected manually via the SCROLL key.

If the RESET key on the VEMD is pressed, the standard pages willreappear on the screen.


CPDS -- Normal Mode




CAUTION

FUELFLI

SYSTEM

STATUS

CAUTION

FUELFLI

CAUTION

FUELFLI

ELEC

VEH

FLIGHTREPORT

CAUTION

FUELFLI

ELEC

VEH

SCROLL

Normal mode in the phases “shut--down, start, relight, flight”

Exception: when shifting from “flight” to “shut--down” phase

RESET

automatically


Derivative Mode with one VEMD Line off

The VEMD consists of a housing with two integral screens and two




The VEMD consists of a housing with two integral screens and twoprocessing modules (lanes) which are each plugged into one of thescreens within the housing. Although they are logically linked, theycanalso operate independently of each other. Therefore, if a screen or aprocessing module fails, the part of the VEMD that is still functioningwill still be able to present the most important data.

If one of the VEMD screens fails in flight, the FLI page will continue to

be displayed on the intact VEMD screen, the CAD will display theCAUTION/FUEL page (degraded caution indication), and theELEC/VEH page will be available when the SCROLL key is actuated.On the ground, the page SYSTEM STATUS can also be selected.

The FLI or CAUTION/FUEL pages will automatically switch to theFLIGHT REPORT page only if the engine RPM drops below 50 % andthe oil pressure in the main transmission is less than 1 bar.


Derivative Mode with one VEMD Lane off




CAUTION

FUEL FLI

CAUTION

FUEL

ELEC FLIGHTREPORT

FLI

-- flight phase -- shut -- down phase --ground phase

ELEC

VEH FLI

VEH

CAUTION

FUELFLI

ELEC

VEH FLI

ELEC

VEH

SYSTEM

STATUS

SCROLL RESET SCROLL

SCROLL

SCROLL


Derivative Mode with CAD off

The CAUTION/FUEL FAIL page will appear automatically on the lower




/ p g pp yVEMD screen.


Derivative Mode with CAD off




CAU

XXX

FLI

SCROLL

ELEC

VEH

FLI

SYSTEMSTATUS

FLI

RESET

SCROLL

FLIGHTREPORT

FLI

automatically

basicpage

ground phase


Derivative Mode with CAD and one VEMD Lane off

IfoneoftheVEMDscreensfailsinflight,theFLIpagewillbepresented




on the intact VEMD screen.

With the SCROLL button the CAUTION/FUEL fail page and theELEC/VEH page can be selected.


Derivative Mode with CAD and one VEMD Line off




FLI

SCROLL

CAU

XXX

SCROLL

ELEC

VEH

SCROLL

FLIGHT

REPORT

RESET

basic page

shut--down phase

automatically

flight phase


Derivative Mode with both VEMD Lines off

If only the CAD is still operating, the CAUTION/BACK--UP page isdisplayed




displayed.


Derivative Mode with both VEMD Lines off




CAU

BACKUP


Maintenance Menu

The maintenance menu is displayed on the VEMD (upper screen).Thesub menues provide access to flight and failure dates The following




sub menues provide access to flight and failure dates. The followingsub menues are possible:

-- Flight Report-- Failure (in preparation)-- Over Limit-- Funct. Times-- Trans Data-- Data Loading

The maintenance mode can only be entered when the engines aredetected in the “shut-down” state. The VEMD screens must beswitched off, the CAD must be switched on.


Maintenance Menu

simultaneous press theEntry to Maintenance Menu: press both keys




SCROLL

RESET

pfour keys and hold untilRELEASE KEY appears

OFF1

OFF2

SELECT

to scroll throughthe fields

Entry to Maintenance Menu:The operation must followwithin two seconds

OFF1

OFF2

p yto switch off

ENTER RESET

enters thesubmenus

DATALOADING

MAINTENANCE MENU

FLIGHT REPORTFAILUREOVERLIMIT

TRANS.DATA

EXIT

FUNCT. TIMES


Flight Report

Flight Report History Page




The Flight Report History page shows CPDS flight numbers from 1 to999 (starts from 0 again) and indicates duration of the respective flight.

Duration counting starts if:

-- N1 RPM engine 1 or engine 2 > 50%

-- XMSN oil pressure is > 1 bar

-- Angle of collective lever CLP > 28.5 (TM) or 17% (PW ).

The Flight Report History can only be entered when the ground stateis detected. The page stores the last 32 flights with failures. They areselectable with the + / -- button.

NOTE No. 1 flight is always the latest flight.




Overlimit

The Overlimit page shows the last 8 flight numbers (0--999) . Byselecting one flight number two counters (Mast Moment higher than




66% and Mast Moment higher than 78%) together with the maximumvalue are displayed for the respective flight. In addition the cumulatedtime for both ranges is shown in two lines below.


Overlimit Menu Page




OVERLIMIT MENU

SELECT NUMBER AND ENTER

FLIGHTNUMBER.

215

MM OVERLIMIT FLT NO. 215

LIMIT TIME MAX

MM > 66% ACC. TIME:

MM > 78%0 mn 15 s0 mn 12 s

68.9 %

31 mn 12 sMM > 78% ACC. TIME: 02 mn 12 s

EXIT PRESS RESET

214213212211

210209208

MM > 66%79.9 %


Transfer Data

Transfer Data is used to copy data from one VEMD lane to the otherin case one of the processor modules has been changed or a

fi ti diff b t th l h b




configuration difference between the processor lanes has beenindicated.


Transfer Data Page




TRANSFER DATA

TRANS. DATA : 1(L) 2(R)

YES

TRANS. DATA : 2(R) 1(L)

NO /

EXIT PRESS RESET


Function Times

The function times page shows the current flights and function timesfor the VEMD modules 1 and 2 and the function times for the CAD.





Function Times Page




FUNCTIONAL TIMES

MODULE 1 FLIGHT TIMES: XXXXXX h

EXIT PRESS RESET

MODULE 1 FUNCT. TIMES: XXXXXX h

MODULE 2 FLIGHT TIMES: XXXXXX hMODULE 2 FUNCT. TIMES: XXXXXX h

CAD FUNCT. TIMES XXXXXX h


Data Loading

With Data Loading a customized configuration file can be uploaded(e.g. modified caution list).




NOTE With the Avionique Novelle Configuration Tool

(software, board for PC, connecting cable tomaintenance connectors) the customer can upload

modified configuration files prepared by

EUROCOPTER. The actual software version

remains unchanged, only the basic configurationfile will be overwritten.


Data Loading Page




DATA LOADING


A/C CONFIG Page

The A/C CONFIG page is displayed on the VEMD (upper screen). Theselected setting option in an equipment data field is modified by the +and -- keys. The next data field to be modified canthen be selected withh SELECT k Th difi d fi i i d b l i h

The CONFIG mode can only be entered when the engine is detectedin the “shut-down” state and the VEMD screens must be switched off,the CAD must be switched on.




the SELECT key. The modified configuration is stored by selecting thedata field VALID with the SELECT key and then pressing the ENTERkey. The system then skips back to the standard MENU page.However, if the data field ABORT is selected and the ENTER key ispressed, the options in the data fields remain unchanged and the

standard MENU page is displayed again.The following parameters can be set on the A/C CONFIG page:

-- AUXILIARY FUEL TANK (N/I),Signifies whether or not an auxiliary fuel tank is installed.

-- BATTERY TEMP.PROBE (N/I),Signifies whether or not a temperature sensor is installedfor battery.

-- SECOND BATTERIE (N/I),Signifies whether or not a second battery is installed.

-- EXTERNAL LOAD (N/I); HOOK, CABLESignifies whether or not a cargo hook or an external hoist isinstalled.

-- FUEL FLOW WITH SENSOR (N/I),Signifies whether or not a fuel flowmeter is installed.

-- FUEL UNIT (LITER), (kg), (lb), (US GALLON), (IMP GAL.)Signifies which unit of measurement is used to indicate thetank contents.

-- UNIT SYSTEM (SI), (IMPERIAL)Determines which system of measurement units is used.

Parameter SI IMPERIAL Altitude m ft

Temp. (TOT, EOT) °C °CRpm/Torque (N1,

TRQ)

% %

Temperature (OAT) °C °FFuel weight Kg lb

Fuel quantity l, US gallon, IMP.gallon

l, US gallon, IMP.gallon

Weights (general) Kg lb

Hour h h

Minute mn mnSecond s s

Electrical power W WFlow rate Kg/h, l/h, US gal./h,

IMP gal./hlb/h, l/h, US gal./h,

IMP gal./hPressure (EOP) bar psi


A/C CONFIG Page

Entry to CONFIG--Mode:The operation must followwithin two seconds

simultaneously press the

four keys and hold untilRELEASE KEY appears

press both keysto switch off




OFF1

OFF2

SELECT

ENTER

+

OFF1

OFF2

SELECT

--

SELECT

ENTER

to scroll throughthe fields

installed not installed valid/abort

and


CPDS Software Versions Overview

The major features of the different CPDS software versions and somechanges depending on h/c serial number are shown in the followinglistings:

V2001A (Part Number: ...05)Integration of PW 206B2 engine.

Mast moment over limit recording




Software Versions

Thesoftwareversioncanbeidentifiedwiththelasttwodigitsinthepartnumber (e.g. part number ...02 corresponds software version V1999).

V1999 (Part Number ....02)Basic Version for EC135 T1 (TM 2B1 engines) and P1 (PW 206Bengines).

Mast moment indication > 50% yellow range, > 78% red range.

Supply tank volumes reverts from blue into yellow if no transfer isprovided or if the supply tanks volumes are below a certain value.

V2000A (Part Number: ...03)Modified mast moment indication:> 50% MM yelllow underlined, > 66% MM red underlined and flashing,GONG, LIMIT in a red box)

Certified for TM engine upgrade 2B1A.

Modified FLI: P1/T1 Transient torque layout change (red dot from 12.5to 14)

V2000B (Part Number: ...04)

Generator current limitation change: Gen. Amps underlined yellowwhen reaching 180 A (before 200 A).

Certified for TM engine upgrade 2B1A_1 (TU45 installed).

Mast moment over limit recording.

CPDS configuration change possible via ARINC 485 bus included.

V2001B (Part Number: ...06)

Mastmoment exceedance can be deleted.

Certification of the TRAINING MODE (single engine) for EC135 P1(PW 206B engines) and EC135 T1 (TM 2B1 engines).

Caution FUEL is integrated.

V2002 (Part Number: ...07)

Certification for Training Mode (dual engine) EC 135 T2 (TM 2B2engines) and EC 135 P2 (PW 206B2 engines); integration of themodified fuel system.

NOTE For the certification status of the software version

and the respective features refer to Flight Manual.


H/C Serial Number Changes Overview

Up to SN 120

CPDS over temperature indication separate light (temperature sensoradjusted to 63 °C




adjusted to 63 C.

Voltage adjustment unit installed under lh/rh cover of the instrumentpanel.

SN 121 and up

CPDS over temperature indication integrated in the CAD caution list(temperature sensor adjusted between 51 and 55 °C.

Voltage adjustment unit installed in the sensor units under the cabinfloor.

SN 169 and up

Only CPDS cockpit is available.

SN 218 and up

Maintenance connector installed in front of the center console(possible retrofit back to SN 169).

SN 250 and up

Modified fuel system (increased volume, modified vent lines andindication system).


Warning Unit

General

The warning unit centrally monitors several systems and provides

Warning Indications

The warning unit accomodates eight warning indications. They appear




visual and audio indications of arising malfunctions.

The unit contains the indication and evaluation units for eachmonitored system as well as a power supply unit. One switch perengine facilitates closure of the fuel valve.

Power Supply

The warning unit is supplied by the ESSENTIAL BUSBAR 1 and 2 viathe overhead panel installed circuit breakers:

-- WARN SYS I-- WARN SYS II

Test

To test the function of the indicator lights and also the audio warnings,a test switch TEST/WARN UNIT is installed in the overhead panel.

red when illuminated and black when inactive.Eachwarning indicationsimultaneously initiates a gong.

All warning indications may be dimmed with the potentiometerINSTR DIMBRT after engaging the associated switch on the overhead

panel.The significance of the warning indications is outlined in the respectivesystem chapters. The following are displayed:

-- LOW FUEL 1-- LOW FUEL 2-- AP. A. TRIMM (Autopilot)-- ROTOR RPM-- BAT TEMP-- BAT DISCH (Battery discharged)-- XMSN OIL P-- CARGO SMOKE


Warning Unit

Safety Guard




FIRE WARNING Eng. 1EMER OFF SW 1Press to release

EMER OFF SW releasedshut off valve is closed

white rim is visible

EMER OFF SW pressedshut off valve is openwhite rim is not visible

EMER OFF SW 1Illuminates together withinstrument lights

ACTIVEIlluminates white, if the EMEROFF SWITCH has been released

Side-view EMER OFF SWITCH


AP. A. TRIMM

The warning AP. A. TRIMM indicates a failure of the autopilot system.It is illuminated for 10 seconds. The signal is triggered by the autopilotcomputers.

R RPM

NOTE BAT DISCH appears if the voltage of the EPU is

below the voltage of the battery and the battery is

discharded via the ESSENTIAL BUSSES only.

XMSN OIL P




Rotor RPM

The ROTOR--RPM warning monitors a total of three limit values. Itreacts in various ways depending on which limit value is exeeded ordropped below.

-- ROTOR RPM < 95% A steady red indication of ROTOR RPM and a pulsed toneis generated. (The pulsed tone can be switched off with

AUDIO RES.)-- Rotor RPM ²106%

The red indication ROTOR RPM flashes and a gong can beheard. (The gong can be switched off with AUDIO RES.)

-- ROTOR RPM ²112%The red indication ROTOR RPM flashes and a continuoustone is generated. (The tone cannot be switched off)

BAT TEMP

The red indication BAT TEMP comes on when there is a batteryovertemperature detected (above 70 °C).

BAT DISCHThe red indication BAT DISCH comes on, when the battery isdischarded more than 2 ampers.

The red indication XMSN OIL P comes on when the oil pressure in themain gearbox is below 0,5 bar.

CARGO SMOKE

The red indication CARGO SMOKE appears, when there is a signalfrom the smoke detector in the rear cargo compartment (optional).

FIRE--Warning with EMER OFF --Switch

The unit consists of the fire warning logic circuit, FIRE indication withswitch EMER OFF SW 1 and ACTIVE-indication resp. FIRE indicationwith switch EMER OFF SW 2 and ACTIVE-indication. The fire warninglogic circuit displays individual fire warnings for engine 1 and engine2 and if necessary activates the fire extinguisher system. Operation ofthe switch EMER OFF SW 1 cuts the fuel supply to engine 1 and the

ACTIVE indication illuminates. Switch EMER OFF SW 2 cuts the fuelsupply to engine 2.

N1 RPM Monitoring

The N1 RPM is monitored for both engines separately. If the speeddrops below 50 % signals are sent to the CPDS/CDS and

-- the ENG FAIL caution is triggered-- the bleed air is switched off-- the fire extinguisher system is activated, if a fire warning is

evident.


Warning Unit -- Adjustment

J1 J2




N1 50%

RTR 106 %RTR 95 %

RTR 112 %

Top View






Switch Unit

General

A number of switches are arranged in the switch unit. They areprovided for:

DC Power Control Switches

In the lower row of the switch unit the DC power control switches areinstalled. These are:




provided for:

-- Engine control (upper row)-- DC power control (lower row)

Engine Control SwitchesFor starting the engines two switches for each engine are provided:

-- FADEC--switch (positions OFF--ON)To power the respective engine electronic system.

-- ENGINE START SWITCH (positions OFF--IDLE--FLIGHT)For automatic engine start and FADEC controlledgoverning in ground idle or flight idle RPM.

To preventinadvertant engine shutdown, the engine start switches areprotected by a manually operated safety guard (to be closed afterengine start).

installed. These are:

-- Two switches (GEN I, GEN II) for generator control with thepositions NORM--OFF--RESET

-- One switch BAT MSTR to control the power supply from

the battery and from an external power source with thepositions ON--OFF--RESET.

NOTE The switch BAT MSTR must be in Position “ON”,

even when the helicopter is supplied by an EPU.


Switch Unit

Start Switch ENG 1FADECControl Switch

Training Selector Switchwith Safety Guard

FADECControl Switch Start Switch ENG 2




Start Switch ENG 1 Control SwitchENG 1

Control SwitchGenerator 1

Control SwitchBattery/Ext.Power

Safety Guard

ENG 2

Safety Guard

Control SwitchGenerator 1


Overhead Console

General

The overhead console which is part of the electrical power system isinstalled in the center of the cabin roof. Busbars and circuit breakers

-- AC busbar 1-- AC busbar 2

Consumers with low energy demand and vital consumers for




supplying the electrical consumers are installed in the overheadconsole. Several systems are activated and/or controlled by switchesin the overhead console.

ComponentsThe overhead console consists of four component brackets and thefront panel containing the components and the busbars on the rear.The front panel consists of three parts with a background lightning andthe labelling of the installed circuit breakers, switches and rheostats.

-- Bus system 1-- Bus system 2

-- Switch unit of the overhead panel

Bus Bars

The following bus bars distribute the current to the individualconsumers:

-- ESSENTIAL busbar 1 (PP10E)-- ESSENTIAL busbar 2 (PP20 E)

-- SHEDDING busbar 1 (PP10S)-- SHEDDING busbar 2 (PP20S)

Additionally max. two bus bars/inverters can be installed for ACvoltage (required for P&R SAS, weather radar, mechanical gyros...):

gyemergency conditions are connected to the two ESSENTIAL busbars.Further DC power consumers are connected to the SHEDDING busbars (not supplied when only the battery is available or in case ofdouble generator failure).

The overhead console is supplied with current by the PRIMARYbusbars 1 and 2 or by the BATTERY busbar.

The BATTERY busbar supplies the ESSENTIAL busbars 1 and 2.Further lines are lead from the electrical master box 1 and 2 to supplythe SHEDDING busbars 1 and 2.


Overhead Console (Example)

AC BUS I AC BUS II

Switch SHEDDING BUS

Switch BUS TIE ISwitch BUS TIE IISwitch AC BUSSEL




SHEDDING BUS I

ESSENTIAL BUS I

SHEDDING BUS II

ESSENTIAL BUS II


Switch SHED BUS

The Switch SHED BUS is a two position switch with the positionsNORM--EMER. The NORM position is protected by a safety guard,which has to be opened before switching in the EMER position.

-- NORMB th SHEDDING b b d (th l SBC1

Switch BUS TIE I / II

The switches BUS TIE I and BUS TIE II are three position toggleswitches with the positions NORM--OFF--RESET. The switches areprotected by a safety guard, which positions the switch in the NORMposition. The following functions are provided:

NORM




Both SHEDDING busbars are powered (the relays SBC1and SBC2 are closed) when the electrical systems aresupplied by a minimum of one generator or by an EPU.

-- EMERThis position is used in order to supply both SHEDDINGbusbars from the battery in case of double generator fail(the relays SBC1 and SBC2 are closed).

-- NORMUpon switching on the BAT MSTR, both bus tie contactorsas well as the battery contactor close in order to connectthe primary busbars and the battery busbar to each other.

-- OFFThe associated bus tie contactor opens/remains open inorder to separate the two primary busbars.

-- In order to reset fault messages and activated protectivefunctions after a bus tie contactor had openedautomatically by a system fault, the switch must be set toRESET before the contactor can be closed again by

selecting the NORM position.Switch AC Bus Select (if two inverters are installed)

The switch AC BUS SELECT is a three position toggle switch with thepositions NORM--INV 1--INV 2.

In position NORM each inverter supplies it’s own bus bar. In case ofinverter failure, the remaining inverter can be switched on in order tosupply both bus bars.


Overhead Console -- Switches and Controls (Example)




NORM

EMER

O

F

F

M

A

X


Pitot--Static System

General

The pitot-static system picks up the dynamic and static pressure of theambient air of the helicopter. A number of drain ports are provided toremove water from the lines Electrical heating elements prevent the

Function

The static ports supply static pressure to the vertical speed indicator,altimeter and airspeed indicator. Ram-air pressure from the pitot tubeand static pressure is supplied only to the airspeed indicator




remove water from the lines. Electrical heating elements prevent thepitot and static pressure pickups from ice accumulation.

Components

The system consists of:

-- Pitot tube-- 2 Static ports-- Ambient pressure sensor-- Static selector valve (only pilot’s side)-- Hose lines

-- Flight instrumentsLocations

The pitot tube is located on the forward RH/LH side of the fuselage.The static ports are located one on each side of the fuselage below theequipment deck. The static selector valve is located on the right-handside of the center part of the instrument panel.

The components are connected with hose lines.The pilot’s pitot-static-operated instruments are located on theright-hand side of the instrument panel.

and static pressure is supplied only to the airspeed indicator.

With the static selector valve it is possible to select the pressure supplyfrom ambient pressure to cabin pressure in case of polluted external

static ports.Pitot/Static Heating (Optional)

With the switch PITOT HEAT in the overhead panel the electricalheating for the pitot tube and the static ports can be switched on.

There are two different versions for the indication in the cockpit:

-- Version 1: A green advisory comes on in the CDS/CPDS if

the heating is switched on.-- Version 2: A yellow caution appears in the respective fieldof the CDS/CPDS if the the heating is switched off.

For the dual pitot/static system two heating systems with two switchesare installed.




Handling of the EC 135

Lifting

General

The helicopter can be lifted with main rotor blades installed or




The helicopter can be lifted with main rotor blades installed orremoved. As a result the helicopter has different center of gravitypositions. For lifting a hoisting device is necessary.

Procedure-- The hub cap must be removed.-- Carefully insert hoisting device with the stamp into the hub

cap support on the rotor mast and attach with bolt.-- Secure the bolt with the safety pin.-- Carefully lift helicopter while observing balance.-- Avoid jerky movements under all circumstances.

NOTE On early helicopter serial numbers the borehole in

the support might be rotated to 45 and the tool canonly be installed after the rotor blades have been

removed.

NOTE Older hoisting device models might be limited to

2000 kg.


Hoisting Device




Bolt

Hub Cap Support

Side View

Safety Pin

Borehole for BoltHoisting Device




Jacking and Shoring

Jacking Bracket




Jack

Tail Boom Support


Weighing

General After completion of the leveling and dimensional check the helicoptermust be weighed.

Tools

-- Determine individual weights of weighing bracket and of 2 jacking brackets.-- Attach 2 jacking brackets to the aft landing gear fittings.

Attach weighing bracket in the center of the front crosstube. Position one jack each with installed force measuringdevice below the jacking brackets and below the weighing




The following tools are necessary for weighing:

-- Two jacking brackets

-- One weighing bracket-- Three jacks-- Weighing device-- Spirit level-- Clinometer

Procedure

-- The helicopter must be placed on a even and solid surfacein a closed draft-free hangar.-- Remove ground handling wheels from helicopter.-- Establish empty weight condition of helicopter in

accordance with Flight Manual (FLM).-- If installed optional equipment according to Equipment List

(EL) is weighed with the helicopter, ensure that theequipment status is recorded at the time of weighing.

-- Ensure that prescribed filling quantities for lubricants andhydraulic fluid are observed. Defuel helicopter using its ownfuel pumps. After defueling appr. 9.45 l (2.5 gal U.S.) equiv.7.6 kg (16.7 lb) of non-consumable residual fuel remains inthe fuel tanks.

device below the jacking brackets and below the weighingbracket.

-- Jack helicopter.

-- Apply spirit level or clinometer on cabin floor and levelhelicopter in horizontal position.

NOTE If weighing is performed with electronic force

measuring devices, more exact measuring results

are obtained by means of several weighing

procedures.Between the weighing procedures the force

measuring devices are to be interchanged in the

counterclockwise direction. The final result of the

weighing procedure is the mean value measured at

the respective weighing point.

-- Read measuring values on the force measuring devicesand record the weighing result in the weighing report (Form204). Calculate net values and moments.







Leveling

GeneralThe helicopter is leveled and dimensions are checked in accordancewith specified procedure. This is to verify all design dimensions. Theleveling data sheet must be kept in the historical record for futurereference. This procedure must be repeated after major modifications




jor repairs after hard landings.

Procedure

The following activities must be performed:

-- Ground the helicopter.-- Remove external equipment if installed.-- Defuel the helicopter.-- The helicopter must be placed on a even and solid surface

in a closed draft-free hangar.

-- Level the helicopter.-- Check the horizontal and vertical measering points.-- Check the angles.-- Record all measuring results in the leveling record.


Measuring Points

X 1766 Y 0 Z --X 5656 Y 0 Z --

X -- Y 0 Z 2800

X 3940 Y -- Z 2350

X 3940 Y Z 2350

1

2

4

5

3




X 3940 Y -- Z 2350

X 5400 Y -- Z 2350

X 5400 Y -- Z 2350X 2160 Y -- Z 1400

X 2160 Y -- Z 1400

X -- Y 1200 Z 2632

X -- Y -1200 Z 263211

10

6

7

5

9

8


Towing and Pushing

General

The EC 135 can be moved on ground by towing or pushing bymanpower.

Tools

-- Two transportation wheels

Pushing

For pushing the helicopter there are the following pushing points in thefuselage area:

-- Fenestron fairing, foreward and integrated control handles.-- LH and RH side shell below the engine deck




-- Two transportation wheels-- Towing bar

Procedure-- Install the two transportation wheels on the skid tubes and

lift the helicopter.-- Push the towing bar on LH and RH side on the skid tubes

and lock it by use of the fixing bolt.

NOTE For towing the helicopter at least one guide and

one person stabilizing the rear structure must be

available.

g-- LH and RH cabin structure-- Landing gear cross tube.

For pushing, the towing bar is not necessary.


Transportation Wheel and Towing BAR

H d li

Hydraulic Jack

BoltPushingPoints




Fixing Bolt

Skid Tube

Towing Bar

HydraulicJack Lever

Transportation Wheel


Parking and Mooring

General

To protect the helicopter from environmental influence, it has to becovered and tied down depending on weather conditions. With thehelicopterparked outdoors, it is recommended to moore the helicopterto the ground and secure the rotor blades by tie-downs.

Procedure

-- All the electrical equipment has to be switched off.-- The helicopter must be grounded at the ground connection

with the ground cable.-- Then all doors, windows and access doors must be closed.




Short-Time Covers

All short-time covers are stowed in a storage sack, which should be

carried on the helicopter during flights.The following short-time covers are available:

-- Short-time cover Fenestron-- Short-time covers engine outlet-- Short-time covers, NACA inlet-- Short-time covers, pitot tube

-- Short-time cover, front windows-- Short-time cover, NACA inlet roof-- Short-time covers, engine inlet-- Short-time cover, NACA inlet cowling

NOTE The engine outlets may be hot!

NOTE Attach the short-time covers with the notice

REMOVE BEFORE FLIGHT so that the notice flag is

clearly visible outside.

-- The main rotor is tied down with a lashbag to the tail boom.-- The main rotor has to be turned in direction of rotation until

one of the blades is aligned with the tail boom.

-- The lashbag must be fitted over the end of the blade andsecured to the tail boom by wrapping the attached belt andsack one full turn around the tail boom.

NOTE Turn the main rotor only in direction of rotation.


Covers

NACA Inlet

Main Rotor

Direction of main rotorrotation




Pitot Tube

Front Windows

Engine Inlet

Fenestron

Engine Outlet

Transport Sack

NACA Inlet

Ground Connection

EC 135Training ManualLifting System

Lifting System





Table of Contents

General Description of the Lifting System 4. . . . . . . . . . . . . . .

Main Rotor Drive 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Driveshafts 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Main Transmission 8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

LH and RH Drives 10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor Output Drive 12




Tail Rotor Output Drive 12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Main Gearbox 14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Lubrication System 18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .XMSN Oil Temperature Indication 20. . . . . . . . . . . . . . . . . . . . . . .

XMSN Oil Pressure Indication 20. . . . . . . . . . . . . . . . . . . . . . . . . .

XMSN High Oil Temperature Caution 20. . . . . . . . . . . . . . . . . . . .

XMSN Oil Chip Caution 20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

XMSN Low Oil Pressure Caution/Warning 22. . . . . . . . . . . . . . .

Oil Distribution System 24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fan Drive 28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Oil Cooling System 30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Main Rotor Hub Shaft 32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Mast Moment Measuring System 34. . . . . . . . . . . . . . . . . . . . . . .

Rotor Brake System 38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Rotor Brake Indication System 40. . . . . . . . . . . . . . . . . . . . . . . . .

Main Transmission Mounts 42. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ARIS Anti Resonance Isolation System 46. . . . . . . . . . . . . . . . . .

Oscillation Damper 52. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Main Rotor System 54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Main Rotor Blade 56. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Rotor Blade Adjustments 68. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .







General Description of the Lifting System

General

The lifting system of the EC 135 is located in the transmissioncompartment on top of the cabin roof, within the center-of-gravity area.Its main components are:

-- Main rotor drive

Rotor Brake System

The rotor brake system permits stopping of the main-- and tail rotor,after the engines have been shut down.

It mainly consists of:

A cockpit mounted brake lever




-- Rotor brake system

-- Main rotor system-- Monitoring system

Main Rotor Drive

The main rotor drive system transmits power from both engines to themain-- and tail rotor as well as to two cooling fans and two hydraulicpumps.

It consists of:-- 2 driveshafts

-- Main transmission

-- Main transmission mounts

-- A cockpit mounted brake lever

-- Flexball cable

-- Brake cylinder-- Brake caliper

-- Brake disk

Main Rotor System

The main rotor system generates the lift and thrust of the helicopter.In conjunction with the tail rotor system, it provides directional control

of the helicopter in flight. Driving forces and control inputs aretransferred to the rotating main rotor through the system components.

Monitoring System

For the important parameters (e.g. rotor RPM, oil pressure and oiltemperature) several sensors are installed. Thesignals are transferredto the cockpit in order to trigger warnings and supply the indicatinginstruments.


Lifting System -- General Arrangement

Main Rotor System




RH Driveshaft

LH Driveshaft

Tail Rotor Drive

Rotor Brake

Main Transmission Mounts

Main Transmission

Swash Plate


Main Rotor Drive

General

The main rotor drive transmits power from both the engines to the mainand tail rotors and the auxiliary units. Additionally it is a structuralcomponent of the helicopter and also transmits all static and dynamicloads between the main rotor system and the fuselage.

Driveshafts

General

Two driveshafts connect the engines to the freewheeling units of themain transmission. They transmit the power of the engines to the maintransmission In addition they correct for any variations in length or




Components Main Rotor Drive

The main rotor drive consists of the following:

-- 2 driveshafts


-- Main transmission mounts

-- Main rotor drive monitoring system

-- Rotor brake system

transmission. In addition, they correct for any variations in length ormisalignment between the engine outputs and the main transmission

inputs. For this purpose two flexible couplings are attached to eachend.

Two different versions (type Bendix or Kaflex) are available.


Engine Drive Shaft

Flange

Side View




Flexible CouplingShaft Tube

Flange

Flexible Coupling

Type Kaflex Type Bendix

Flexible Coupling


Main Transmission

General

The main transmission transfers the power from both engines to themain rotor system, tail rotor and the auxiliary units. All mounting points,attachment fittings and oil lines are integral with the transmissioncasing. Two freewheel units incorporated in the input drives allowpower to be transmitted only from the engines to the main

Leading Particulars

Weight approx. 143.5 kg

Gear reduction Main rotor 14.923

Tail rotor 1.183

Speed Drive 5898 RPM




p y gtransmission.

Components

The main transmission is of modular design. It mainly consists of:

-- LH and RH input drives

-- Tail rotor drive

-- Main gearbox

-- Lubrication and cooling system

-- LH and RH accessory drives

Speed Drive 5898 RPM

Main rotor 395 RPMTail rotor output 4986 RPM

Oil quantity approx. 8.0 l

Oil type O--156; MIL--L--23699C

Material Aluminium alloy


Main Transmission -- Modules

LH Fan Gearbox

Main Gearbox

LH Hydraulic Pump Drive

RH Hydraulic Pump Drive




LH Input Drive

RH Input Drive

Tail Rotor Drive

RH Fan Gearbox

FWD


LH and RH Drives

Assembly

The drive consists of:

-- Freewheel shaft

-- Freewheel unit

-- Cover shaft with seal

-- Ball bearing and roller bearing

Freewheel Unit

The engines drive the input drive shafts in clockwise direction. In thisdirection, the freewheel clutches are interlocking the driving and drivenparts.

The freewheel clutches are effective in the following situations:

-- Starting the engines: Only one turbine drives initially andh f h l l h h h d i i I ill l k




g g

-- Drive pinion

Function

The driveshaft connecting the engine to the main transmission isattached to the triangular flange of the free wheelshaft. The bevel gearof the drive pinion meshes with the bevel gear of the intermediateshaft. The correct gear mesh (gear backlash and gear tooth pattern)is ensured by placing a shim of the appropriate thickness between theball bearing and transmission casing. The shaft seal in the cover seals

off the rotating freewheel shaft at its outboards end.

the freewheel clutch to the other drive is overrun. It will lock

if both engines are running at the same RPM.-- One engine becomes inoperative: Its freeweel clutch is

overrun and prevents the engine from being driven by themain transmission.

-- Both engines become inoperative: Both freewheel clutchesare overrun and the main rotor can turn without anyadditional friction from the engine (autorotation).


Freewheel Assembly

Seal

Drive Pinion




Bearing

Bearing

Freewheel Unit

HousingEngine Shaft,Driving Part

Gearbox Drive Pinion,Driven Part

Clutch Free

Clutch under Load

Engine Shaft Stopped

Gearbox Drive Pinion,decoupled from Engine Shaft

Sense of

Rotation


Tail Rotor Output Drive

General

The tail rotor output drive consists of:

-- Connecting pad

-- Cover with shaft seal

-- Pinion

-- Ball bearing




g

AssemblyThe connecting pad provides the attachment point for the rotor brakedisc adapter and the tail rotor driveshaft. The pinion bevel gearmeshes with bevel gear of the pinion shaft. The correct gear mesh(gear backlash and gear tooth pattern) is ensured by placing a shimof the appropriatethickness betweenthe ballbearing on the pinion andtransmission casing.The shaft seal in the cover seals off the rotatingconnecting pad at its outboard end.


Tail Rotor Output Drive




O--Ring

Shaft Seal

Cover

Spacer

Connecting Flange

O--Ring

Screw

Pinion


Main Gearbox

Power Flow

The engines drive both input drives of the main transmission. The inputdrive bevel gears mesh with the bevel gears of both intermediateshafts.

An input shaft connects each intermediate shaft to its respective oilpump through a spline connection. The spur gears of the intermediateshafts drive the collector shaft, in which the main rotor hub--shaft is

Gearbox

The following assemblies are installed in the gearbox for the purpose

of transmitting power and reducing speed:

-- Two intermediate shafts each with an integral bevel gear, aspur gear for driving the collector shaft, a larger spur gearlocated below the bevel gear for driving the intermediatespur gear, and a spline for driving the oil pump.




splined to the inside.

Through its integral spur gear, the collector shaft drives the pinionshaft. The bevel gear of the pinion shaft meshes with the bevel gearof the tail rotor output drive.

The large spur gear on each intermediate shaft drives an idler gear.The idler gears in turn mesh with the spur gears of driveshafts anddrive both fan gearboxes. The hydraulic pump drives, which aresplined to these driveshafts, rotate at the same speed.

The bevel gears of driveshafts mesh with the output pinion gears. Theflange-mounted fans are positively splined to the output pinion gearsand rotate at the same speed.

-- A collector shaft with an integral spur gear which drives thepinion shaft of the tail rotor.

-- A pinion shaft with integral spur gear and bevel gear fordriving the tail rotor output drive.

The main rotor shaft is splined to the inside of the collector shaft andis held in position by a mast nut. Mounted on the upper casing of thegearbox is a support tube which surrounds part of the main rotor shaft.The support tube provides the sliding surface for the up and down

motion of the swash plate.


Main Gearbox -- Geartrain and RPM (at 100%)

Hydraulic Pump 2 Drive5146 RPMIntermediate Shaft 1 and Oil Pump 1 Drive

1696 RPM

Cooling Fan 1 Drive12666 RPM

Hydraulic Pump 1 Drive5146 RPM

Collector Shaft andMain Rotor Hub Shaft

395 RPM




Cooling Fan 2 Drive12666 RPM

Intermediate Shaft 21696 RPM

RH Input Drive5898 RPM

Tail Rotor Output Drive4986 RPM

LH Input Drive5898 RPM

Oil Pump


Main Gearbox, Lateral Cut, View in Flight Direction

Intermediate Shaft

Main Rotor Hub Shaft

Sliding Sleeve

Intermediate Shaft

Collector Shaft




Oil PumpOil Pump


Main Gearbox, Longitudinal Cut

Upper Roller Bearing

Inner Seal Ring

Upper Bearing Outer Race

Spacer Tube

Sliding Sleeve

Collector Shaft

Seal

Upper Bearing Inner Race




Tail RotorOutput Drive

Cover

Lower Ball Bearing

Lower Roller Bearing

Hub Shaft Nut

Hub Shaft NutLocking Device

Collector Shaft


Lubrication System

General

The main transmission is provided with a wet sump oil system forlubrication and cooling. Because of redundancy, the lubricationsystem comprises two oil pumps located in the lower casing of thegearbox. The main components of the system are:

-- Filler neck

-- Oil filter

Oil Filter

An oil filter located in the central oil passage separates thecontaminants from the oil. The housing of the oil filter is fitted with abypass valve (np 3.5 bar) and a mechanical filter contaminationindicator (np 2.1 bar). If the filter becomes clogged, the oil will bererouted through the bypass valve thereby maintaining the propersupply of oil to the system.

An oil pressure transducer measures the oil pressure in the central oil




-- Spray tubes

-- LH and RH oil pumps

-- Oil sight glass

Oil is added to the system via the filler neck. The oil level is indicatedby the oil sight glass. Oil is drained off through a valve which housesthe magnetic plug.

Oil Pumps

The main transmission is provided with a redundant lubrication systemcomprising two oil pumps located in the lower casing of the gearbox.These pumps are driven by the intermediate shafts throughinterconnected driveshafts. The oil pumps draw oil from the oil sumpand convey it through a central oil passage. If either pump should fail,the remaining pump is able to convey enough oil to meet systemdemands. Failure of an oil pump is detected by a low-pressure switch

and is visually indicated in the cockpit. In the central oil passage, anoil temperature transmitter measures the oil temperature and an oiltemperature switch monitors the max. permissible oil temperature.The associated indicators are located in the cockpit.

An oil pressure transducer measures the oil pressure in the central oil

passage. Visual indication of the pressure is provided in the cockpit.The oil is conveyed to both oil coolers and from there to the lubricatingpointsthrough theintegral oil passages in the casing. Installed at theselubricating points and accessible from the outside are spray tubeswhich provide for optimum lubrication of the components.

Oil Cooler

The oil coolers are mounted to the RH and LH side of the main

transmission. They are split into two sections. The smaller section of each cooler, which is connected directly to the main transmission,serves for cooling the main transmission oil (50% each side).

For this, ambient air is drawn by the cooling fans and forced throughthe oil coolers via air ducts. From there the air is directed overboard viaoutlet ducts (See also chapter “Power Plant”, Oil Cooling System).


Main Transmission -- Oil System

Temperature Transducer

Temperature Switch(Triggering at approx. 115 °C)

XMSN OIL T CDS/CPDS Caution

Oil Filter with By-pass Valve

Oil Pressure Transducer

Oil Cooler Bearings

Pop Out Indicator




Supply

Scavenge

Magnetic Chip Detector andDrain Valve

XMSN CHIPCDS/CPDS Caution

Oil Tank

Oil Pumps with By-Pass Valve

(opens at approx. 8 bar)Pressure Switch XMSN OIL PCDS/CPDS Caution

Check Valves

Pressure Switch XMSN OIL PCDS/CPDS Cau-tion

XMSN OIL P

Warning Unit


XMSN Oil Temperature Indication

General

The oil temperature of the main gearbox is measured by a transducermounted to the gearbox at the oil filter housing. The temperature isindicated in the cockpit on the oil temperature and pressure unit or onthe VEMD in °C.

XMSN Oil Pressure Indication

XMSN Oil Chip Caution

General

For the detection of magnetic chips in the oil system, a chip detectoris fitted in the common suction line of both oil pumps. It is installed bya bayonet connection in the XMSN oil drain plug (a check valve closeswhen the chip detector is removed).

Accumulation of particles bridge a contact gap of the detector magnetand close the circuit to the CDS/CPDS.




General

The oil pressure is measured by a transducer mounted to the gearboxin the central oil passage. The pressure is indicated in the cockpit onthe oil temperature and pressure unit in bar.

Minimum 0.5 bar

Continuous operation 0.5 to 7.8 bar

XMSN High Oil Temperature Caution

General

The oil temperature caution caption is triggered by an oil temperatureswitch installed at the main transmission oil filter housing. The switch

closes thecircuit to the CDS/CPDS at a temperature of approx. 115û

C.The indication at the MISC CAUTION display will be:

-- XMSN OIL T

The indication at the MISC CAUTION display will be:-- XMSN CHIP


Main Transmission -- Monitoring

FWDMain Transmission





Speed Pickup for Rotor RPMIndication and Warning

Chip Detector

Oil Temperature Switch


Oil Temperature Transducer


XMSN Low Oil Pressure Caution/Warning

General

To warn the pilot in case of low oil pressure in each of the XMSNlubrication systems two pressure switches are installed downstreamof the oil pumps. The switches are installed at the lower front side of the main transmission.

Low Oil Pressure Caution

Each oil pressure switch closes when the pressure at the associated




p p

pump outlet is below 0.5 bar.

The associated indication are as follows:

-- XMSN OIL P Caution SYS I or II on CDS/CPDS

Low Oil Pressure Warning

In case of low oil pressure in both XMSN lubrication systems (bothpump outlet pressure switches sense a pressure below 0.5 bar) a low

pressure warning will be sent additionally to the CDS/CPDS cautioncaptions.

The associated indications are as follows:

-- XMSN OIL P Caution SYS I and II on CDS/CPDS

-- XMSN OIL P Warning on the warning unit

-- Gong in the headset with 3 seconds intervals


Main Transmission -- Oil Pressure Switches




Oil Pressure SwitchSYS I

Oil Pressure SwitchSYS II




Main Transmission -- Componentsof Lubrication System

Oil Filter

Filler Neck

Spray Tubes

FWD

Contamination Indicator




Spray Tubes

Spray Tubes

LH Oil Pump

RH Oil PumpSight Glass

Oil Cooler


Main Transmission Oil Service

The following oil type is approved for the main transmission:

-- MIL--L--23699

The oil quantity is approx. 8.0 liter.

Oil Level Sight Glass

The main transmission oil level can be checked by a sight glass,located at the RH rear side of the main transmission.

The “MAX” and “MIN” marks indicate an oil level of approx. 9, resp. 7




liters.


Main Transmission -- Oil Service

MIN

MAX

Oil LevelSight Glass

Cap

O--Ring

FWD




Filler Neck

Adaptor for Oil Drain Hose

Chip Detector

Electric Plug

FWD


Fan Drive

General

A fan drive gearbox consists of:

-- Gearbox housing

-- Idler gear with ball bearing

-- Driveshaft with bevel gear and bearings

-- Output pinion gear with ball bearings

-- Cover with shaft seal




Configuration and Function

The intermediate shaft of the main gearbox drives the idler gear andthe driveshaft of the fan gearbox. The driveshaft is splined to thehydraulic pump. A cover fitted with shaft seals off the driveshaft at itsupper end where the hydraulic pump is connected. The bevel gear of the driveshaft drives the output pinion gear of the fan, which is runningin an oil bath.


Fan Gearbox

Flange for Hydraulic Pump




Idler Wheel

Driving Wheel

Flange for Fan Housing


Oil Cooling System

GeneralBoth engines as well as the main transmission of the helicopter areequipped with internal, independent oil circuits. These ensurepermanent lubrication and cooling of highly stressed componentsunder all operating conditions. To keep the oil temperature withinlimits, a oil cooling system is installed in the helicopter.

Independant cooling circuits are availble for the:

LH Engine

Oil CoolerThe oil coolers are mounted to the RH and LH side of the maintransmission. They are split into two sections. The smaller section of each cooler, which is connected to the main transmission by bushingsdirectly, serves for cooling the main transmission oil (50% each side).

The larger section of each cooler is connected to the associatedengine by oil hoses. This section serves for cooling the engine oil.

C li Ai Fl




-- LH Engine

-- RH Engine

-- Main Transmission

Components

The oil cooling system consists of the following:

-- 2 cooling fans

-- 2 inlet airducts

-- 2 outlet airducts

-- 2 dual section oil coolers (engine / main transmission)

-- 2 thermal controlled bypass valves in the engine circuits

-- several hoses and connectors

Cooling Fans

The cooling fans are mounted on the front side of the maintransmission RH and LH. They are driven by the main transmissiongeartrain (12665 RPM at 100%).

Cooling Air Flow

Ambient air which enters the air intakes is drawn by the cooling fansand forced through the oil coolers via the inlet air ducts. From there theair is directed overboard by the outlet ducts.


Oil Cooling System -- General Arrangement

Fan Drive

To/fromMain Transmission

FWD




Fan Drive

CoolingFan

Inlet Duct

Outlet Duct

Oil Cooler

To/fromEngine

Inspection Door



GeneralThe main rotor hub shaft transmits the driving moment to the main rotorblades which are connected to the hub. In doing so, it also performsthe function of a rotor head.

The main rotor hub shaft assembly consists of the followingcomponents:

-- Rotor hub shaft with integral flanges

Hub cap support

Bonding JumperFour bonding jumpers are screwed onto the hub cap support with oneend and to bonding studs at the rotor blades. This allows staticdischarge of the rotorblades.

Hub Cap Support

The hub cap support, which is manufactured from aluminum alloy, isattached by screws to the upper hub flange of the main rotor hub shaft,and seals off the open end of the hub shaft




-- Hub cap support

-- Rotor hub cap

Configuration

The main rotor hub shaft, which is hollow and is formed with two hubflanges at its upper end, is a one-piece forging made of steel alloy. Thehub flanges provide for the attachment and securement of the main

rotor blades.Formed 180° apart on the shaft are connectors which provide amounting for the rotating scissor clamps. On the lower end of the shaftare the seating surfaces for the mast bearings and the mast splinewhich meshes with the main transmission.

The upper hub flange is marked with the numbers 1 thru 4 at the bladeattachment areas, with the numbers ascending in the clockwise

direction. This identification is important for relating the bladeattachment areas to their respective blades.

and seals off the open end of the hub shaft.

The helicopter can be lifted by a hoisting device attached to the hubcap support.

Rotor Hub Cap

For aerodynamic reasons a rotor hub cap is installed. It is a compositeconstruction which can be delivered in two different types. There are

two different hub cap supports possible:

-- standard rotor hub cap

-- quick-removable rotor hub cap for blade folding system(optional)

Attachment to their respective hub cap supports is by screws in thecase of the standard hub cap and by bayonet connections and safetyscrews in the case of the quick-removable hub cap.



Standard Rotor Hub Cap

Standard Hub Cap Support

Bonding Jumper




Bonding Jumper

Rotor Hub Shaft

Spline

Thread for Shaft Mounting Nut

Upper Flange

Lower Flange

Connectors for Levers.Two off, 180 apart

Teflon Covered Bushings

Upper Hub ShaftBearing Seating

Cap




Mast Moment Measuring System

PP 20E

Circuit BreakerMAST MM

Signal Amplifier UnitSignal Processing Unit

Coupling Rotor--Stator




Strain Gauge Bridge

Cumulated Counter

Value

Green, Yellow, RedBars and Limit Light

MMEX XXXX

Mast MomentIndication VEMD







Mast Moment Measuring System

Strain Gauge Bridge(b d d i t th t)

Rotor

Sensor Amplifier Unit




(bonded into the mast)

Stator

Rotor

Lower Gearbox Cover

Signal Processing Unit


Rotor Brake System

GeneralThe hydro-mechanical rotor brake system enables the main and tailrotors to be brought to a standstill, and locks them against furtherrotation for a limited period of time. With the brake lever applied andlocked, the hydraulic pressure in the rotor brake system will bemaintained for a longer period of time before slowly dissipating. Anelectrical switch lights up a caption in the cockpit indicating system thatthe rotor brake has been engaged.

NOTE Th b k l b d d h

FunctionThe rotor brake is actuated by a brake lever. Before it can be operated,the brake lever must be released from its detent by actuating a lockingpawl which allows the brake lever to be pulled downward until itengages. The maximum force is limited by the damper spring after thebrake lever has reached the mechanical end stop. To release the brakelever, the locking pawl on the brake lever must be pressed.

NOTE The fluid reservoir must be filled with brake fluidDOT l




NOTE The rotor brake may only be operated under thefollowing conditions:-- the engines have been shut down,

-- the rotor speed is down to 50 % of its nominal

speed

-- OAT > --30

C

System Components

The rotor brake system mainly consists of:

-- Brake lever (located in the cockpit)

-- Bowdenflex cable

-- Damper (force limiter spring)

-- Brake cylinder with fluid reservoir

-- Brake caliper

-- Brake disk

-- Micro switch for CDS/CPDS caution ROTOR BRK

DOT--4 only.


Rotor Brake System

Reservoir for Brake

Fluid

Hydraulic Hose

Brake Support




Micro Switch

Brake Disk

Tail Rotor Drive

Shaft

Brake Caliper

Damper

Brake Cylinder

Brake Lever

Lever

Bowdenflex Cable


Rotor Brake Indication System

General

The rotor brake indicating system indicates an engaged rotor brake.For this a microswitch is installed at the brake caliper mountingslideway. The slideway itself is installed in the rotor brake support ina way that it can move laterally against a spring by approx. 1 mm.

If the rotor brake is engaged and the brake disk starts turning, thebrake caliper will move together with the slideway against the springand depress the microswitch.

The indication at the CDS/CPDS MISC caution display will be:




-- ROTOR BRK

NOTE With an engaged rotor brake and a stillstanding

rotor, the indication may be on because of

manufacturing tolerances and has to be checked.

It has to come on in the moment the rotor starts

turning and the brake is engaged.


Rotor Brake Indication SystemRotor BrakeSupport

Slide

Slide Bolt




Micro Switch

ROTOR BRK

Micro Switch

Break Support

Top View

Brake Caliper

SYSTEM I SYSTEM IIMISC




Main Gearbox -- Attachment

Vibration Isolator ARIS(Z A i )

FWD




Torque Strut(X Axis)

Side Load Strut(Y--Axis)

(Z Axis)

Emergency Stop

Emergency Stop







Gearbox Struts

Main Gear BoxGearbox Cover




Torque Strut

Bracket on the Transmission Deck

Side Load Strut

Bushing

Bushing

Bushing

Bolt


ARIS Anti Resonance Isolation System

Principle

In order to isolate a vibration between the rotor system and the aircraftfuselage the principle of the spring/mass damper is used.

The spring rate, the force transmitting unit and the mass weight haveto be defined in such a way the main rotor frequency induces the antiresonance oscillation in the spring/mass system. Thus the H/C rotorsystem and the damping mass vibrate with the same frequency, onlywith 180° phase shifted. Therefore the forces generated by the rotor

system in downward direction are compensated by the forces createdby the damping mass in upward direction and vice versa.




This system is only effective in the vertical axis (z--direction) andtowards the adjusted frequency.


Principle of Passive Anti--Resonance Vibration Isolation

Rotor induced forces

Vibration of transmission caused byrotational forces on the rotor system

Oszillations of the mass damper




No forces Equal forcesacting inoppositedirections

No forces No forcesEqual forcesacting inoppositedirections

Fuselage vibrations

Fuselage forces


General

The system consists of 4 uniaxial hydro-mechanical vibration isolaters.They carry all weight and lifting forces transmitted by the maintransmission. They are attached to the airframe by 4 bolts each andto the main transmission by a special spherical bearing and one bolteach. For “fail safe” purposes an emergency stop is mounted aroundeach damper.

The purpose of the system is to reduce the loads and vibrationsgenerated by the main rotor to the helicopter fuselage.

FunctionThe vibrations generated by the main rotor cause periodic movementsof the main transmission relative to the fuselage which in turn causes

At the upper end of the secondary bellows is a mass jacket to whichis attached a pendulum rod which acts asa guide for the mass and alsoaccomodates the additional weights.

A pre-loaded compression spring together with the secondary bellowsproduce an operating pressure within the self-contained unit of approx. 6 to 7 bar, thereby ensuring the functional integrity of thevibration isolator for all operating conditions.

The emergency stop which is formed in the shape of a cylindrical potand fits over the corrugated portion of the primary bellows is attachedto the transmission deck of the fuselage by screws.

If the primary bellows of the vibration isolator should fail, thetransmission will be supported either by the fixed stop ring or thed h bl i h




of the main transmission relative to the fuselage which in turn causesaxial movement of the primary bellows.

In response to the travel of the primary bellows, the secondary bellowsproduces a longer stroke as determined by the ratio of their respectivecross-section areas. The resultant inertia forces (force generator)cause the pressure of the glycol solution in the vibration isolator tofluctuate. The spring and pressure forces on the isolator attachmentpoint on the fuselage overlap each other. At the anti-resonancefrequency, this results in the forces transmitted to the fuselage beingcancelled and consequently the vibrations being reduced.

The primary bellows are provided with an adapter at the bottom endfor connecting them to the fuselage, while at the top end they areformed with a forked lug for connecting them to the main transmission.

The forked lug is fitted with bushings. Above the bellows section, theprimary bellows are formed with an integral ring above which is anannular groove which accomodates a split emergency stop ring.

detachable emergency stop rings on the emergency stop.

NOTE Earlier versions are equipped with a combination

of 6 aluminium / steel weights used for fine tuning.

In newer versions the pendulum rods are emptyand the mass jacket weight is higher. These

versions don’t require an adjustment.


ARIS -- Vibration Isolators

Secondary

Bellows

Water/GlycolSolution

Filling and Bleed Port

(Manufacturer only)

Pendulum with Tuning

Emergency StopRing (Splitted)

EmergencyStop

Vertical Movement of Mass/Spring Unit




PrimaryBellows

CompressionSpring

Mass Jacket

gMass (6 off)

Bearing Cage

with Bearings

Vibration Isolator

Locking ScrewPendulumProtrusion


Clearance

When ready installed the clearance between stop ring and emergencystop must be a certainmeasure. Formeasuring this clearence, a feeler

gauge is used at four places 90° apart and the mean value has to becalculated.

The clearance is adjusted with shims to the nominal value 1.0 --0.3 mmduring installation.

NOTE The clearance will change with the temperatureand therefore can’t be used for failure detection.

Adjustment




A main rotor speed of 100% NR means that the main rotor rotates at6.6 revers per second. This results in a 4/rev vibration frequency of 26.3 Hz. The natural vibration frequency of the ARIS is adjusted to thisfigure.

Failure Detection

At +20 °C the pendulum rod will prodrude for approx. 8--9 mm. Theprotrusion varies with the ambient temperature, but generally it can bestated, that as long as the pendulum rod protrudes the ARIS is stillserviceable.

In case of pressure drop (e.g. crack in one of the bellows) the internalspring and the inner bellows expand and the pendulum rod willdisappear.


ARIS -- Measurement of Clearance

Vibration Isolator

Stop RingMeasuring Points

Nominal Clearance




Emergency Stop

1.0 -- 0.3 mm

Main Transmission Deck

Shim


Oscillation Damper

General

The aircraft is equipped with a mass/spring damper to reduce lateralvibrations (y direction). It is mounted to the fuselage and compensatesfor lateral vibration from the main rotor system.

Location and Assembly

The y--damper is mounted to the stringer below the LH floor panel.

The damper assembly consists of two weights, which are adjustablefor mass, bolted to the springs. The location of the weights on thesprings is also adjustable. On each weight it is possible to attach upt 6 dditi l i ht ( dj ti h t ) f t i Th i ith

A main rotor speed of 101.5% NR means that the main rotor rotates

at 6.7 revers per second. This results in a 4/rev vibration frequency of 26.7 Hz. The natural vibration frequency of the y damper is adjustedto this figure.

NOTE If the H/C flies permanently in higher altitudes, the

efficiency of the damper can be adjusted by

removing a certain amount of tuning sheets(according service engineering information).




to 6 additional weights (adjusting sheets) for tuning. The springs, withthe weights attached, are mounted to a common support.

Function

The damper is energized by lateral oscillations of the fuselage. The

natural frequency of the damper can be adjusted by adjusting themass of the weights or moving the weights on the springs. If thedamper frequency is tuned to the same frequency as the fuselageoscillations, it will vibrate in exact opposition to the fuselage vibrations.This induced vibration of the damper will react in direct opposition tothe fuselage vibrations and cause a reduction in fuselage lateralvibrations.

The y--damper is adjusted, to give the lowest level of vibrations, at

101.5% NR instead of 100% NR. This is in order to achieve the bestcompromise of vibration levels when the rotor speed increases to104% NR at high density altitudes.


Y--Damper

Mass M 21

Tuning Sheets

Support

Y--Damper 2




z

x y

Mass M 22

Mass M 12

Mass M 11

Spring

Spring

Tuning Sheets

Y--Damper 1


Main Rotor System

General

The main rotor system consists of a bearingless, hingeless 4--blademain rotor, main rotor shaft with integral hub, control elements, and therotor-related indicators. By using modern composite materials, thisrotor system provides the flapping, lead-lag and blade pitch changefunctions without the installation of complicated ball and elastomericbearings. This type of construction is beneficial in terms of maintenance, cost and weight.

System Components

The components of the main rotor system are:

Swash Plate

The swashplate is the connecting link between the rotating rotor andthe stationary components of the control system. It is mounted to asliding sleeve, free to slide on a main gearbox mounted support tube.

Rotating Control Rods

The four rotating control rods transmit the control inputs from theswashplate to the main rotor blades. For flight control adjustment

(track and balance), the control rods are length-adjustable.

Driving Unit




p y

-- Four main rotor blades

-- Main rotor hub-shaft

-- Swash plate

-- Four rotating control rods

-- Driving unit

Main Rotor Blades

The four main rotor blades generate the lift and propulsion required forflight. Each blade is attached to the hub-shaft by two identical bolts.

Main Rotor Hub--Shaft

The main rotor hub-shaft transmits the driving torque from maintransmission to the main rotor blades. It also takes up rotor forces andmoments and passes them on to the main transmission.

Two scissors assemblies provide for synchronous rotation of theswashplate bearing ring with the rotor mast.


Main Rotor System

Main Rotor Blade

Hub-Cap Support

Hub-Cap




Swash Plate

Main Rotor Hub-ShaftRotating Control Rod

Scissors Assembly(Driving Unit)


Main Rotor Blade

General

The main rotor blade is manufactured from fiber composite materials. A blade root having low bending and torsional stiffness (Flex Beam)performs the functions of both the flap and lag hinges and the bladepitch bearings.

A pitch control cuff is integrated in the blade skin to provide a rigidconnection with the airfoil section of the blade. The pitch angle of themain rotor blade is changed through a pitch horn on the pitch controlcuff. During this feathering motion, the pitch control cuff is kept

centered about the blade root by a bearing support and a sphericalbearing.

Two elastomeric lead lag dampers provide sufficient in plane

Blade number 1 (yellow colour code) is the reference blade. Thesettings (pitch link length and trim tab position) must not be changed

during maintenance in order to store the basic rotor adjustment(min./max. pitch angle). All blades can be replaced individually due tothe manufacturer basic settings. The numbers and colour codes for theblades 2, 3 and 4 are mainly used as a reference for the track andbalance equipment.

NOTE If the basic adjustment is changed the relationship

between the rotor thrust and the collective pitch

lever position will be out of tolerance. Depending

on the amount of deviation the autorotation RPM

and the general H/C performance will be




Two elastomeric lead-lag dampers provide sufficient in--planedamping of the main rotor blade to prevent ground and air resonance.

The surface of the main rotor blade is provided with a protective coatof PUR lacquer to protect the composite materials from solar radiationand environmental and weather influences.

Color Marking

Each of the four main rotor blades is identified with a different color.The upper hub flange of the main rotor hub-shaft is coded with thenumbers 1 thru 4 on the blade attachment areas. In order to avoidhaving to readjust the control settings and the blade track whenremoving or installing the same main rotor blades, the main rotorblades are reinstalled so that their respective colors are pairedcorrectly with number codes on the hub flange.

and the general H/C performance will beinfluenced.

Color to Number Code Realationship

-- Yellow = number 1

-- Green = number 2-- Blue = number 3

-- Red = number 4


Main Rotor Blade

Metallic ErosionProtection

Transition Area Pitch

Airfoil Section




PU--ErosionProtection

Control Cuff

Damper Connectionwith Pitch Horn

Transition Area PitchControl Cuff to Airfoil


Blade Root

The blade root has the following functional areas:

-- Blade fitting area (1)Serves to attach the main rotor blade to the rotor hub of the main rotorshaft and is fitted for this purpose with two Teflon--coated bushings.

-- Soft flapping section (2)

This area enables the main rotor blade to flap up and down.

-- Soft torsion section (3)

Enables the main rotor blade to twist about its feathering axis tochange the blade pitch angle.

-- Soft lead-lag section (4)

The in--plane rigidity of the pitch control cuff is obtained through theunidirectional orientation of its carbon fibers in the trailing and leadingedge of the control cuff. Lead--lag rigidity is necessary to enable

lead-lag movements of the main rotor blade to be transmitted directlyto the lead-lag dampers without significant losses.

To prevent denting of the pitch control cuff -- especially on the lesscurved upper and lower surfaces -- it incorporates a sandwich structureand a hard foam filler core.

Two drain holes are provided on the underside of the pitch contol cuff at the outboard end adjacent to the blade airfoil section. These serve

to vent the pitch control cuff and to allow water which has condensedin or penetrated the pitch control cuff to drain off.

The integration (transition area) of the pitch control cuff into the bladebod pro ides a positi e and force transmitting connection hich




Enables in-plane motion of the main rotor blade.

Pitch Control Cuff

The pitch control cuff is provided with a transition area where it is

integrated with the aerodynamic portion of the blade, and with adamper connection at its open end. The pitch control cuff, whichpermits neither torsional nor lead-lag movements, surrounds the bladeroot and is rigidly connected to the adjacent airfoil section.

Torsional stiffness is required so that the control inputs can betransmitted through the pitch control cuff to the airfoil section of theblade.

body provides a positive and force transmitting connection whichtransmits the control inputs to the aerodynamic portion of the blade.Part of the forces and moments generated by the main rotor blade aretransmitted through this connection to the pitch control cuff.

A positive twist of +16° built into the blade in the region where the pitchcontrol cuff joins the airfoil section provides the airfoil section with acorresponding preset pitch angle and brings the flexbeam into anunloaded (untwisted) mid position.


Main Rotor Blade -- Control Cuff

Filler CoreInplane Stiffener

FlexbeamSandwich Construction

Inplane Stiffener




1

2

3

4

1 Blade Fitting Area2 Soft Flapping Section3 Soft Torsion Section

4 Soft Lead--lag Section

Control Cuff




Main Rotor Blade -- Blade Fitting Area and Pitch Control

Upper Lead-Lag Damper

Bearing Support

Spherical Bearing

Blade Bolt (2 off)




Special Nut

Safety Pin (2 off)

Pitch Control Cuff Seal

Lower Lead-Lag Damper

Pitch Horn

Control Cuff

Blade RootSleeve

Expansion Boltwith Rubber Cap


Airfoil Sectional Cut

Blade Core

The hard-foam blade core provides the supporting structure for theblade contour and stabilizes the blade skin.

Blade Spar

The blade spar consists of glassfiber rovings. They run from the bladetip to the blade root, around the bushings in the blade fitting area, andback to the tip. They absorb the tension and bending forces.

Lead RodThe lead rod in the blade leading edge determines the requiredposition of the blade center of gravity in the chordwise direction.




Blade Skin

The blade skin, which is made up of GRP plies, surrounds the spar,lead rod and blade core. It ensures that the aerodynamic portion of theblade is provided with the necessary torsional stiffness. The skin plieson the upper and lower surfaces of the blade converge at the bladetrailing edge where they are squeezed together to complete a torsionbox.


Main Rotor Blade -- Airfoil Section

Airfoil Section

Control Cuff withFlex Beam Section

Erosion Protection

Lead Rod

Spar




Spar

Blade CoreBlade Skin

Trailing Edge




Lead Lag Dampers and Bearing Support

The lead-lag dampers are attached to the damper connection on thepitch control cuff by screws installed through the bottom aluminum

plates. The top steel plates of the dampers are connected by nuts tothe ends of the bearing support, thereby connecting the lead-lagdampers to each other through the bearing support. Both lead-lagdampers are preloaded upon their connection to the bearing support.This prevents tension loading of the elastomer material during controlinputs and blade flapping movements. Tension loads would greatlyreduce the service life of the lead-lag dampers.

The lead-lag dampers are installed at a tilt in relation to the rotor planedue to the canted damper connection (see View V). This layoutenables a kinematic coupling to be obtained between the lead-lagmotion and the pitch angle of the main rotor blade. This pitch-lagcoupling effects a large part of blade lead-lag damping during flight

Operating principle of the lead lag damper and bearing supportassembly explained on the basis of its response to blade lagmovement:

-- The damper connection of the pitch control cuff makes alead movement in relation to the blade fitting,

-- the bottom aluminum plate is deflected forward, while thetop steel plate is restrained by the bearing support,

-- the layers of elastomers sandwiched between the steeldisks become deformed, absorb energy and, in doing so,dampen the lead motion of the main rotor blade.




coupling effects a large part of blade lead lag damping during flight.

The bearing support is mounted in blade fitting through a sphericalbearing which allows it to pivot and tilt. The bearing support togetherwith the lead-lag dampers support theopen end of thepitch control cuff

and center it around the blade root.


Pitch Control Cuff and Blade Root

V

A A

1 Lead--Lag Damper2 Control Cuff Seal3 Blade Fitting Area

4 Expansion Bolt with Cap5 Sperical Bearing6 Bearing Support7 Bottom Aluminum Plate8 Elastomer Layer9 Steel Disc10 Top Steel Plate11 Damper Connection on Pitch Control Cuff

12 Flexbeam13 Balance Washers14 Bolt for Bonding Jumper

7

89

11

10

14




Section A -- A

View V rotated 90, without Control Cuff Seal1

2

3

4

5

6

7 1112 13

14


Rotor Blade Adjustments

Manufacturer Adjustments

In the EC 135 main rotor all four blades can be replaced individually.On a rotor test stand the deviation of the dynamic behaviour from themaster blade is detected and corrected. In order to stay within themanufacturer limits the following parameters have to be adjusted.

Longitudinal Moment (Static Spanwise Balancing)

The longitudinal moment can be adjusted by changing weights in thecenter of the balance chamber which is exactly in the center of gravityline in the longitudinal axis. To determine the individual setting aspecial weighing equipment is necessary.

NOTE Any change of the longitudinal moment (e. g.application of paint in different radius stations of

After the measurements on the rotor test stand weights can be shiftedforward and backward in order to achieve the master blade track level.

The plastic spacers between the metallic weights allow a lateraltransfer of weight without influence on the longitudinal moment.

Pretrack Value

For the first rotor or blade adjustment the rotating pitch links normallyare set to a basic length. As a fine tuning towards the master blade thebasic length can be altered according the measurements on the rotortest stand. The pretrack value is a dimension in +/--[mm] for thechange

of the basic pitch link length and is stamped on the respective controlcuff and the rotor blade log card. Thus the necessary flight time for thetrack and balance adjustment can be reduced.

NOTE E ti t bl d l d




pp p

the rotor blade) will influence the blade behaviour

significantly and abnormal vibrations can occur.

Lateral Moment (Chordwise Balancing)The lateral moment determines the lift and therefore the track level of the rotor blade under different pitch angles. With the adjustment of thelateral moment the characteristic of the master blade can betransferred to all production blades.

By shifting mass behind the longitudinal center of gravity line theincrease of the lateral moment creates more lift with a higher track leveland vice versa. When leaving the production line the balance chambernormally is equipped with 12 weights (6 in front, 6 behind the centerof gravity line). To harmonise production tolerances brass or severalcombinations of brass and tungsten weights can be used.

NOTE Every time one or more rotor blades are replacedthe pretrack value has to be adjusted at first, even

for blade number 1 (yellow reference blade). For

any further track adjustment the pitch link length

or the trim tab setting of blade number 1 must notbe changed.


Plastic Spacer

Balance Chamber

Metallic Weightfor Lateral Moment




Metallic Weightfor LongitudinalMoment

Compression Spring

Blade Tip


Customer Adjustments

Track and Dynamical Balancing

Track adjustment of the main rotor blades is performed on the

helicopter by following means:

-- Adjusting the length of the rotating control rod.

-- Bending the trim taps.

Dynamic balancing of the main rotor is performed on the helicopter byadding or removing balance washers to or from the pitch control cuff.

Track Level Adjustment

The track level has to be measured in hover and in forward flight. Thebladenumber 1 has to be taken as reference blade and all other bladeshave to be brought into the deviation tolerance given in themaintenance manual.

NOTE Normally the track adjustment has to be done prior

to the balancing because a track change will againcreate an imbalance due to the moving center of

gravity of the rotor blade when flying higher or

lower. Modern track and balance computers are

able to combine both adjustments and to reduce

vertical vibrations by a certain track spread. Thistrack spread and all other adjustments have to

stay within the manufacturer limits given in the

maintenance manual.




The track level in hover flight is adjusted by changing the length of therotating pitch links (longerpitch link makes the blade fly higher andviceversa).

Further deviations to the blade number 1 in forward flight can becorrected by changing the trim tab setting (bending the trim tab upmakes the blade fly higher and vice versa).

Main Rotor Balancing

Dynamic balancing of the main rotor is performed by adding orremoving washers to or from the pitch control cuff. In order to eliminatean in plane imbalance weights can be found on one or two blades.


Rotor Blade Adjustments

Stabilizer (fixed setting)

Trim Tabs

Balance Washers




a a ce as e s

EC 135Training ManualFuselage

Fuselage





Table of Contents

General Description 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Cabin Structure 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Main Fuselage Structure 8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Doors 14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Service Covers 20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Windows 22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .





General Description

General

The fuselage serves as platform for the helicopter systems, crew,passengers and payload. The exterior shape of the fuselage isdictated by the major functions during operation and typical usage of light helicopters.

Components

The components of the fuselage are:

-- Cabin structure-- Main fuselage structure

-- Rear structure

-- Doors and service covers




-- Windows

Modular Concept

The modular concept simplifies the assembly of the helicopter andpermits the replacement of individual modules without the necessityof disassembling the entire fuselage.

Materials

The following materials are used:

-- Aluminium

-- Titanium-- Composit Materials (glass fiber and carbon fiber composite)

-- Acrylic glas


Fuselage

Main Fuselage Structure




Cabin Structure

Rear Structure


Cabin Structure

General

The cabin structure comprises the forward section above the cabinfloor. It is designed to function as a frame. It consists of:

-- Cabin framework

-- Cabin ceiling

-- Control post

Cabin Framework

The cabin framework is a one-piece structural component. It isconstructed as a hollow profile made of composite material (mainlycarbon-fiber). The framework provides the structural support formounting the windshields, the nose windows, the pilot/copilot doorsand the sliding doors to the passenger compartment. The upper forkend of the windshield center post houses the overhead panel

Control Post

The control post is installed between the cabin floor and the cabin roof.It only houses the vertical control rods for main rotor control. Thecontrol post is made of aluminum alloy. It is displaced slightly to thestarbord side of the helicopter to allow the pilot to have anunobstructed view to the rear left.

NOTE The control post is a non load carrying structure. It

houses the control rods only.




end of the windshield center post houses the overhead panel.

Threaded inserts in the area of the window frame profiles are providedfor installation of the front and nose windows.

Cabin Roof

The cabin roof covers the cabin framework. It also functions as a fairingfor the main rotor control rod system.

The cabin roof is made of carbon fiber composite material (partlysandwich). The roof is riveted to the cabin framework. To allow accessto the control rods and an upper guidance unit, a handhole is provided

in the upper right side of the cabin roof dome.

NOTE The cabin roof is a non load carrying structure.

NO STEP!


Cabin Structure

FWD

Center Post

Handhole

Carbon Fiber




FWD

Carbon Fibre

Glass Fiber Plies in theLower Section

Cabin Roof



General

The main fuselage structure is the part of the fuselage that carries all

the loads transmitted by the main transmission from the main rotorsystem and all the loads caused by the engines, landing gear and tailunit.

Components

The main fuselage structure consists of the following:

-- Body structure

-- Floor structure

The body and floor structure are ridgidly attached to each other.

Body Structure

The predominantly aluminum alloy body structure is composed of

Side Panels

The side panels, which provide the framework on the sides of the body

structure, consist of frames 4 thru 7 and stringers. The outer skin,which is aluminum alloy, is riveted to the frames and stringers.

Integrated in the side panels are maintenance steps. The left-handside panel also incorporates a housing for accomodating the fuel fillerneck.

The outer skin of each side panel is provided with cutouts for the aftwindow panes and the cooling vents.

Attached to the outside of both side panels is a center door rail forguiding the respective sliding door.

Transmission Deck

The transmission deck which takes up the load of the lifting system




The predominantly aluminum-alloy body structure is composed of individual assemblies which are:

-- Side panels (2 off)

-- Transmission deck

-- Engine deck

-- Rear structure attachment cone

-- Equipment deck

The body structural components are rigidly attached to each other.

The transmission deck, which takes up the load of the lifting system,consists of frames 4 thru 5 and longitudinal beams. It is attached byrivets to the side panels. On the transmission deck six mounts for main

transmission installation are provided. The transmission deck skin isaluminum alloy.



5 6 74a4 85a

Equipment Deck

Transmission Deck

Rear Structure Attachment Cone




FWD

4 8 Frame 4 to Frame 8

RH Side Panel

LH Side Panel

Engine Deck




Airframe StructureRear Structure Attachment ConeEngine Deck

Transmission Deck

Transmission Mounts

RH Side Panel

Floor Structure

Equipment Deck

Frame 8

Frame 7




LH Side Panel

Frame 6

Frame 5

Frame 4a

Frame 4Frame 3

Landing Gear Fitting

Frame 2Frame 1


Cabin Floor

The cabin floor supports the seats and parts of the interior furnishingsof the helicopter. It is an aluminum honeycomb sandwich construction

and comprises the following sections:-- Foreward floor

-- Aft floor

-- Left and right cable channel cover

Located in the forward floor are cutouts through which the flight controlelements and wiring harnesses are routed. The forward floor providesthe points of attachment for the pilot seats, controls and consoles. The

bottom end of the control post is also bolted to the forward floor.

Integrated into the removable aft floor are tracks running in alongitudinal direction. These enable the helicopter to be configuredwith passenger seats or items of special operational equipment.

The removable side channel covers cover the area of the floor

The fuel tanks are located between frames 3 and 5 and behind frame5, respectively.

Lower Shell

The lower shell, which is a one-piece composite structure, enclosesthe subfloor structure and supports the fuel tanks. It is riveted to thesubfloor structure.

A maintenance hole is provided in the lower shell between frames 1and 2 and between 2 and 3, respectively.

Running laterally below each frame 2 and 5 is a tunnel which isoccupied by a landing gear crosstube.

In the area behind frame 3 and in front of and behind frame 5, the lowershell is stiffened to provide a firm mounting base for the fuel pumps.

A lower door rail for guiding the respective sliding door is integrated inthe upper edge of each side of the lower shell between frames 2and 4




between the forward and aft floors and the cabin side shell.

Subfloor StructureThe subfloor structure, which is a aluminum-alloy construction,supports the cabin floor and the landing gear. It is made up of frames1 thru 6 and two longitudinal beams. The structure is riveted to the sidepanels through the frame and the lower shell.

Disposed between the longitudinal beams behind frame 1 and in frontof frame 2 is a transverse bridge.

A forward and an aft landing gear fitting are riveted to each of the twolongitudinal beams.

and 4.


Floor Structure

Cabin Floor

Subfloor Structure

1

23

4

56

4a




Lower Shell

FWD


Doors

General

The helicopter fuselage is fitted with six entrance doors to provide

access to the cockpit, passenger cabin and cargo compartment.

Cockpit Doors

The cockpit doors (pilot doors) are hinged doors located left and rightat the foreward part of the cabin frame. In the standard version theycan not be jettisoned.

The cockpit doors are a carbon-fiber composite construction with a

seal fitted to their circumference. They are installed to the cabinframework via two hinges with integral bearings and two clevis fittings.The upper one is attached by rivets and the lower one by screws.

The rear edge of the pilot door is fitted with locking devices at the topand at the bottom. They are operated through the exterior or interiordoor handle and the interconnecting lever and tubes The claws of the

Cockpit Door Windows

The pilot door windows are made of 3--mm--thick acrylic glass. They

are positioned on a layer of adhesive sealant in the door structure andsecured to the latter by countersunk screws and dimpled washers.

The pilot door windows incorporate smaller sliding windows which aremoved on rails by means of a handgrip bonded to the pane. The slidingwindows are held by friction in the selected open position on the rails. A mechanical detent locks them in the closed position so that theycannot be opened from the outside.




door handle and the interconnecting lever and tubes. The claws of thelocking devices engage with the mating fittings on the cabinframework. The pilot door can be locked with an integral door lock. Agas spring holds the unlatched pilot door wide open.

In a second version the gas spring is removed and the door can belocked in the full open position in the vicinity of the pitot tubes.


Pilot Door

Locking Device Top

HingesDoor Handle

GasSpring

Handle for Locking Device




Locking Device

Spring


Sliding Doors

The sliding door is a carbon-fiber composite construction. It is fittedwith a door seal around its entire circumference except for the edge

adjacent to the pilot door. Fitted to the forward top and bottom cornersof the sliding door are the upper arm and lower guide which areprovided with a runner and a roller, respectively. The sliding door ismoved on its upper arm andlower guide along an upper railin thecabinframework and a lower rail in the lower shell. Another arm with anintegral runner is fitted on the rear edge of the sliding door. By meansof this arm, the sliding door also runs on a rail located aft in the sidepanel.

The sliding door is opened and closed via the exterior door handle orinterior door handle, and the associated locking mechanism. Latchingof the sliding door is provided by an inner tube which matches with afitting in the cabin framework above the sliding door, and by a lockwhich matches aft with a corresponding fitting in the side panel.

For flight with open sliding door the locking mechanism for the open

Emergency Exit

The clamping seal of the sliding door window is formed with four slits.Of these, the two lateral inner and outer slits are each fitted with a filler

(PVC cord with matching profile) which expands the circumference of the clamping seal so that the window is held firmly in the door frame.The filler in the inner or outer lateral slit can be pulled out of theclamping seal by means of an emergency handle on the inside andoutside of the the sliding door. To prevent of an inadvertant pulling, theemergency handles are protected by pushbutton-fixed covers. Afterthe filler has been removed, the window pane can be pressed out of the sliding door.




For flight with open sliding door the locking mechanism for the openposition has to be installed and the speed limits have to be obeyed.

Sliding Door Windows

The sliding door windows are made of 3--mm--thick acrylic glass. Theyare fitted in the sliding doors with a peripheral clamping seal whichenables them to be removed quickly to provide a mean of escape inthe event of an emergency.


Sliding Door

Upper Armwith Runner

Guard Cover

EmergencyLoop Strap

Clamping Seal

Filler

Sliding Door Pane




Lower Guidewith Roller

Aft Arm withRunner


Rear Doors

Therear door structure is a carbon fiber/glass fiber hybrid construction.The edges of the rear doors are fitted with a door seal. Attached byscrews to each rear door are two fittings through which the rear doorsare connected to the main fuselage structure. Attached by screws tothe inside of each rear door is a fitting to which is installed a gas springfor holding open the unlatched rear doors. Two locking mechanismsare installed on the edge of the right--hand door which, when the doorsare closed, clasp the mating sleeves on the edge of the left--hand door.Both rear doors are latched together from the outside and then lockedwith a key.

Rear Door Windows

The rear door panes are made of 2 mm thick acrylic glass. They arebonded to the rear door structure and are secured by screws.





Rear Doors

Gas Spring

Rear Door




Door Fitting

Locking Mechanism


Service Covers

General

Installed on the fuselage are a number of service covers which can beremoved to get access to components inside the helicopter.

Handhole Cover

The handhole cover, which is constructed of GRP, has a seal bondedto its inside edges. It is attached by screws to the cabin roof cowlingand when removed provides access to the upper main rotor controllinkage.

Nose Cover

The nose cover, which is of fiberglass honeycomb panel construction,has a seal bonded to its inside edges. Installed in the nose cover is afixed position landing light. The nose cover is attached to the cabinframework by stud fasteners. Removal of the nose cover providesaccess to the landing light instrument connections components of the

Middle Cover

The middle cover is of aluminum sheet metal construction. It isattached to the lower shell by means of stud fasteners. Removal of themiddle cover provides access to flight control components and to the

engine emergency control connections.

For helicopters equipped with a cargo hook the middle cover is fittedwith a hood. A cover is attached to the hood to provide access tocomponents of the cargo hook.

Tank Covers

The forward main tank cover and the aft main tank cover are

constructed of aluminum sheet metal. They are provided with aprotective plastic edging. Each cover has a round opening in which theboot of the associated fuel drain valve is inserted. The covers areattached by screws to the lower shell. Removal of the covers providesaccess to the equipment plates of the fuel system.

The supply tank cover is constructed of aluminum sheet metal. It hast d h l i hi h th b t f th f l d i l i t d




access to the landing light, instrument connections, components of thecabin heating and ventilation system, and the windscreen wiper motor.

Foreward Access Cover

The forward access cover is a fiberglass honeycomb panelconstruction which is attached to the lower shell by stud fasteners.When the stud fasteners are opened, the nose cover hangs from thelower shell by means of four cables with snap hooks on their endswhich clip onto brackets on the forward access cover and the lowershell. Removal of the forward access cover provides access to flight

control components and to the blower of the cabin heating andventilation system.

two round holes in which the boots of the fuel drain valves are inserted.The cover is attached by screws to the floor shell. Removal of thecover

provides access to the two equipment plates of the fuel system.

Rear Structure Covers

The RH and LH tail boom covers are of composite construction. Theyare attached by screws to the tail boom. Removal of the coversprovides access to the antenna connections, wiring harnesses and theflux valve.

The lower and aft vertical fin covers are of composite construction.They are attached by screws to the Fenestron structure. Removal of the covers provides access to the inside of the Fenestron structure forinspection purposes.


Service Covers

Handhole Cover

Vertical Fin Covers

Tail Boom Cover LH

Tail Boom Cover RH




Nose Cover

FWD Access Cover

Middle Cover(Standard)

Middle Cover (alternativewith Cago Hook)

Tank Covers


Windows

Windshields

The windshields are made of 5 mm thick acrylic glass. Optional

windshields with a hard, scratch--resistant surface coating are alsoprovided. The windshields are positioned on a formed sealing strip anda layer of adhesive sealant in the cabin framework and secured to thelatter by countersunk screws, dimpled washers and sealing washers.The bottom edge of the windshields is not attached by screws to thecabin framework, but is held against it by a metal retaining strip . Ametal strip is installed between the windshild, which is attached byscrews to the center post of the cabin framework. It is installed flush

with the adjacent windshields to provide a flat, continuous surface forthe windshield wiper. The joint between the windshields and the cabinframework is not rigid but designed to give the windshields a limiteddegree of movement relative to the cabin framework. In consequence:

-- Varying degrees of heat expansion in the cabin frameworkand the windshields are compensated for and

Nose Windows

The nose windows are made of 2--mm--thick acrylic glass and

reinforced with 1 mm thick Orlon around the edges. They arepositioned on a formed sealing strip and a layer of adhesive sealantin the cabin framework and secured to the latter by countersunkscrews and dimpled washers. The upper edge of the nose windows isnot attached byscrews to the nosespar, but is heldagainst itby a metalretaining strip which itself is attached by screws to the nose spar.

Side Windows

The side windows are made of 2 mm thick acrylic glass. They arepositioned on a layer of adhesive sealant in the side panels andsecured to the latter by round-head screws and washers.

Cleaning of the Windows

NOTE Use only approved cleaning agents Unapproved




-- Stresses imposed on the windshields due to deformation of

the cabin framework are prevented.For this purpose, the diameter of the washer holes is greater than theshank diameter of the mating countersunk screws.

NOTE Use only approved cleaning agents. Unapproved

cleaning agents may contain harmful solvents that

could cause crazing.


Windshield, Nose and Side Windows

Metal Strip

LH Windshield




LH Side Window

LH Nose Window

Metal Strip

EC 135Training ManualTail Unit

Tail Unit





Table of Contents

Principle of the Fenestron 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Unit 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Horizontal Stabilizer with End Plates 8. . . . . . . . . . . . . . . . . . . .

Tail Boom 10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor Drive 12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Vertical Fin with Fenestron 16. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor 18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor Gearbox 24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor -- Inspection 30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .




Principle of the Fenestron

Direction of Air Flow

Thrust of Tail Rotor

Torque Main Rotor




Sense of Rotation Main Rotor

FWD


Tail Unit

General

The rear structure is the aft section of the fuselage. It stabilizes thehelicopter in flight by means of the vertical fin with the integratedFenestron tail rotor and provides the lever arm on which the thrust of the tail rotor counteracts the torque of the main rotor system. The rearstructure is mainly constructed of composite materials.

Components

The rear structure of the EC 135 consists of the following assemblies:

-- Tail boom

-- Horizontal stabilizer with end plates

-- Vertical fin with Fenestron structure




Horizontal Stabilizer with End Plates

General

The horizontal stabilizer dampens pitching motions of the helicopter

around the lateral axis during forward flight. The horizontal stabilizerhas an asymmetric profile which is curved on the underside and isequipped with a spoiler on both sides. The pitch angle is a permanentfactory setting. The horizontal stabilizer aerodynamically stabilizes thepitch attitude of the helicopter in cruise flight and dampens pitchmotion.

When viewed in the direction of flight, the end plates are permanentlyoffset to the right, thereby enabling them to reduce aerodynamicallythe thrust power required of the tail rotor system in cruise flight.

Design

The horizontal stabilizer passes through the tail boom. Above andbelow the cutout on each side of the tail boom is an attachmentbracketthrough which a single bolt is installed to secure the horizontalstabilizer to both sides of the tail boom.




The horizontal stabilizer is a shell-type structure made of carbonfiber-reinforced plastics.

Attached by 6 screws to each outboard end of the horizontal stabilizeris an end plate which is a honeycomb sandwich construction.

Fitted to the outboard sides of the end plates are the navigation lights.For easy removal/installation the two parts of the spoiler are bolted onthe R/S side while riveted only on the L/H side.


Horizontal Stabilizer and End Plates

Horizontal StabilizerBolt

Spoiler(Riveted)

Spoiler(Bolted)




End Plate

Nut


Tail Boom

General

The tail boom connects the rear structure to the main fuselage

structure. It supports the vertical fin, tail rotor systems and thehorizontal stabilizer. Running along the top of the tail boom are the tailrotor drive shaft, hydraulic lines and the tail rotor flex ball control.

Design

The tail boom is a sandwich structure consisting of a Nomex core withcarbon fiber-reinforced facings in which is embedded copper foil toensure electrical conductivity.

The conically-shaped tail boom is built up of two half sections joinedtogether by bonding and additionally secured by rivets. Thealuminum-alloy connecting frame is riveted to the inside of the tailboom. To prevent corrosion, the mating surfaces are isolated fromeach other by layers of sealing compound. The tail boom is bolted tothe connecting frame 8 of the main fuselage structure through itsconnecting frame.

Access to the interior of the tail boom is provided by maintenancecovers. Routed inside the tail boom are cable ducts for the electricalcables.

When communication/navigation systems such as the VHF, VOR, ADF, and radar altimeter (optional equipment) are installed, the tailboom is fitted with antenna connections to which the respectiveantennas are installed.

Fairing

A detachable fairing made of fiber-reinforced plastic provides a

covering for the tail rotor drive shaft, hydraulic lines, and the ballbearing control. The fairing is fitted by spring-loaded fasteners to thetail boom.

On the connecting frame, a bulkhead plate is attached.




Fittings

In the areas where the fittings are installed, the half sections are locallyreinforced. The aft end of the tail boom is provided with two cutoutswith integral fittings for attaching the horizontal stabilizer. Bolted atintervals along the top of the tail boom are five bearing supports forsupporting the tail rotor drive shaft. The first three brackets aresupported by vertical struts in the structure in order to stabilize the

entire system.


Tail Boom

Fairing (Carbon Fiber)

Support Fitting

Bearing Support(Aluminium)

Fitting for Horizontal

Connecting Flange

LongDrive Shaft

ForewardShort DriveShaft

Hydraulic Hoses

FWD




Maintenance CoverCable Duct

BulkheadPlate

Connecting Frame(Aluminium)

Fitting for Horizontal

Stabilizer

Tail Boom (Nomex Sandwich)

Tail Boom

Antenna Attachment

FWD U--Profile

Vertical Strut


Tail Rotor Drive

General

The tail rotor drive transmits the power from the main rotor

transmission to the tailrotor through a system of shafts, flexiblecouplings and the tail rotor gearbox.

Components

The tail rotor drive train consists of the following parts:

-- 3 shafts with flexible couplings

-- Tail rotor gearbox

Drive Shafts

The tail rotor drive shaft assembly consists of:

-- Foreward drive shaft with two couplings

-- Center drive shaft with 6 bearings

-- Aft drive shaft with two couplings





Tail Rotor Drive Shaft

Foreward Drive Shaft

Center Drive Shaft

Bolted Flange

Bolted Flange




Flexible Coupling

Gearbox InputShaft

Aft Driveshaft


Foreward-- and Aft Drive Shaft

The foreward and aft drive shafts are built up as follows:

-- Tube

-- Adapers-- Flexible Couplings

The tubes consist of carbon fiber. The three-armed adapters consistof titanium and are riveted and bonded to the ends of the tubes.

The foreward drive shaft is connected via the flexible couplings andflanged couplings to the tail rotor output drive of the main transmissionand to the center drive shaft.

The aft drive shaft is connected via flexible couplings directly to thecenter drive shaft and to the tail gearbox input flange.

Flexible Coupling

The flexible couplings consist of packs of steel discs which are heldtogether by assembled flanged sleeves and washers. The flexiblecouplings correct for misalignment and variations in length.

Center Drive Shaft

The center drive shaft is built up as follows:

-- Tube

-- Two removeable flanges-- 6 roller bearings with rubber sleeves

The tube consists of steel. The bolted and the removable flangesconsist of titanium.

The removable flanges are connected to the tube by spring bushingswhich are secured by bolts, nuts and special washers.

The center drive shaft is supported by 6 sealed roller bearings, whichare mounted on top of the tail boom by bearing supports. The innerraces of the bearings are embedded in rubber sleeves, which help todampen vibrations, and account for misalignment.





Drive Shafts -- Tail Rotor

Rubber Sleeve

Ball Bearing

Center DriveShaft

Aft Drive

Flange

ForewardDrive Shaft

FlexibleCoupling

Spring Bushing

Rivets

Adapter




Flange

Flexible Coupling

t eShaft

Bolt

Special Washer


Vertical Fin with Fenestron

General

The vertical fin together with the integral Fenestron structure form a

unit. The upper region of the vertical fin has an aerodynamic function,while the Fenestron structure below it encloses the tail rotor system.

The yaw control of the helicopter is made possible by the Fenestron.

Design

The vertical fin is constructed of Nomex honeycomb with carbonfiber-reinforced facings. Embedded in the outer facing plies is a copperfoil which ensures electrical conductivity. The vertical fin is built up of two half sections joined together by bonding and additionally securedwith rivets. It is riveted to the tail boom via a connecting flange.

A fin tip fairing, which incorporates the anti-collision light, is screwedto the open upper end of the vertical fin.

Screwed to the underside of the Fenestron airframe is a tail bumperwhich increases the yaw stability and protects the tail boom againstimpacts, e.g. ground contact during flare. A static discharger is fitted




at the fin tip fairing as well as at the tail bumber.


Vertical Fin with Fenestron

Fin Tip Fairing

Vertical Fin

Support Fitting

Guide Vane

Half Fairing

Static Discharger

Gearbox Cover




Tail Bumper

Stator Hub

Static Discharger




Principle of Tail Rotor

S f R t ti

FWD




Sense of RotationTail Rotor

Yaw Control


Components

The tail rotor consists of the following:

-- 10 tail rotor blades

-- Hub body-- 10 inner bearings

-- 10 outer bearings

-- Pitch change spider

-- Center flange

-- Fairing

Tail Rotor Blades

The tail rotor blades are constructed of aluminum alloy and consist of the blade air foil and the blade root. The tail rotor blade air foil isformedwith a built-in spanwise twist. It has a nonlinear profile whichprogressively changes from the blade neck to the root twist. The bladeroot is hollow. It has two bearing surfaces and, a bore for receiving twobushings and the blade bolt, and a pitch horn. The tail rotor blades aresupported in the hub body by the mating sliding bearings. This

t bl th t il t bl d t f th d h th i

Fairing

A fairing protects the components within the hub body and is fitted withfasteners and plate nuts. At the center of the fairing is a bore which isused to detach the fairing. The bore is sealed by a plug.




arrangement enables the tail rotor blades to feather and change theirpitch angles. Bolted to the pitch horn is a ball segment which connectsthe tail rotor blade to the pitch change spider. The hollow blade rootserves to accomodate the tension-torsion bar to which the rotor bladeis attached by bushings and a blade bolt.


Tail Rotor

Fairing

Control Rod

Guide

Output Gear Wheel

Center Flange

Pitch Change Spider

Tail Rotor Blade

Thrust Nut

Locking Washer

Outer Bearing

Inner Bearing




Fairing


Thrust Nut

The thrust nut is screwed to output gear wheel of the tail rotor gearboxand securs the tail rotor. It is prevented from rotating by the lockingwasher. The thrust nut transmits the tail rotor thrust to the Fenestron

structure through the tail rotor gearbox and the stator.

Pitch Change Spider

The pitch change spider is attached to the pitch horn of the tail rotorblades through ball joints. It is the central pitch changing componentsfor all of the tail rotor blades.

Center Flange

The center flange is bolted to pitch change spider and is connected tothe control rod and guide of the tail rotor gearbox. Interposed betweenthe guide in the tail rotor gear box and the center flange is an settingshim by means of which the pitch of the tail rotor blades can be set.

Control inputs move the control rod and the guide, which in turn movethe pitch change spider axially through the interconnected centerflange. Simultaneously, the pitch angle of all the blades is changed by

the same amount via the pitch horns mounted on the pitch change

Hub Body with Bearings

The hub body houses the tail rotor components. In the hub body, thetail rotor blades are each supported in an outer and an inner bearing.On the hub body rear side 6 threads for bolts and balance washers are

installed.

NOTE For balancing work the bolts have to be numbered

from 1 to 6 beginning at the speed reference mark

in counter--clockwise direction.

Splined Hub Flange

The splined hub flange is connected to the hub body by screws and,through its internal spline, is splined to the pinion of the output gearwheel. It connects the tail rotor to the tail rotor gear box.

Tension--torsion Bar

The tension-torsion bar consists of a stack of steellaminates which areheld together bay a shrink sleeve. The tension-torsion bars retain thetail rotor blades within the hub body and connect them to the hub

flange The tension torsion bar absorb centrifugal forces The low




the same amount via the pitch horns mounted on the pitch changespider.

Chinese Weights

The Chinese Weights or propeller moment weights dynamicallyreduce the control forces.

NOTE There are different chinese weights mounted to the

left and to the right.

flange. The tension-torsion bar absorb centrifugal forces. The lowtorsional stiffnes of its steel laminates enables pitch angle variation onall the tail rotor blades.

Attach Ring

The attach ring together with the tension-torsion bars and the hubflange are attached to the hub body by bolts and associated nuts.


Tail Rotor Control

1

1

11

11

2

3

8

9

9

8

46

3

View from side of the tail rotor fairing

7

10

Bolt/Washer for Balancing

Thread for Bolt/Washer(6 Positions)

12




1 Hub Body2 Splined Flange3 Attach Ring4 Pitch Change Spider

5 Bushing6 Tension--torsion Bar

7 Bushing with Chinese Weight8 Inner Bearing Ring9 Outer Bearing Ring10 Ball Joint

11 Tail Rotor Blade12 Plate

45

67

10

10


Tail Rotor Gearbox

General

The tail rotor gearbox is a single-stage, spiral-toothed bevel gear. Itdoes the following:

-- Drives the tail rotor

-- Reduces the speed from the drive shafts

-- Diverts the direction of power flow through 90° by means of two bevel gears

-- Transmits tail rotor forces and moments through the statorto the fuselage

The tail rotor gearbox houses the components which control the tailrotor. These components transmit the control inputs from nonrotatingto the rotating parts of the tail rotor.

Components

The tail rotor gearbox consists of the following:

-- Gearbox housing

-- Input casing

Design / Function

The gearbox housing is made of aluminum alloy. Installed in thehousing are the input pinion gear and output gear wheel which areattached by the flanges of their supporting bearing outer races to thegearbox housing. The gearbox housing is provided with an inputcasing and an output drive casing which are both fitted with a shaftseal.




Input casing-- Output casing

-- Input drive flange

-- Input pinion gear

-- Output gear wheel

-- Control unit (comprising casing, control rod, guide)


Tail Rotor Gearbox

Input Casing

Input Pinion Gear

Gearbox Housing

Control Unit

Output Gear Wheel

Output Casing

Shim

Guide

Sight Glass




Input Drive Flange

Lip Seal

Input Casing

Drain Plug

Electrical Chip Detector


Input Drive Flange

The input drive flange which transmits torque to the input pinion gear,is formed with a three-arm flange and a splined shaft which mesheswith the internal spline of the input pinion gear.

Input Pinion Gear

The input pinion gear, which drives the output gear wheel, consists of a spiral bevel gear, a double ball bearing, and a special nut securedby a locking ring.

Output Gear Wheel

The output gear wheel, which drives the tail rotor, consists of a spiral

pinion gear, a double ball bearing, and a special nut secured by alocking ring. The tail rotor is splined to the pinion of the output gearwheel through the splined hub flange.

Control Unit

The casing, control rod and guide together comprise the control unitwhich is installed inside the output gear wheel. Control inputs causethe Fenestron actuator to move the contol unit in an axial direction. The

control unit transfers control movements to the tail rotor.The control unit casing comprises the casing itself and an integratedcontrol rod which is connected to the input lever of the tail rotor controllinkage so that the casing cannot rotate.

Installed inside the casing is a control rod and a double ball bearingwhich is held in the housing by a special nut and secured by a nutretainer.

The components inside the casing provide for the transition fromnonrotating to rotating movement of the tail rotor controls.

The axial movement of the control unit casing is transferred throughthe double ball bearing to the pivoted control rod and guide. Thecontrol rod and guide are connected to the tail rotor blades through thecenter flange and the pitch change spider of the tail rotor, causing themto rotate at the same speed as the tail rotor.

An setting shim is interposed between the guide and the central flange.




g p g gThe thickness of the setting shim determines the position of the centralflange and, when adjusted, affects the pitch of the tail rotor blades.


Tail Rotor Gearbox

1

2

3

4

56

7

1 Input Drive Flange2 Input Casing

3 Input Pinion Gear4 Output Gear Wheel5 Guide6 Control Rod7 Casing of Control Rod8 Output Casing9 Gearbox Housing10 Double Ball Bearing

11 Nut12 Nut Retainer13 Setting Shims14 Pitch Change Spider

5610 1211 13 147

FWD




8

9

Control Unit


Oil System

Installed in the lower region of the gearbox housing is a valveincorporating a magnetic plug which is fitted with an electrical chipdetector. The magnetic plug is retained within the valve by a bayonet

coupling. When the magnetic plug is removed, the valve closesautomatically to prevent oil from flowing out.

The oil in the tail rotor gearbox is drained by means of a hose with anadapter which fits into the valve. An oil level sight glass, which hasminimum and maximum markings, enables visual inspection of the oillevel.

The oil filler neck of the gearbox housing is fitted with a strainer and a

cap.The gear wheels and bearings of the tail rotor gearbox are providedwith splash lubrication.

The tail rotor gearbox is cooled by the circulating oil and via thegearbox housing.

Balancing Installation

For balancing the tail rotor a velocimeter and a magnetic speed pickupare installed at the tail rotor gearbox The wiring leads to a receptacle




are installed at the tail rotor gearbox. The wiring leads to a receptaclein the circuit breaker panel 1 which is situated in a recess in the LHcargo compartment side cover.

NOTE A dummy velocimeter pickup or a operating pickupcan be installed.


Tail Rotor Gearbox

Strainer

Cap

Adapter forOil Hose

Oil Level Sight GlassMIN and MAX Markings

Input Casing

Input Drive Flange

Gearbox Housing

Housing forSpeed Sensor

Magnetic Pickup

Revolution Marker

V l i tCircuit Breaker Panel 1




DCRECEPT

TR&BALINFLT

3MJA

19VVA

10 5

Magnetic Plug

Electrical Plug

Velocimeter

Velocimeter

Receptacle forTrack&Balance

Circuit Breaker Panel 1


Tail Rotor -- Inspection

Clearance Check of the Tail Rotor Blades

The clearance at any position of the tail rotor blades and the Fenestron

structure must not be less than 3.5 mm.

Procedure

The clearance of all rotor blades is measured with a gauge atposition 1 (lowest part of the Fenestron duct). The blade with theminimum clearence is rotated with 45° steps and the clearance ismeasured at each position.

CorrectureCheck whether paint was applied too thickly in the affected area of theFenestron structure when the paint coat was previously renewed ortouched up. If this is found to be the case, reduce the thickness of thepaint coat by the excessive amount. However, the paint must not beremoved to the point where the light blue primer coat is exposed.

If this does not apply, disassemble the tail rotor and inspect screws

and laminated tension-torsion bars for wear. Replace worn parts andreassemble the tail rotor. After the tail rotor has been reassembled,




measure again the clearance between blade tips and the Fenestronstructure.


Tail Rotor -- Clearance Check

3

4

5

6

7

82

Tail Rotor Blade

1 8 positions at which clearance is measured




1Fenestron Structure

Allowable Clearance 3.5 mm

EC 135Training ManualFlight Control

Flight Control




EC 135

Training ManualFlight Control

Table of Contents

Principle of Flight Control 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Flight Control of the EC 135 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Collective Control 8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Cyclic Control 12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Mixing Lever Assembly 14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Swash Plate 18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Rotating Control Rod 20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Driving Unit 22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Track&Balance Installation 24. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Trim System 26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tail Rotor Control 32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Hydraulic System 38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Pressure Supply Systems 42. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Hydraulic Actuators 52. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Description of the Follow Up Principle 52. . . . . . . . . . . . . . . . . . .

System Description MHA/EHA 56. . . . . . . . . . . . . . . . . . . . . . . . . .

Mechanical Override 60. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Electro-- Hydraulic Actuator EHA 64. . . . . . . . . . . . . . . . . . . . . . .Indication and Testing System 68. . . . . . . . . . . . . . . . . . . . . . . . . .




Fenestron Actuator 70. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Three Axis Stability Augmentation System SAS 72. . . . . . . . . .

Yaw Stability Augmentation System 72. . . . . . . . . . . . . . . . . . . . .

Pitch & Roll Stability Augmentation System 76. . . . . . . . . . . . . .

Pitch Damper (DPIFR) 80. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EC 135






EC 135


Principle of Flight Control

General

The attitude and airspeed of the EC 135 is controlled by adjusting the

angle of incidence of the main and tail rotor blades.

Flight Control

Three types of controls are necessary to fly the helicopter:

-- Collective control of the main rotor

-- Cyclic control of the main rotor

-- Tail rotor control

The pilot gives control signals by:

-- Collective pitch lever (left hand)

-- Cyclic control stick (right hand)

-- Tail rotor pedals (feet)

Collective Control

Changing the angle of incidence equally on all four main rotor bladesincreases or decreases the main rotor thrust. This is called collective

l

Tail Rotor Control

The tail rotor control is in principle the same as the collective control

of the main rotor system. Adjusting the angle of incidence of the tentail rotor blades collectively varies the thrust, reacting against the mainrotor torque. The helicopter stands still in hover, if these forces areequal. If not, the helicopter will turn around its yaw axis.




control.

Cyclic Control

The cyclic control adjusts theangle of incidence of two opposite bladesperiodically and inverse. By means of this results a horizontal force.The helicopter will tilt and move in the direction of the horizontal force.

Cyclic control consists of lateral control (left and right movement) andlongitudinal control (forward and backward movement).

EC 135


Flight Control

Cyclic ControlMain Rotor

Collective ControlMain Rotor

Yaw ControlTail Rotor

FWD

FWD




EC 135


Flight Control of the EC 135

Components

The flight control of the EC 135 comprises the following systems:

-- Main rotor control

-- Tail rotor control

Main Rotor Control

The main rotor control consists of two systems:

-- Collective control

-- Cyclic control

Components

The most important components of the main rotor control are:

-- Collective lever

-- Cyclic stick grip

-- Trim system

-- Control linkage, non boosted section-- One mechano-hydraulic actuators (MHA)

Tail Rotor Control

The main components of the tail rotor control are the following:

-- Pedal assembly

-- Ball bearing control cable

-- Electro-mechanical actuator (SEMA)

-- Fenestron actuator




-- Two electro-hydraulic actuators

-- Mixing lever gear unit

-- Control rods, boosted section



EC 135


Collective Control

Signal Input

The collective signals are given by pulling the collective pitch leverupward or pushing downward. Pulling creates climb, pushing descent.

Collective Pitch Lever

The collective pitch lever is located on the left side of the pilot seat. Thesecond lever is located on the left side of the copilot seat. Bothcollective pitch levers are mechanically linked via a torsion tube.

Friction Brake

To prevent undesired movement of the collective lever during flight, afriction brake acts on the torsion tube. The desired friction against themovement of the pitch lever can be set by the adjusting screw.

Collective Pitch Stop

The collective pitch stop is an elastic stop which limits the angle of attack of the main rotor blades in fast, high density altitude flights.

During an emergency condition i.e. autorotation landing it may be

necessary to exceed this elastic stop. This will increase the collectivecontrol force because of a spring force to overcome.

NOTE Final adjustment of the collective pitch stop is

determined during maintenance check flight. Theactual mechanical stop is compared to the rotor

thrust given by the measured torque under therespective outside air conditions (PA, OAT). If there

is a difference to the calculated volume in the

diagram, the mechanical stop can be adjusted bychanging the number of shims under the flange.



EC 135


Control Transmission

The signals are transmitted via a torsion tube, located underneath thecockpit floor, several control rods and bell cranks to the input controllever of the dual hydraulic boost unit. Here the signals are force

amplified. The amplified signals are transmitted via a control rod to thecollective control fork, which is part of the mixing lever assembly.

The collective control fork lowers or lifts the sliding sleeve, whichcreates the intendet simultaneous variation of the angle of incidenceon all four rotor blades.




EC 135


Collective Control

Control Rod

Control Rod

Upper Guidance Unit

Main Rotor Actuator




Collective Pitch Lever

Collective Shaft

EC 135


Cyclic Control

Signal Input

The cyclic control signals are given by moving the cyclic stick left orright (lateral control) and by pushing or pulling it (longitudinal control).

Cyclic Stick

The cyclic sticks are located in front of the pilot’s and copilot’s seat.Both sticks are linked via a torsion tube and a linkage mechanismunderneath the cabin floor.

Control Transmission

Longitudinal control inputs are transmitted via the cyclic shaft to ahorizontal control rod which leads to the lower guidance unit beneaththe control post.

Lateral control inputs are transmitted via a linkage which is connectedabove the cyclic shaft to the control stick, to a bell crank and to ahorizontal control rod which leads to the lower guidance unit beneaththe control post.

The lower guidance unit transfers longitudinal and lateral controlinputs as thrust motions to one vertical control rod each.

The left and the right bell crank of the upper guidance unit transmit the

The longitudinal control lever tilts about the axis of the correspondingbearing bushing and displaces the control ring of the swashplateforward to the right via a cyclic control link when pushing the stick

forward or backward to the left when pulling the cyclic stick aft.The lateral control lever tilts the swashplate forward to the left whenpushing the cyclic stick to the left and backward to the right whenpushing the stick to the right.

Vibration Decoupling Unit

The linkage for decoupling the vibrations is located between the upperguidance unit and the mouning plate of the main rotor actuator. This

unit supresses control inputs induced by vibrations from the main gearbox relatively to the fuselage. If there is a displacement between themain gearbox and the upper guidance unit, the decoupling rod causesa tilting of the guidance unit for compensation.




g pp gthrust motions to one horizontal control rod each.

One horizontal control rod displaces the input lever of the longitudinalcontrol piston (LH) and the other one displaces the input lever of thelateral control piston (RH) at the main rotor actuator.

Boosted inputs are transmitted behind the pistons to the longitudinalcontrol lever or to the lateral control lever of the mixing lever gear unit.

EC 135


Cyclic Control

Cyclic Stick

Vertical Control Rod(Lateral Control)

Vertical Control Rod(Longitudinal Control)

Upper Guidance Unit

Main Rotor Actuator

Horizontal Control RodPitch Axis

Vibration Decoupling Unit

Lower Guidance Unit

EHA for SAS(Pitch Axis)

EHA for SAS (Roll Axis)

Horizontal Control RodRoll Axis

Lateral ControlRod

Lateral Trim Control Rod

FWD

Cabin Floor




Cyclic ShaftBearing Support

Long. Control RodLong. TrimControl Rod

Cyclic Shaft

EC 135


Mixing Lever Assembly

General

The purpose of the mixing lever assembly is to transmit the threeamplified main rotor control signals (collective, longitudinal and lateral)to the swashplate.

Main Components

The main components of the mixing lever assembly are:

-- Collective control fork

-- Two cyclic control levers

Collective Control ForkThe collective fork is supported by the hinged support mounted on topof the main transmission. At the forked end it is connected to theslidingsleeve.

Cyclic Control Levers

The two cyclic control levers are mounted one on each side of the

collective control fork. As seen in flight direction, the lateral controllever is mounted to the RH side and the longitudinal control lever ismounted to the LH side of the collective fork.




EC 135


Mixing Lever Assembly

Short Control Rod Lateral

Cyclic Lever Lateral

Collective Fork

Swash Plate

Short ControlRodLongitudinal

Main Gear Box

Connecting RodLateral




Shim Plate

Cyclic Lever LongitudinalConnecting RodCollective

Connecting Rod

Longitudinal Hinged Support

EC 135


Transmission of Control Signals

Collective:

For increasing the vertical lift of the helicopter the swash plate has tobe raised evenly by the collective fork and the sliding sleeve (point 1

to point 1’).Thus the pivot points of the lateral and longitudinal levers have to beraised as well in order to avoid a cyclic input to the swash plate (point2 to point 2’ and point 3 to point 3’).

Longitudinal input (example forward flight):

The longitudinal lever raises point 3 to point 3’ and thereby tilts theswash plate. Thus the rotating pitch links, which are mounted at the

loading edge of the rotor blades, provide the maximum input approx.90° prior the tail position of the blades. Due to the gyroscopic effect,inertial blade mass and rotor characteristics the blades deliver thehighest lift at the tail position. the lowest lift is evident at the noseposition. The rotor plane tilts forward which causes the helicopter to flyforward.

For a rearward flight the swash plate has to be tilted in the oppositedirection (lowering of point 3) and the rotor plane will tilt to the rearaccording the principle described above.

Lateral input:

NOTE Transmission of cyclic signals is totally

independant of collective control inputs. Collectivecontrol signals are transferred to both, the sliding

sleeve and the two short control rods.




The lateral input for left and right follow the same principle as thelongitudinal control. Point 2 has to be raised or lowered and thehelicopter will turn left or right.

EC 135


Transmission of Cyclic and Collective Signals

Cyclic Control SignalCollective Control Signal

Swash Plate

Sliding Sleeve

Collective Fork

Longitudinal Lever

Lateral Lever

Short Control Rod

Axis a

1

2

3

3

3’

Axis a

3’

2’

1’




Input: Increase Thrust Input: Forward Flight

EC 135


Swash Plate

General

The swash plate transfers the rotor blade pitch change controlmovements from the stationary cyclic or collective control input to therotating blades.

Sliding Sleeve

The collective controlinputs move the sliding sleeve up or down. Insidethe sleeve two teflon liners are attached, which permit easy slidingmovement on the gearbox mounted support tube. Two bearing boltsat the top of the sliding sleeve retain the cardan ring. Two ball bearings

at the lower side connect to the collective control fork of the mixinglever unit.

Cardan Ring

The cardan ring contains four bearings, two for pivoting the slidingsleeve and two for pivoting the control ring. This arrangementconstitutes a gimbal mounting which enables the interconnectedcontrol ring to tilt in all directions about the vertical axis.

Control Ring

The stationary control ring transmits the cyclic inputs via the swashplate bearing to the rotating bearing ring It is connected to the mixing

Swash Plate Bearing

The swash plate bearing is a douplex ball bearing which connects thenonrotating control ring to the rotating bearing ring.

NOTE The swash plate bearing is the only rotating part of

the helicopter that is lubricated by grease.

Bearing Ring

The bearing ring is rotated synchronously with the rotor through thetwo scissors assemblies. The four forked lugs provide the attachementpoints for the rotating control rods. The connecting bolts from the twolevers integral with the bearing ring provide the attachment points for

the scissors assemblies.

Located within the bearing ring is a soft-iron pin which provides theimpulses for a magnetic pick-up for track and balance purposes.




plate bearing to the rotating bearing ring. It is connected to the mixinglever assembly by two control rods.

Also at the control ring provision is made for installation of a speedpickup for track and balance purposes.

EC 135


Swash Plate Assembly

Speed Pickup Mount

Connecting Bolt for

Scissors Assembly

Outer Ring

Inner Ring

Split Cover

Duplex Ball Bearing

Control Ring Nonrotating

Bearing Ring, Rotating




Teflon BushingControl Fork Bearing

Cardan Ring

Control Ring, Nonrotating

EC 135


Rotating Control Rod

General

The purpose of the rotating control rods is to transmit the flight controlsignals to the main rotor blades. Four rotating control rods are installed

between the rotating part of the swash plate and the pitch horns at therotor blades.

Components

Each rotating control rod consists of:

-- Two bearing rod ends

-- Two counter nuts

-- Two keyed washers

-- Rod body

Configuration

The bearing rod ends are screwed into the rod body by a coarse thread(MJ10x1.25) on one side and a fine thread (MJ10x1.00) on the otherside. The rod ends are secured in the rod body by a keyed washer and

a counter nut on each side. The counter nuts are additionallylockwired. To prevent corrosion inside the rod body of, the upper endis sealed by a sealing compound.

NOTE The metric threads of some high loaded bolted

connections might be designed according the MJstandard. Due to modifications in the thread root

area an improved stability is achieved. In addition

the self locking behaviour has been improved due

the selected relationship of thread diameter andpitch.For combinations or exchangeability of MJ and

standard ISO M threads the remarks in the IPC

have strictly to be followed. For identification the

letters “MJ” are imprinted on bolts/nuts.

WARNING The threads of the rod ends are marked by red

paint. These red areas must not be visible afteradjustment/installation.




NOTE The coarse thread must be located on the top. If

not, the adjustment for the blade track by rotating

the rod body is not as described in themaintenance manual.

EC 135


Rotating Control Rod

Sperical

Bearing withCoarse Thread

SphericalBearing with

Counter Nut

Keyed Washer

Rod Body




gFine Thread

EC 135


Driving Unit

General

The driving unit connects the swash plate to the rotor mast. Its purposeis to drive the rotating part of the swash plate. The driving unit connects

the bearing ring of the swash plate with the scissors clamp at the mainrotor mast.

Attachment

The driving unit is connected to the main rotor mast by two integratedlugs. Each of the two scissors assemblies are connected to the swashplate by means of a spherical bearing and a swash plate installed bolt.

NOTE The lettering OUTER SIDE on the lever faces

outboard.



EC 135


Track&Balance Installation

General

For balancing the main rotor two velocimeter and a magnetic speedpickup are installed. The wiring leads to a receptacle in the circuit

breaker panel 1 which is situated in a recess in the LH cargocompartment side cover.

Lateral Velocimeter

The lateral velocimeter is installed on the main transmission, thevertical velocimeter is located in the nose of the helicopter in the areabelow the copilot’s seat under the forward floor, next to the cyclic shaft.

Magnetic Pickup

The magnetic pickup is installed at the control ring of the swash plate. An iron interrupter pin is mounted in the rotating bearing ring of theswash plate.

NOTE Dummy velocimeter pickups or operating pickups

can be installed.




EC 135


Track&Balance Installation

Magnetic Pick

up

Cable Assy

Velocimeter M/R LAT

Bearing Ring

Interrupter Pin

Main Transmission

Receptacle forTrack&Balance

Circuit Breaker Panel 1

Forward Floor




DCRECEPT

TR&BALINFLT

3MJA

19VVA

10 5

Track&Balance

VelocimeterCyclic Shaft

Receptacle forDC Power SupplyT&B--Equipment

EC 135


Trim System

General

As the EC 135 is equipped with hydraulic boost units for main rotorcontrol, which amplify the control signals, no real control forces are

necessary at the control stick.Forbetter handling of the helicopter an artificial controlforce, givingthepilot a reference for stick displacement is desireable. For that reasontrim actuators with artificial force feel springs are installed in thenon--boosted section of the cyclic controls.

During flight the pilot does not only move the stick for a short time, e.g.flying a turn, but also for along time , e.g. during cruise. Holding the

cyclic stick against the artificial control force would fatique the pilot.Therefore the artificial control force can be trimmed to zero in eachstick position by electric motors and clutches in the trim actuators.

Trim Actuator

The longitudinal trim actuator is installed beneath the cabin floorcentered directly behind frame 1 and in front of the cyclic shaft. Theidentical lateral trim actuator is installed beneath the cabin floorcentered behind the cyclic shaft and in front of frame 2.

In the housing of an actuator there is mounted a DC motor, anelectro-mechanical clutch, a eddy current brake, a position sensor and

i f tifi i l f f l

It is installed beneath the cabin floor off-center right in front of thetorsion tube.

Control Board

The control board for the trim system is installed beneath the cabinfloor right behind the cross beam attached to the cabin floor. On thecontrol board there are mounted two relays for control of the DCmotors.

4--Way Trim Switches

The 4--way trim switches are installed on top of both cyclic control stick

grips, respectively.The desired trim position of the cyclic control is adjusted by the 4--waytrim switches.

Push Buttons

The push buttons ATT TRIM REL to release the trim position areinstalled on top of both cyclic stick grips, respectively.

Dual Controls

If dual controls are installed, the 4--way trim switch priority is set to trimaft / right, regardless whether the trim signal is triggered by the pilot orthe copilot




a spring for artificial force feel.

Trim Linkage

The longitudinal trim rod connects the output lever of the longitudinal

trim actuator with the torsion tube of the cyclic shaft for longitudinalcontrol.

the copilot.

Circuit Breaker

The circuit breakers TRIM ACT and ATT TRIM REL are mounted in the

overhead console.

EC 135


Trim System -- Locations

CDS

AUDIO

RES

4--Way Trim Switch ATT TRIM

Push Button ATT TRIM REL

Circuit Breaker ATT TRIM ACT

Circuit Breaker ATT TRIM REL

Cabin Floor




Trim ActuatorLateral

Trim ActuatorLongitudinal

Trim SystemControl Board

Cross Beam

EC 135


Function

The function of the longitudinal and lateral trim actuator is identical.

By operating the 4--way trim switch at the cyclic stick, the DC motor inthe trim actuator drives the primary reducer (irreversible wormgear)

and transmits the movement to the closed electrical clutch. With theclutch the primary reducer is connected to the secondary reducer andthe motor movement is transmitted to the output shaft. Via the outputlever and a control rod, the stick is moved into a new force free neutralposition.

The running direction of a trim motor is changed by a polarity reversal.The on--board circuitry with the relais and the two DC motors enablesfour running directions: Forward, aft, left, right.

When operating the 4--way trim switch only one of the four contactscan be closed. When releasing the switch, all four contacts are againopened.

During a cyclic control input the trim actuator output lever movestogether with the cyclic controls. With the trim actuator deenergized nomovement of the reduction geartrain is possible. By the relativemovement between the two plates, the spring becomes twisted, thus

creating an artificial control force.

Depressing the ATT TRIM RELEASE push button at the cyclic stickenergizes the electric clutch in the trim actuator. The clutch opens andseparates the secondary reducer fromthe primary reducer. This allows

After releasing the ATT TRIM RELEASE push button, a new force freestick position is maintained.

NOTE In case of accidental jamming of any internal trim

actuator parts, a higher control force has to be

applied to break a shear pin in the affected trimactuator output shaft. This allows free movement

in the respective direction without an artificialcontrol force. In that case the trim system in the

associated direction is disabled, too.




p y p ythe secondary reducer to turn and the spring to move in the force freeposition. To smooth this movement a damping device mounted withthe secondary reducer gives a torque resistance proportional to

speed.

EC 135


Trim System; Trim Actuator

Longitudinal Trim Rod

Cabin Floor

LateralTrim Rod

Frame 1

Cyclic Shaft

Side View

Top ViewFWD

SpringMovable Gear

Gear withShear Pin

Position Sensor

Electrically Activated Coupling

CentrifugalFriction Brake

DC Motor

Output Lever




LongitudinalTrim Actuator

Lateral Trim Actuator

Output Lever

EC 135






EC 135


Trim System -- Functional Diagram

PP10SPP10EPush Button ATT TRIM REL(Pilot)

4--Way Trim Switch ATT TRIM(Pilot)

Control Board

Push Button ATT TRIM REL(Copilot)

4--Way Trim Switch ATT TRIM (Copilot)

Forward

Rear

Right

Left

LRVH

Push Button

4--Way Trim Switch ATT TRIM

Left

Forward

Right

Rear




M M

Longitudinal Actuator Lateral Actuator

ATT TRIM REL


Tail Rotor Control

General

The tail rotor control changes the angle of incidence of the tail rotor

blades collectively. The tail rotor control is used for the yaw control.Control inputs are made by the pilot via the pedal assembly. The pedalinputs are superimposed by inputs from the Yaw Stability AugmentionSystem (YAW--SAS) via an electro-mechanical actuator. The inputsare boosted hydraulically and transmitted to the control spider whichchanges the blade angles.

Components

The tail rotor controls consist of the following assemblies:

-- Pedal assembly

-- Ball bearing control cable

-- Yaw--SAS actuator

-- Fenestron actuator (booster)

Pedal Assembly

The pedal assembly consists of:

-- 2 pedals

-- 2 pedal control rods

-- Bellcrank lever

Yaw Actuator

The yaw actuator is an actuator with an integral position feedback

(Smart electro-mechanical actuator, SEMA). It converts the stabilizingsignal produced by the fibre optic gyro (FOG) into a correspondingmechanical input to the tail rotor control linkage.

The series-connected yaw actuator operates between the ball bearingcontrol and the hydraulic Fenestron actuator. In consequence,stabilizing inputs from the yaw stability augmentation system and thecontrol inputs from the pilot are superimposed on each other.

Following a stabilizing input, the yaw actuator automatically recenterswithin its maximum stabilizing stroke range to ensure full stabilizinginput authority. The authority in the yaw actuator control is 8%.




Bellcrank lever

The pedal assemblys of the pilot and copilot are linked by a connectionrod.


Tail Rotor Control

Fenestron Actuator

Ball Bearing Control Cable

Yaw Actuator

Coupling for ConnectionRod to the Copilot’s PedalAssembly

Bell Crank Lever

HydraulicPressure Tube




Pedal Assembly

Assembly

Control Rod


Ball Bearing Control Cable

The ball bearing control cable (FLEXBALL) consists of a double--rowarrangement of steel balls leading through captive ball cages. Thesteel balls roll between two outer races and a center core. A flexballcasing encloses the races. Due to this construction the center core is

able to transmit identical tensile and compression forces.





Ball Bearing Control Cable (Flexball)

Casing

Outer Race

Center Core

Steel Ball

Ball Cage





Function of the Tail Rotor Control

The angle of incidence of the tail rotor blades can be varied within arange of --16.8û thru +34.2û.

If e.g. a control input “yawto the left” is made by actuating the left pedalof the pedal assembly, this input is transmitted as a tension motion viacontrol rods and the guidance unit to the ball bearing control.

The ball bearing control actuates a control rod in the Fenestron andthus the input of the yaw actuator. The yaw actuator superimposesadditional control inputs of the yaw stability augmentation system. Thepart of the control rod located behind the yaw control actuator pulls theinput lever.

The Fenestron actuator increases the force at the input lever andaxially shifts therotating control spider via its piston rod to the right. Thelevers of the control spider convert the axial motion into a positive twistof the rotor blades.





Tail Rotor Actuator

Return PipePressure Pipe

Input Lever

Control Rod

Bleed Valve





Hydraulic System

General

The hydraulic system is used to boost the manual control inputs of the

pilot. At the same time the reset forces of the rotor blades are blocked.Components

The hydraulic system consists of the following components:

-- Two identical pressure systems

-- Main rotor actuator

-- Fenestron actuator

-- Indicating and testing system

Leading Particulars

Operating Pressure 103 bar

Return Pressure 1.40 -- 1.75 bar

Hydraulic Fluid acc. MIL--H 5606 (F)

Fluid Capacity 1.0 l (SYS1), 1.2 l (SYS 2)

Reservoir Capacity 0.8 l

NOTE To prevent a contamination and blockage, it is

recommended that hydraulic fluid stored in cans

should not be used when it is older than 3 years.

Location

The components of the hydraulic power system are installed on the

front of the main transmission and in the cockpit. Two pressure supplysystems are installed on top of the fan gearboxes. The fan gearboxesare attached to the left-hand and right-hand forward side of the maintransmission. The main rotor actuator is installed in the center of theforward side of the main transmission. The Fenestron actuator isinstalledinside the stator hub of the Fenestron. Hydraulic lines connectthe pressure supply systems to the main rotor actuator and theFenestron actuator. The components of the indicating and testing

system are part of the pressure supply systems. The related switchesand displays are installed in the overhead panel and in the instrumentpanel.




y


Pressure Supply System

Lateral Control Rod Output Lever

Mixing Lever Unit

Input Lever

FWD

Main Transmission

Pressure SupplySystem 1

Pressure Supply System 2

Collective Control Rod Refill Port System 2

Refill PortSystem 1




Actuator

Longitudinal Control Rod


Redundancy Provision

The hydraulic power system is a dual system. It has two identicalpressure supply systems, system 1 and system 2, that operateindependently. Under normal operating conditions both pressuresupply systems simultaneously generate the entire pressure for

boosting the main rotor controls. System 2 in addition also boosts thetail rotor controls. If one of the pressure supply systems fails, theremaining system continues to supply the main rotor actuator. Thiscauses the operating force of the mechano-hydraulically operatedmain rotor actuator to decrease to half.

Only the failure of system 2 causes the tail rotor control to operatewithout pressure. Failure of system 1 has no effect on the Fenestronactuator.





Hydraulic Power System

System 1

System 2

CDS/CPDS

System 1 System 2

Main Rotor Actuator

Relais Relais

TestSwitch

Pump Pump

Reser-voir

Reser-voir

ValveBlock

ValveBlock

Fenestron

Actuator

HYD PRESS HYD PRESS





Pressure Supply Systems

General

The pressure supply systems 1 and 2 are two identical systems. Theyindependently supply the hydraulic actuators with operating pressure.

Components

Each pressure supply system consists of:

-- Hydraulic pump

-- Reservoir

-- Valve block

-- Hydraulic lines

NOTE To prevent the hydraulic systems from

contamination an external ground cart must not beconnected. System tests can be carried out by

operating the hydraulic pumps with a special tool.

To refill the systems a container with a hand--pump

and a fine filter is available.





Flight Control

Pressure Supply System

Sight Glass

Reservoir

Return Line Port

Level Indicator

Maintenance Port

Bleed Valve

Supply Line Port

Leak Oil Port

Valve Block

Filter

Pressure Switch

Solenoid Valve

MIN Marker

MAX Marker




Pump


Flight Control

Hydraulic Pump

The hydraulic pump is an integral part of the pressure system. Allconnections (i.e. pressure line, suction line case drain) are made bychannels and bores in the valve block.

The pump is conventional piston type wherein a cylinder barrelcontaining nine pistons is driven by the accessory drive of the maintransmission.

The pistons are constrained by the rotating part of the backplate andball--and--socket--joints shoes which are hydrostatically balanced. Asthe barrel rotates, the pistons intaking and discharding fluid through astationary valve surface (control plate) on the port cap. The length of the piston stroke, and thereby the displaced volume is determined by

the angle of the nonrotating part of the backplate. This angle iscontrolled by a spring acting against system pressure on the cam of the nonrotating part.

NOTE The longer the stroke of the pistons, the larger thevolume of fluid delivered.

Leading Particulars

Speed 5145 RPM

Preloaded pressure in the reservoir 1.40--1.75 bar

Reservoir Capacity 0.8 l

Low pressure relief valve Opens at 6.5 barHigh pressure relief valve Opens at 122 bar

Pressure switch (increasing pressure) Opens at 82.7 bar

Pressure switch (decreasing pressure) Closes at 69 +/-- 3.4 bar





Flight Control

Splined Shaft

Inlet Port

(from Reservoir)

Outlet Port

(to Valve Block)

Piston orPlunger

Adjustment Screw(Factory Set)

Control

PistonSpring

Backplate(Fixed Part)

Backplate(Rotating Part)

Case Drain

Piston orPlunger

Fluid FlowDecrease

Fluid FlowIncrease

Backplate

Backplate(Rotating Part)

Inlet Port(from Reservoir)

Outlet Port(to Valve Block)

Pump Shaft

Barrel

Hydraulic Pump

Seal Drain




Backplate,(Fixed Part)


Flight Control

Reservoir

The reservoir stores the hydraulic fluid. The necessary preloadpressure is generated by the double actuated piston in the reservoir.The operating pressure applies a force on the smaller piston. As aresult the larger piston pressurizes the reservoir. With the ratio

between the both piston areas (1:60) and an operating pressure of 103bar, a return pressure of 1.40 -- 1.75 bar is created in the reservoir toprepressurize the pump suction side.

A pressure relief valve avoids a damage of the reservoir caused byoverpressure. It opens at a pressure of 6.5 bar and relieves outhydraulic fluid to the leak oil port.

Both the reservoirs with the valve blocks attached to their forward side,

are installed on the hydraulic pumps. A support bracket also attachesthem to the main transmission.

The sight glass on the top of the reservoir serves as an indicator for theamount of air in the system.

A fluid level indicator is installed on the rear side of the reservoir.

NOTE The sight glass must be half full of hydraulic fluid

minimum. Otherwise the system has to be bled. Asave flight operation is assured as long as fluid is

visible in the sight glass.

Valve Block

The valve block contains all the valves and control lines to control andtest the hydraulic system.

Directly after the hydraulic pump there is a non return valve to preventa reversal of the fluid direction.

The filter prevents the system from contamination.

The high pressure relief valve prevents overloading of the system. Thevalve opens at a pressure of 122 bar and excessive pressure isreleased to the return side.

A solenoid valve, the shut off valve and the pressure switch are partof the indication and test system. Energizing the solenoid valve causes

the shut off valve to close. The resulting decrease in pressure causesthe pressure switch to close and to send a signal to the cockpit for lowpressure caution indication.

Maintenance

For maintenance purpose the following ports are available:

-- Bleed valve/sightglass for detection and bleeding of trapped air (in system 2 a second bleed valve is mounted

at the fenestron actuator).

-- Maintenance port for pressure monitoring (high pressureside).

-- Maintenance port for draining and refilling the system (lowpressure side).




NOTE Due to internal piping the refill port is mounted at

the actuator carrier plate in front of the main gearbox opposite of the respective system.


Flight Control

Hydraulic Valve Block -- Non Pressurized

Bleed Valve

Level Indicator

MIN MAX

Low Pressure Piston

Vent

Reservoir

Low Pressure Relief Valve

Non Return Valve

Pressure Monitoring

(Maintenance)

Filter

Solenoid Valve

Pressure Switch

High Pressure Relief Valve

Shut-Off Valve

Pressure Out

Return In

Sight Glass




Port to Drain System

Seal Drain

Pump


Flight Control

Hydraulic Valve Block -- Normal Operation

The hydraulic pump delivers a constant pressure of 103 barvia thenonreturn valve and the filter to following locations:

Location 1: Small piston chamber (left section) of the reservoir pistonunit

Result: The force at the piston rod due to the high pressure in the smallchamber creates the low pressure in the large piston chamber (rightsection) with a relationship of 60:1.

Location 2: Right side of the shut off valve

Result: The force generated by the high pressure piston (right side)and the spring force override the force created by the low pressure

piston and keep the shut off valve in the opened position.Location 3: center section of the shut off valve

Result: As the shut off valve is being kept in the open position the highpressure outlet is pressurized. The pressure switch is open andtherefore the caution HYD PRESS in the CDS/CPDS is suppressed.In this situation the respective main rotor actuator system is suppliedwith high pressure. The returning fluid from the actuators is recycledby the hydraulic pump or led to the reservoir, depending on the flowdemand.

Location 4: Solenoid Valve inlet

Result: In this situation none





Flight Control

Hydraulic Valve Block -- Normal Operation

High Pressure

Low Pressure





Flight Control

Hydraulic Valve Block -- Test

Test Function activated:

For the single system test on ground one system has to be shut off withthe spring loaded test switch in the overhead panel. During the test thesolenoid valve is activated and opens the high pressure inlet forthe left

side of the shut off valve.

Result: the piston of the shut off valve travels to the right end stopbecause the force created by the larger piston surface and the highpressure is greater than the force created by the spring and the smallerpiston surface with high pressure applied.

The Pressure outlet is blocked and the pressure switch closes(Caution HYD PRESS in the CDS/CPDS for the respective system

comes on).The pressure outlet line and the main rotor actuator of the deactivatedsystem are connected to the return pressure as long as the testsituation is evident.

Test function deactivated:

The test switch is released to the norm position, the solenoid valvecloses the high pressure inlet for the left shut off valve piston and the

shut off valve reverts to the open position again. The fluid of the leftpiston chamber is pushed into the low pressure line which is openedsimultaneously.

Result: The pressure switch opens again (caution HYD PRESS goesoff) and the main rotor actuators are supplied with high pressure again.

NOTE Both hydraulic systems can be tested with this

procedure. Only when testing system 1 (system 2is inactive) there is no pressure supply to the

fenestron actuator.

WARNING Never activate the hydraulic test switch inflight.





Flight Control

Hydraulic Valve Block -- Test

High Pressure

Low Pressure





Flight Control

Hydraulic Actuators

General

Due to the high reset forces which react on thecontrols when changingthe blade pitch, hydraulic actuators transmit boosted control inputs to

the rotor system.

The main rotor actuator consists of three adjacent hydraulic actuators.It is installed at the front part of the main rotor gearbox by means of anattachment and supply plate.

Assembly

The hydraulic actuator mainly consists of:

-- Servo valve

-- Boost cylinder

Description of the Follow Up Principle

Fluid Flow

System pressure is supplied from the pump via the valve block to thecontrol spool. Depending on the control spool position the left or rightside of the piston is pressurized. The boost piston moves in thecorresponding direction. The low pressure fluid from the notpressurized chamber is led back to the return line into the reservoir.

With the control spool in the neutral position, no boost pistonmovement is possible, because the pressure line as wellas both return

lines are closed. The boost piston is hydraulically blocked.

Control Input

The input control rod is moved to the right. At the moment of the input,the boost piston cannot move, because it is still hydraulically blocked.Therefor, when the control input rod moves to the right, the controllever turns around the pivot point at the boost piston. The control spoolin the control valve is pulled to the left by means of the connecting rod

and the lever. This opens the right port of the servo valve, directinghydraulic pressure into the right chamber of the boost cylinder. In thesame moment the return line of the left chamber opens and the fluidmoves back to the reservoir.





Flight Control

Hydraulic Actuator -- Basic System Function

Boost Cylinder

Boost Piston

Output toSwash Plate

Pivot Point

Control Lever

Input Control Rod

Pump

Reservoir

Connecting Rod

Control Spool

Lever

Pressure Line

Return Line

Starting Input





Flight Control

Reaction of the Boost Actuator

The hydraulic pressure in the right chamber of the boost cylindercauses the piston to move to the left. Low pressure fluid from the leftboost cylinder chamber is ported to the servo valve and to the reservoirvia the return line.

With the boost piston moving to the left and a constant movement atthe input control rod to the right, the middle point of the control leverbecomes to the pivot point where the control lever turns around. Thecontrol spool remains pulled to the left end stop by the connecting rodas long as the input continues.

Input Stop

When there is an input stop, the upper bearing of the control leverbecomes the pivot point. As the control spool is still in the openposition, the boost piston moves until the control spool is pushed backin the closed position by the connecting rod and the lever.

With the control spool in the neutral postion no further hydraulic flowis possible and the boost piston becomes hydraulically blocked again.This short time delay is not feelable in the controls.





Flight Control

Hydraulic Actuator -- Basic System Function

Pivot Point

Movement here

Input Stop

Pivot Point

Movement here until the controlspool is in neutral position and

blocks hydraulically the boost piston

Movement

Boost Cylinder

Boost Piston

Output to

Swash Plate

Pivot Point

Control LeverConnecting Rod

Input Control Rod

Lever

Pressure Line

Return Line

Continued Input




Flight Control

MHA -- Non Pressurized

Input Rod

Connecting Rod

Control Lever

Input Lever

Valve Sleeve

Control Spool

Shut-Off/Bypass Valve

Strong Spring

Pressure Port P2with check Valve

Return Port R2

Pressure Port P1with Check Valve

Return Port R1

Boost Piston




Test ButtonShut Off/Bypass Valve

Weak Spring


Flight Control

MHA -- Pressurized, no Movement





Flight Control

MHA -- Pressurized with Movement





Flight Control

Mechanical Override

Purpose

In order to assure the function of the hydraulic system in case onecontrol spool jams, a mechanical override is installed to each system.Because the control spools of the two systems are mechanically linkedto each other, a jammed control spool in one system would causeblocking of the corresponding control spool within the other system.

Assembly

The control spool is moving in a valve sleeve, which is kept in positionby two springs. A test button is installed to the springs housing.

Function

In case of a jammed control spool, every control input will shift thecontrol spool and the valve sleeve together against the spring forces.The first displacement of the valve sleeve causes the opening of thecontrol line to return pressure, thus first the shut-off valve closes andthen the bypass valve opens. A bypass around the boost pistonchambers of the respective system is established.

NOTE In case of a jammed control spool an increasedcontrol force in the affected axis will be observed.





Flight Control

MHA -- Mechanical Override of System 1

Control Spool Blocked

Normal Situation





Flight Control

System Test

A test button, installed to each spring housing allows checking thevalve sleeve for free movement. Pressing the test button will first closethe gap between button and sleeve then, increase of applied force willcause the displacement of the valve sleeve. The test button returns to

its normal position because of the spring forces and after the returnpressure has been built up.

NOTE If, after closing the gap, no further movement is

possible against the spring force, the valve sleeve

may be blocked in the housing or the control spoolmay be jammed in the valve sleeve





Flight Control

Valve Sleeve Test

Test Button

Gap

Control Spool

Valve Sleeve

Springcompressed

Springs forMechanicalOverride




Normal Position Closed Gap Position Displaced Valve Sleeve Position


Flight Control

Electro-- Hydraulic Actuator EHA

General

In addition to the mechanical inputs by thepilot thegyro based Stability Augmentation System SAS superimposes the control output to themain rotor lateral and longitudinal axis in system 1.

Function

The basic functions concerning boost piston and control spool aresimilar to the mechano-hydraulic actuator as described for thecollective axis.

In order to allow the control cylinder inputs to the control spool andthereby to the control output the mechanical linkage is modified. As

long as the SAS is inactive the control cylinder is centered by twosprings and the control spool moves only after an input coming fromthe pilot.

When the supply line from P1 to the electro valve is pressurized thecontrol pressure builts up via the solenoid valve and closes the by passvalve.

Thus the operating pressure can be directed into one of the control

piston chambers by the piston unit in the electro valve. The position of the piston unit is controlled by the SAS computer via electromagneticsignals to the servo valve coils. The position sensor signal is used asa feedback signal for the control loop in the SAS computer.

With both control piston chambers connected no differential pressurebuild up and no influence from the SAS is possible.





Flight Control

EHA -- Normal Operation with SAS Input

SAS Control Piston

Shut-Off/Bypass Valve

Solenoid Valve

Servo Valve

Control PressureChamber

Position Sensor




To SAS Computer


Flight Control

EHA -- SAS Decoupled

The complete SAS (P&R and YAW SAS) can be switched off by thepilot manually. In this case the solenoid valve is activated directly bya switch in the cockpit.

The control pressure will be relieved to the return line and the spring

force will open the by pass valve. Then the control piston will becentered from present position. The restrictor in the by pass valvecauses a delay in order to avoid a control input. Therafter the controlspool and the boost piston move only after a mechanical input via theflight controls.

NOTE In case of hydraulic system 1 failure the P&R SAS

will be inoperative.





Flight Control

EHA -- SAS Decoupled





Flight Control

Indication and Testing System

General

Each system has a pressure switch to monitor the operating pressure.Power is supplied through the busbar PP10E resp. PP20E and the

related circuit breakers.

With system pressure above approx. 83 bar, the pressure switch isopen and the related relay is not energized. There is no CAUTIONindication.

System pressure of less than approx. 69 bar closes the pressureswitch and energizes the related relay. The CAUTION indication HYDPRESS is displayed on display segment SYSTEM I or SYSTEM II on

CDS/CPDS. The range of hysteresis between 69 and 83 bar is bymeans of the different friction in the pressure switches.

Components

The components of the indicating and testing system are:

-- Pressure switch for System 1 / 2

-- Solenoid valve for System 1 / 2

-- Shut-off valve for System 1 / 2-- Circuit breaker HYD--P SYS 1 / 2

-- Relay for System 1 / 2

-- Display system CDS/CPDS

-- Test switch (spring loaded)

Test Procedure

As both hydraulic systems operate simultaneously one system has tobe switched off to test the other. Testing System 2 (test switch in

position SYS 2) system 1 is switched off (and vice versa) via thesolenoid valve. The pressure in System 1 drops and the pressureswitch activates the CDS/CPDS caution HYD PRESS in system 1.With small control inputs on ground the pilot can test the response of the respective system.

NOTE Testing System 1 the pedal forces will increase

because System 2 and therefor the fenestron

actuator is switched off.

WARNING The test has to be performed on ground only.





Flight Control

Hydraulic System -- Indication and Testing System

Relay (SYS II)

Relay (SYS I)

Circuit BreakerHYD P SYS I Circuit Breaker

HYD P SYS II

Test SwitchHYD SYS I/II





Flight Control

Fenestron Actuator

General

The Fenestron actuator is used for boosting the inputs for the tail rotorcontrol. It is bolted to the tail rotor gearbox. It transmits pedal inputs to

the control spider for changing the angle of incidence of the tail rotorblades. Integrated in the Fenestron actuator are the stops for themaximum and minimum control range. The actuator is supplied withpressure by the pressure system 2.

Function

Without hydraulic pressure the two springs with different spring ratekeep the bypass valve (weak spring) in the opened and the shut--off

valve (strong spring) in the closed position.

Thus the power piston can travel freely and the pilot is able to giveinputs to the tail rotor rotor by means of the mechanical linkage only.

When operating pressure fills the shut--off valve inlet chamber and thecontrol chamber through the hollow piston rod, the valve unit starts totravel to the left. First the by pass closes (weak spring), second theshut-off valve opens and gives the pressure free to the control spool

inlet.

The input lever is connected with the piston rod of the power piston viathe control lever. Pulling the input lever displaces the control spool tothe right and the operating pressure enters the left power pistonchamber which causes again a movement to the right as long as theinput lever continues to travel (and vice versa).

The control spool closes as soon as the required position of the powerpiston has been reached (input lever stops the movement) due to thefeedback of the control lever.

The movement of the power piston is stopped and the power piston iskept in its position until a new control input is made.

If the pressure drops in system 2, the shut-off valve closes and theby--pass valve opens. Both piston chambers of the boost cylinders areconnected and the mechanical control can displace the power piston.

The control spoolnormally travels in the valve sleeve which is centeredby two springs. If the control spool is blocked the valve sleeve can be

shifted against the spring pressure. Thus the control line is directlyconnected to the return line. If the pressure drops in the control line,the bypass valve switches the system off via the shut-off valve unit asdescribed above. The pilot will feel slightly higher control forces in theaffected axis because one of the springs at the valve sleeve has to becompressed.

The function of the test button corresponds the System Test of themechanical override in the MHA/EHA schematic.





Flight Control

Fenestron Actuator

Control Line

Valve Sleeve

Control Spool

Input Lever

Control Lever

Power Piston

Bypass Valve

Shut-Off Valve

Strainer

Return Port

Pressure Port

Test Button

Output Lever

Weak Spring

Strong Spring Control Chamber





Flight Control

Three Axis Stability Augmentation System SAS

General

The helicopter can be equipped with an optional 3-axis stabilityaugmentation system (SAS).

The 3-axis stability augmentation system comprises the followingindependent subsystems:

-- Yaw stability augmentation system (standard equipment)

-- Pitch and roll stability augmentation system (option)

Yaw Stability Augmentation System

General

Theyawstability augmentationsystem applieslimited authority controlinputs to the tail rotor control linkage.

The yaw SAS operates independently of the other flight controlsystems and provides the following functions:

-- Enhancement of the dynamic yaw stability

-- Damping of gust effects on the yaw axis

The system is designed for “feet-on” operation, thereby requiring thepilot to provide helicopter yaw control by operating the pedals. In turn,the pilot experiences improved handling qualities while at the sametime retaining full control input authority.

System Components

The yaw stability augmentation system consists of the followingcomponents:

-- Fiber optical gyro FOG-- Yaw actuator

-- Circuit breaker YAW SAS

-- Cut-off switch SAS DCPL

-- Re-engagement switch SAS CONT





Flight Control

3 Axis SAS (CDS Version) -- Locations

Pitch EHA

P&R SAS Computer

Yaw Gyro

Yaw SEMA

Roll EHA

T i A t t

Cyclic Stick

Overhead Panel

T i A t t

CDS

Pitch Gyro

Roll Gyro

Pitch Gyro (DPIFR)

Pitch SEMA (DPIFR)




Trim ActuatorTrim Actuator


Fiber Optical Gyro FOG

The fiber optical gyro (FOG) is installed on the engine deck within thestructure of the tail boom attachment cone between frame 7 and frame8. It can be accessed when the avionic plate is lowered.

The fiber optical gyro controls helicopter acceleration around the

vertical axis. A variation in the yaw rate within a specific frequencybandwidth causes the FOG to transmit an electrical stabilizing signalto the yaw actuator. The FOG is equipped with an electronic validitycontrol loop to monitor the operational readiness of the system.

Yaw Actuator

The yaw actuator is installed in the Fenestron structure. It is anactuator with an integral position feedback (SMART

electro-mechanical actuator SEMA). It converts the stabilizing signalproduced by the FOG into a corresponding mechanical input to the tailrotor control linkage.

The series-connected yaw actuator operates between the ball bearingcontrol and the hydraulic Fenestron actuator. In consequence,stabilizing inputs from the yaw stability augmentation system and thecontrol inputs from the pilot are superimposed on each other.

Following a stabilizing input, the yaw actuator automatically recenterswithin its maximum stabilizing stroke range to ensure full stabilizinginput authority.

Circuit Breaker YAW SAS

The circuit breaker YAW SAS is located in the top right-hand sectionof the overhead panel.

Switch SAS DCPL

The cut-off switch SAS DCPL is located on the extreme left on theupper end of the cyclic stick grip.

In the case of blockage of the yaw actuator, the system can bedisengaged through the cut-off switch SAS DCPL. The cut-off switch

interrupts the engage signal to the FOG.

Switch SAS CONT

The re-engagement switch SAS CONT is located in the top left--handcorner of the cyclic stick grip and is used to reactivate the system afterthe cut-off switch has been operated (reactivation is also possible bydepressing circuit breaker YAW SAS) . The re-engagement switchreconnects the engage signal to the FOG.

CDS/CPDS Display

The Caution YAW SAS appears in the MISC field if the Yaw SAS isdecoupled





Functional Schematic -- Yaw SAS

FOG

SAS

PP20E

P&R

Y RST

SEMA

YAW SAS

DCPL

Fenestron Actuator

CDS/CPDS Display

Blade Pitch Change

Re-engagementSwitch

Cut-Off Switch

Yaw Rate

Pilot Yaw Control Inputs Pilot + Yaw Actuator Control Inputs

Flexball Cable

Cut-Off

Switch





Pitch & Roll Stability Augmentation System

General

The pitch and roll stability augmentation system, which is also anindependent system, is used for stabilizing the attitude of the

helicopter about the longitudinal and lateral axes. It applies limitedauthority stabilizing inputs to the main rotor controls.

System Components

The pitch and roll stability augmentation system consists of thefollowing components:

-- Pitch and roll SAS computer

-- Electro-hydraulic actuators (EHA) (2 off)-- Circuit breaker P/R SAS for 28 V DC

-- Circuit breaker ROLL SAS and PITCH SASfor 26 V AC / 400 Hz

-- Cut-off switch SAS DCPL

-- Re-engagement switch SAS CONT

-- 2 Attitude gyros

Pitch and Roll SAS Computer

The pitch and roll SAS computer is located in the left--hand sidechannel in the floor structure and uses the input signals from theattitude gyros to compute the stabilizing input signals for theelectro-hydraulic actuators (EHA). An integral electronic validitycontrol loop within the SAS computer monitors operational readinessof the system. Position signals from both trim actuators are used by the

SAS computer to determine whether the pilot is overriding an SAS

control input. This prevents the SAS from working against pilot stickinputs.

A position sensor (LVDT) in the electro--hydraulic actuators (EHA)

supply the SAS computer with actuator position feedback signals.

Electro-Hydraulic Actuators

The electro-hydraulic actuator (EHA) is integrated into the housing of the mechano-hydraulic actuator in the main rotor actuator.

The electro-hydraulic actuator (EHA) in the pitch and the roll axesconverts the electrical stabilizing signals to mechanical inputs. Whenthe electro-servo valve is excited, a hydraulic control cylinder operatesto move the control spool of the mechanical-hydraulic actuator MHA,thereby adding stabilizing inputs to the MHA of the respective axis. Asa result, the stabilizing inputs from the pitch and roll stabilityaugmentation system are superimposed on the pilot stick inputs.

Following a stabilizing input, the EHA automatically recenters withinits maximum stabilizing stroke range to ensure full stabilizing inputauthority.

Circuit Breaker P/R SAS (DC System)

The circuit breaker P/R SAS is located in the upper LH section of theoverhead panel. The busbar PP10E supplies the P&R SAS system 28V DC through the circuit breaker P/R SAS.




SAS computer to determine whether the pilot is overriding an SAS


Functional Schematic -- Pitch and Roll SAS

SAS

PP10E

P&R

Y RST

CDS/CAD

P/R SAS

VG / HOR

26VAC I DCPL

VG / HOR

26VAC II

Lateral Trim Actuator

Long. Trim Actuator

Roll Attitude

Pitch Attitude

Pilot Control Inputs

MHA for Pitch Axis

MHA for Roll Axis

Blade Pitch

Blade Pitch

SASComputer

EHA

EHA

EHA + PilotControl Inputs

Re-engagementSwitch

Cut--Off Switch

Fast Erect

Cut--Off

Switch





Circuit Breaker Roll SAS and Pitch SAS (AC System)

The SAS computer is also supplied with 26 V AC / 400 Hz from busbar26 V AC BUS I and II through the circuit breaker ROLL SAS andPITCH SAS .The circuit breaker ROLL SAS is located in the upper LHsection, the circuit breaker PITCH SAS in the upper RH section of theoverhead panel.

The system is operative when its power supply is on. It becomesinoperative when the power is removed by pulling one of the threecircuit breakers.

Cut-Off Switch SAS DCPL

The cut--off switch SAS DCPL is located on the extreme left on theupper end of the cyclic stick grip.

If the electro-hydraulic actuators should become jammed, the systemcan be disengaged by actuating cut-off switch SAS DCPL.

The cut-off switch removes the engage signal to the SAS computer.

Re-engagement Switch SAS CONT

The re-engagement switch SAS CONT is located in the top left--handcorner of the cyclic stick grip and used to reactivate the system after

the cut-off switch has been actuated (reactivation is also possible bypulling and depressing the circuit breaker P/R SAS). There-engagement switch reconnects the engage signal to the SAScomputer.

Attitude Gyros

Depending on the equipment of the helicopter, there is one artificialhorizon installed in the instrument panel and one vertical gyro installedin the subfloor assy. As an equipment variant there are two gyrosinstalled in the subfloor assy.

The attitude gyros detect changes in the pitch and roll attitude of thehelicopter. These changes are applied to theSAS computer in the formof electrical signals. The roll signal comes from the vertical gyro 1, thepitch signal comes from the vertical gyro 2 or from the artificial horizon.

CDS/CPDS Display

The annunciation P/R SAS is displayed on the CDS/CPDS when thepower supply is interrupted or a fault occurs in the EHS, SAScomputer,

or attitude gyro.





Pitch Damper (DPIFR)

General

For Dual Pilot IFR certification an additional pitch damper has to beinstalled in order to compensate excessive pitch changes (e.g. EHArunaway).

System Components

The pitch damper system comprises the following:

-- Pitch Gyro

-- Pitch SEMA

-- Switch P&R / Y / P DAMPER RST

-- Circuit Breaker PITCH DAMPER

-- Indication P DAMPER

Pitch Gyro

The pitch rate gyro (FOG, Fibre Optic Gyro) is installed in the LH sidechannel near to the SAS computer and measures angular changes of the helicopter in its pitch axis.

The pitch rate gyro provides digital signals for control of the pitch

SEMA.

The power supply for the system is provided via the P DAMPER circuitbreaker located in the overhead panel.

Pitch SEMA

The pitch SEMA is integrated in the horizontal control rod which leadsfrom the upper guidance unit to the main rotor actuator for longitudinal

control.

The SEMA is installed in series with the pilot’s longitudinal control. Itsends limited control signals directly to the actuator without the cyclicstick being moved.

The actuator and a servo control loop are contained in the pitch SEMA

casing.The electronics of the servo control loop includes a monitoring systemwhich detects and corrects internal defects in the servo control loopitself and control signal errors.

In the SAS mode, the pitch SEMA only works as a rate damper and isactive when the pitch EHA is out of order and has not centered in themiddle during NORM operation.

Switch P&R / Y / P Damper RST

The switch P&R / Y / P DAMPER is locatedon the left onthe upper endof the cyclic stick grip. The 3--way switch is used to engage theindividual functions.

Circuit Breaker

The circuit breaker PITCH DAMPER is installed in the overhead panel

and supplied via the ESS. BUS II.

Indication PITCH DAMPER

A failure of the pitch damper is indicated with the caution P DAMPERin the MISC field of the CDS/CPDS. Additionally, the indication lightPITCH DAMPER on the left side of the Warning Unit comes up.





Pitch Damper -- Indication and Switch

P&R // Y // P DAMPER RST

Pitch Damper

YawSAS

2--Axis P&R SAS

PITCH DP/R SAS

YAW SAS

CDS/CPDS

3--Way Switch: Movement to engage the respective system.

SAS DCPL

Warning Indication PITCH DAMPER




EC 135Training ManualLanding Gear

Landing Gear





Table of Contents

Landing Gear 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Measurement of Ground Clearance 8. . . . . . . . . . . . . . . . . . . . .





Landing Gear

General

The landing gear carries the the weight of the helicopter on the groundand absorbs landing impact loads. It is attached through four fittingsto the lower part of the floor structure.

To prevent the fuselage from beeing over--stressed during touchdown, the bearing rings on the cross tubes are swivelling in theirbrackets, so that all forces are absorbed by bending the cross tubes.

The landing gear consists of two aluminum cross tubes and twoaluminum skids which are clamped together by aluminum skid shoes.

In its basic configuration, the landing gear is equipped with parallel

skids for landings on prepared surfaces. A skid track of 2.3 m providesthe helicopter with good stability when standing on the ground.

The landing gear can be fitted with optional equipments likeEmergency Flotation System, Multi--Purpose Carrier or High LandingGear to meet changing operational requirements.

To meet changing ground conditions, the landing gear can be fittedwith the optional supplementary landing provisions.

Components

The landing gear consists of:

-- Two crosstubes

-- Two skids

-- Four skid shoes

-- Four bonding jumpers

-- Two entrance steps

Crosstubes

The two crosstubes of the landing gear cushion landing impact loadson the fuselage by bending.

The crosstubes are mounted laterally approx. 2 m (6.6 ft) apart. Thelanding gear is attached by the crosstubes to four landing gear fittingswhich are integral with the floor structure of the fuselage.

Both crosstubes are hollow aluminum tubes of circular cross section.They are connected to the landing gear fittings through four bearingrings which are each held by two fitted bolts in the landing gear fittings.Each bearing ring is retained by a set screw ring clamped on thecrosstube.

For the purpose of jacking the helicopter, a jacking bracket can bepositioned below each of the 4 landing gear fittings.

The helicopter can be weighed by installing a weighing bracketcentrally on the forward crosstube.

Touch Down Limitations

The two aluminum cross tubes can absorb all forces, resulting from

touch down speeds up to approx. 1m/s (depending on helicopter massand ground harness).

Higher touch down speeds will result in plastic deformations of thecross tubes. Touch down speeds between 1 m/s and 2.5 m/s will notdamage the fuselage.





Landing Gear

Entrance Step

Skid

Crosstube

Skid Shoe

Protection Plate

Bonding Jumper

Bushing

Bearing Ring





Skids

Both skids, which are aluminum tubes of circular cross section, arecurved upward at their forward ends.

On the underside of each skid, one small aft and two bigger forwardskid protective plates are attached by screws. The skid protectiveplates are exposed to a high degree of wear because they are in directcontact with the ground.

Skid Shoes

The four skid shoes connect the skids to the crosstubes to form aspatial frame. They make for a stiff connection, thereby giving thelanding gear stability.

Each skid shoe is connected to the crosstube by a single bolt. The

saddle-shaped end of the skid shoe retains the skid through two splitclamps which are each tightened by two screws.

Bonding Jumpers

Bonding jumpers are installed between the crosstubes and skids andthe crosstubes and the floor structure to electrically connect theisolated attaching hardware. The bonding jumpers enable staticelectricity to be discharged from the surface of the helicopter to theground.

Entrance Steps

The two entrance steps which are crosstube--mounted above theskids, are provided to give boarding assistance to crew members andpassengers. The aft end of each entrance step is raised somewhat tofacilitate access to the lower maintenance step.

The V--profiles of the entrance step are of fiber composite constructionand the upper part is made of aluminium. They are each attached totheir respective crosstube by two fittings.





Measurement of Ground Clearance

General

If deformation of cross tubes is evident or suspected, the groundclearance of helicopter must be measured.

-- The ground clearance at the forward cross tube must notbe less than 460 mm.

-- The ground clearance at the aft cross tube must not beless than 360 mm

Procedure

The measurement must be carried out from a point in the middle of thefuselage located directly in front or behind the cross tubes. If the

minimum value is not reached, the respective cross tube must bechanged.

NOTE The measurement must be taken with a non loadedlanding gear. For this purpose the helicopter must

be jacked.





05 9

Measurement of Ground Clearance

Min. 460 mm Min. 360 mm




EC 135Training ManualPower Plant

06 1J lF i i d i f i l

Power Plant




06 3July 2002For training and information only







General Description of Power Plant

General

The power plant system of the EC 135 comprises all systems andsubsystems necessary for proper engine operation and control.

The EC 135 is equipped with two identical engines (Version: P1, P2,

T1, T2). They provide the driving power for the main and tail rotor andfor all secondary units.

The power plant system comprises the following components:

-- Engines

-- Engine indication

-- Fuel system

-- Power management-- Engine starting

-- Ignition

-- Oil cooling system

-- Engine mounts

-- Firewalls

-- Fire protection

-- Engine drain lines

Engines

In the EC 135 the following engine variants are possible:

-- T1 Turbomeca ARRIUS 2B, 2B1, 2B1A, 2B1A_1T2 Turbomeca ARRIUS 2B2

-- P1 Pratt&Whitney 206B,P2 Pratt&Whittney 206B2

NOTE The following desciptions refer to the versions T2

and P2, major deviations from the earlier

versionsT1/P1 are described in an overview page

at the end of the chapter.

Training manuals for these engines are published by Turbomeca orPratt&Whitney.

Engine Indication

The engine indicating system provides the pilot with information on theperformance parameters required during flight. The indicating systemalso provides information on engine malfunctions.

Fuel System

The aircraft fuel system provides fuel storage and supply to bothengines.

Power ManagementPower management and speed control of the engines is accomplishedby a Full Authority Digital Engine Control (FADEC) per engine. This isachieved with one single channel Electronic Engine Controller (EECof version P&W), resp. single channel Electronic Engine Control Unit(EECU of version TM). All control functions are monitored andimplemented when requested either by the electronic or by pilotsinputs. In the event of electronic failure or for training purposes control

is maintained by reverting to a manual back-up mode.






Power Plant -- General Arrangement

Oil Cooling System

Airframe Fuel System

Firewalls

Engine (T)Engine (P&W)






Fuel System

Fuel Storage System

General

The fuel storage system comprises two bladder tank cells mounted in

series in the lower shell of the fuselage. Access to the parts installedin the fuel cells (i.e. pumps, QTY--sensors, drain valves) is provided byfour removable equipment plates through the helicopter floor shell.The impact-resistant bladder tanks are made of reinforcedrubberiezed nylon fabric.

Main Tank

The main tank cell is located in the center bottom shell of the fuselage

between frame 3 and 5.

Threaded bolts, vulcanized into the topside of the tank, provide pointsfor attaching the tank to the underside of the cabin floor. Velcro strips,bonded to the topside of the tank and to the underside ofthe cabin floorprovide for additional stability.

Two fexible hoses are routed from the fuel pumps to the rear of thetank. They are connected to the split supply tank. The hoses

connected to the two sections of the supply tank are routed throughthe overflow channels located in the upper rear area.

Split Supply Tank Cell

The split supply tank is located in the bottom shell between frames 5and 8. To meet the certification requirements of total engine separationthe tank consists of two sections separated by a fuselage mountedfence. The fence has about the same height as two overflow channels

which connect the supply tank to the main tank. The tank isconstructed to accomodate the fence.

Two overflow channels, vulcanized into the front side of the tank cell,connect the supply tank to the main tank.

Flanges in the L/H and R/H side walls provide for engine supply linesconnection. One flange in the R/H sidewall allows the connection of anauxiliary fuel tank (option). One flange on the R/H rear top provides forconnection to the vent system.

Expansion Tank

The expansion tank (approx. 14.5 l, polyethylene material)

accommodates an inreasing fuel volume in case of a warm up and isfixed with a belt in an aluminium box to the R/H side of the cabin. Twoventing hoses are routed in vicinity of frame 5 upwards and along theunderside of the engine deck to the L/H side of the cabin, where theyare connected to the port vent outlets in the bottom shell.

Venting System

The refueling venting hoses of the main tank and the supply tank are

fixed to the LH inner topside of the fuel cells and routed into the fillerneck. In both hoses an air no fuel valve is integrated.

On the RHupper side the supplytankand the main tank are connectedto expansion tank by an additional vent line embedded into the cabinfloor.

From H/C SN 250 and up this vent line is separated between the twoexpansion tank inlets. In addition a short vent line between the main

tank and supply tank is installed.






Fuel System -- System Components

Ground Connection

Vent Hoses

Feed Line Engine 1

Cabin Floor Embedded Vent Line

Shut--off Valve Engine 1

Front Equipment Plate

Venting Hose

Flexible Fuel Hose

Air No Fuel Valve 1

Rear Equipment Plate

Hook--and--Pile Tape

Expansion Tank

Inlet for Aux. Fuel Tank (Option)

Shut-off Valve Engine 2

Filler Neck

Equipment Plate System 1

Equipment Plate System 2

Overflow Channnels

Split Wall

Feed Line Engine 2

Air No Fuel Valve 2 (not visible)

Vent Line (SN 250 and up)

Vent Line separated(SN 250 and up)



06yg y



Fuel Distribution System

General

The fuel distribution subsystem transfers fuel from the main tank intothe split supply tank and from there via shut off valves to the engines.

Components

The fuel distribution subsystem comprises the following components:

-- Two main tank mounted transfer pumps

-- Transfer lines

-- Two supply tank mounted prime pumps

-- Engine feed lines with shut-off valves

Transfer Pumps

The transfer pumps deliver fuel from the main tank to the split supplytank via transfer lines. The capacity of the transfer pumps is such thateach of them delivers more fuel to the supply tank than the engines canconsume. The surplus fuel returns to the main tank via the overflowtubes. This guarantees that the supply tank is always filled, as long asthere is fuel in the main tank. The pumps are powered with 28 V DCand have a dry operating time of approx. 20 minutes.

Transfer Lines

Two flexible hoses are routed from the transfer pump outlet ports to therear of the main tank. There they are connected to each other and tothe split supply tank. The connecting hoses to the two sections of thesupply tank are routed through two overflow channels in the upper reararea.

Prime Pumps

The prime pumps deliver fuel to the engines via the feed lines duringengine start. The pump in the left tank chamber supplies the left engineand the pump in the right tank chamber supplies the right engine. Withboth engines running, the engine driven pumps draw in the fuelthrough the prime pumps. Thus the prime pumps can then be switchedoff.

The prime pumps are identical to the main tank mounted transferpumps, but there is no check valve installed in the pump outlet.

NOTE The prime pumps are only to be switched on for

engine start and some emergency procedures acc.

to flight manual.

Foam Core

An airframe mounted foam core is integrated in the RH supply tankshape and therefore the fuel quantity is reduced by 4 kg in order toavoid a simultaneous flame out of both engines when the fuel tanksbecome empty.





Fuel Transfer System

FWD

Check Valve

Transfer Line

Overflow Channel

Shut--Off Valve

Adapter for Aux. Fuel Tank

to Engine 1

FWDTransfer Pump

Aft Transfer PumpCheck Valve

to Engine 2

Filler Neck

Prime Pump 2

Prime Pump 1

Chamber Divider

Foam Core

Main Tank 452 kg

LH Supply Tank 48 kg

RH supply Tank 44 kg

Rear View

Total Fuel 544 kg

up to SN 250 SN 250 and up

474.5 kg

49.0 kg

44.5 kg

568.0 kg





Equipment Plates

The four equipment plates in the fuel tanks are identical andaccomodate the fuel supply components. The components are:

-- Fuel pump

-- Fuel sensor

-- Low fuel sensor (supply tank only)-- Drain valve

The equipment plates are interconnected respectively with a groundcable to the fuselage.

Check Valve

The fuel pumps of the main tank are each equipped with a check valve

attached to the pump outlet port. The check valve prevents fuel fromflowing back to the main tank if a transfer pump should fail.

Fuel Sensors

The four capacitive fuel sensors, unequal in size, are attached to theequipment plates in an upright position.

NOTE Due to crash safety the fuel sensors must not

contact the ceiling of the fuel cell. To increase theaccuracy of the fuel indication when the system is

completely filled, from SN250 and up, all four fuel

sensors are longer. Therefore a cut out is

integrated in the cabin floor above each sensor.

Low Fuel Sensor

The fuel sensors in the supply tank are eqipped with low fuel sensors(NTC--thermistors) used for the LOW FUEL indication.





Equipment Plates

Drain Valve with Rubber Collar

Engine Supply Hose

Fuel Sensor

Prime PumpTransfer Pump

Check Valve

Equipment PlateMain Tank

LH Equipment Plate

Low Level Sensor

Transfer Hose

Foam Core





Fuel Pumps

The four fuel pumps are identical. They are powered by a 28 V DCmotor, supplied by the electrical system. The DC motor of a pump ismounted in a separate compartment. It can be removed from theequipment plate without defueling the tank.

Two transfer pumps are attached to the equipment plates of the main

tanks in an upright position.

One fuel pump (prime pump) is mounted inside each of the two supplytank sections. This configuration makes provision for individualdefueling of the sections.

The fuel pumps of the main tank are each equipped with a check valveattached to the pump outlet port.

The delivery rate of the earlier pump version is 6.6 l/min (manufacturerGlobe Motors) and for newer versions 12.5 l/min (manufacturerTestfuchs) with a pressure of approx. 1 bar.

Drain Valve

Two drain valves are located inside the main tank and two inside thesupply tank. They are attached to the equipment plates, which areinstalled at the lowest point in the tank. Access to the valves is given

from the underside of the fuselage through access doors.The drain valves are opened by depressing the valve body. Anintegrated valve spring automatically closes the drain valve after thevalve body is released.

NOTE Drain each tank into a container and check for

presents of water, until only fuel emerges.

WARNING If the helicopter is parked on a slope watermight be left in the tanks even after the tankshave been drained.





Fuel Pump and Drain Valve

Pump Housing

Motor PumpCartridge

Locking Pl

ug

Valve Lever

Open Position

Closed PositionDrain Valve

Drain Tool

FuelFuel





Fueling System

General

The filler neck of the helicopter is used to fill the main tank. The groundconnection provides static discharge after landing and during fuelingthe helicopter.

Filler NeckThe filler neck is located between frame 4 and 5 at the lower left endof the side panel. The access door can be locked by a key. The fillerneck is constructed for gravity fuelling with a max. rate of flow of 100liters per minute.

Air No Fuel Valve

The air in the tank, displaced by fueling is bled through two ventingports on thetopside of the tank as wellas through venting hosesroutedto the filler neck coming from the rear section of the tanks.

If the fuel reaches the level of the venting valves the opening is closedvia the ball mechanism and the fuel tanks can be filled up to maximum.The fuel coming from the front port of the main tank or the rear port of the supply tank causes the valves to open and drains back to themain/supply tank.

Ground Connection

The ground plate is located on the outside of the LH rear side panelof the fuselage. The ground bushing extends outside and is located tothe right of the access to the filler neck. The ground bushing isconnected to the fuselage by a flexible ground strap.

WARNING Connect the ground cable before fueling thehelicopter.





Power Supply and Monitoring of the Fuel Pumps

General

The switches and circuit breakers for the fuel pumps are located in theoverhead panel.

Switches and Circuit Breakers Main Tank

The following switches/circuit breakers for the main tank are installed:

-- Switch FUEL PUMP XFER--F

-- Switch FUEL PUMP XFER--A

-- Circuit Breaker XFER--F--Pump

-- Circuit Breaker XFER--A--Pump

Switches and Circuit Breakers Supply TankThe following switches/circuit breakers for the supply tank areinstalled:

-- Switch FUEL PUMP PRIME I

-- Switch FUEL PUMP PRIME II

-- Circuit breaker PRIME--P ENG I

-- Circuit breaker PRIME--P ENG II

Precision Resistors

The precision resistors (shunts) for the current measurement of thetransfer pumps are located on the backside of the overhead panel ina mounting unit.





Fuel Pumps -- Switches and Circuit Breakers

Circuit Breaker XFER--A--PUMP

Circuit Breaker PRIME--P ENG II

Switch FUEL PUMP XFER--A

Switch FUEL PUMP XFER--F

Switch FUEL PUMP PRIME II

Switch FUEL PUMP PRIME I

Circuit Breaker XFER--F--PUMP

Circuit Breaker PRIME--P ENG I

ENGI

ENGII

OFF

M AX





Power Supply Transfer Pumps

The transfer pumps are supplied via the following busbars:

-- FWD transfer pump with Essential busbar 1

-- Aft transfer pump with Shedding busbar 2

Power Supply Prime Pumps

The prime pumps are supplied via the following busbars:

-- Transfer pump engine 1 with Essential busbar 1

-- Transfer pump engine 2 with essential busbar 2

Monitoring

The electrical circuits of the transfer pumps are monitored. In case of a defective pump, a dry running pump, or a switched off pump cautionindication is displayed at the CDS/CPDS MISC field.The indications are:

-- F--PUMP AFT

-- F--PUMP FWD

The pumps are monitored via a shunt. When the power consumptionis higher than 5 Amps (blocked pump), or longer than 3 min lower than2 Amps (dry running pump), the caution will be triggered.

Indication

As long as the prime pumps are switched on, in the CDS/CPDSCaution panel SYS I and/or SYS I the following indication will bedisplayed:

-- PRIME PUMP





Fuel Supply Lines and Shut--Off Valves

General

The fuel is supplied to the engines by two hoses, both equipped withshut off valves.

Components

The fuel supply system consists of the following:

-- Engine feed lines

-- Shut--off valves

-- Switch EMER OFF SW I

-- Switch EMER OFF SW II

-- Circuit breaker FUEL --V ENG I

-- Circuit breaker FUEL--V ENG II

Engine Feed Lines

The flexible fuel hoses are connected to the ports on both sides of thesplit supply tank. From the ports they are routed through the rearbottom shell, along the LH and RH side panel of the fuselage and upto the engine deck. All hoses are of size DN 08 and made of spiralfabric-- reinforced teflon tubing. All fuel hoses routed through the

fuselage are protected with fuel resistant hoses (DN 32). Additionalventing minimizes fuel vapor collecting in the hose system.The fuelhoses located above the engine deck are adapted to fit the specificinstallation requirements of engine. They are made of metal.

Shut-Off Valves

The fuel shut--off valves are used to perform emergency shutdown of the engines and also for normal maintenance activities. The valves areoperated by a 28 VDC electrical motor.

The shut-off valves are installed in sealed housings in the L/H and R/Hside shell. The housings are vented to the ambient.

Power Supply

The shut--off valves are supplied by the following busbars:

-- Shut-off valve engine 1 with Essential busbar 1

-- Shut-off valve engine 2 with Essential busbar 2





Fuel Shut-Off Valves

Switch EMER OFF SW I

Switch EMER OFF SW II

Circuit BreakerFUEL--V ENG I Circuit Breaker

FUEL--V ENG II

Warning Unit

Overhead Console

Electrical Connector

Cover

Hou-sing

Lower Fuel Hose

Upper Fuel Hose

Clamping Nut

Frame 6

Fuel Shut--off Valve

Upper Port

Lower Port





Operation

The shut--off valves are controlled by the EMER OFF SWITCH I resp.EMER OFF SWITCH II, located in the Warning Unit. The switches areguarded push--to release switches (FIRE -- Buttons).

-- When the switches are released, the valves close.

-- When the switches are depressed, the valves are open.

Monitoring

The positions of the shut--off valves are monitored and displayed at theCDS/CPDS SYS I/SYS II and at the warning unit adjacent to theEMEROFF SWITCHES.

When the valves are open (normal position):

-- No indication

If an EMER OFF SWITCH is released, the following indications willappear:

-- ACTIVE (Warning Unit) will be ON continously.

-- FUEL VALVE on the CDS SYSI/II is displayed as long asthe valve is transient.

-- F VALVE CL on the CDS/CPDS SYSI/II is displayed when

the valve is closed.

NOTE When no FIRE Warning is evident, only the

shut--off valve will close when operating the EMER

OFF SWITCH.





Fuel Shut--Off Valves -- Function

FIREI

EMEROFFSW

ACTIVE

1

I

P P 1 0 E

FIREII

EMEROFFSW

ACTIVE

II

2

1

1

P P 2 0 E

PCB

M

PCB

M

MS 1 MS 2 MS 2 MS 1

PR P R

SYS I MISC SYS II

1

2

12

4

5

6

7

8

9

10

11

12

13

1 Warning Panel SYSTEM II CDS/CPDS2 Valve Actuator3 Limit Switch MS 1 (totally closed)4 Switch EMER OFF SW 25 Circuit Breaker FUEL--V ENG II6 Fuel Shut--Off Valve Position Indication7 Busbar PP 20E8 Warning Unit9 Busbar PP 10E10 Circuit Breaker FUEL--V ENG I11 Switch EMER OFF SW 112 Limit Switch MS 2 (totally open)13 Waning Panel SYSTEM I CDS/CPDS

1 Fire Warning Logic within Warning Unit

2 Instrument Lighting

3

FUEL VALVE

F VALVE CL

Fully ClosedClosing Opera-tion





Fuel Quantity Indication System

General

The fuel quantity indicating system provides the pilot with informationsabout the fuel quantity and system malfunctions. The relevant data isgathered with the aid of sensors, digitally processed and displayed onthe CDS/CPDS.

The fuel quantity indication system mainly consists of:

-- Four fuel transmitters

-- Indication on the field FUEL of the CDS/CPDS

Fuel Quantity Transmitter

Each fuel quantity transmitter consists of two concentric tubes,installed on the equipment plates in the fuel tank . The inner and outertube form the plates of a capacitor.

As the fuel level changes, the amount of fuel between the twocapacitor tubes changes. This changes the value of the dielectricum,thereby varying the capacity of the fuel quantity transmitter. Anoscillator circuit, consisting of a precision resistor and the transmitter,changes its frequency proportional to the fuel mass in the tank. Theoutput frequency varying between 8 kHz with a full tank and 13 kHz

with an empty tank is digitally processed and displayed in theCDS/CAD.

The measurement accuracy amounts to 6% with maximum fuelcontent and 4% with decreasing fuel content. Inaccuracies resultingfrom pitch-- attitudes of the helicopter are taken into account (attitudecompensation).

Inaccuracies resulting from different fuel types and temperatures

(density) are within the 6% resp. 4%.





Fuel Quantity Indication

ProcessingUnit

ProcessingUnit

Electronics

ProcessingUnit

Electronics

ProcessingUnit

Split Supply Tank

Main Tank

Fuel Sensor

Fuel Sensor

CDS

CPDS

Fuel Sensor





Fuel Quantity Indication CDS

General

The fuel quantity indication (FUEL display) is located in the CDS. It issubdivided in the following sections:

Numeric indications (white) for MAIN, SPLY 1, SPLY 2 and AUX (main

tank, supply tank 1 supply tank 2 and optional auxiliary tank).Bar indication (white) for MAIN, SPLY 1, SPLY 2 and AUX.

LOW warning indication (red) for SPLY 1 and SPLY 2. The LOWwarning indication for SPLY 1 and 2 is triggered by the CDS software,when the respective supply tank chamber indication is less than 28 kg.

FREE condition indication (white) for main tank. The FREE indicationcomes on when the free volume in the main tank and both supply tanks

is greater than the current fuel volume of the auxiliary tank (if installed).

XFER (white) if an auxiliary tank is installed. The XFER indicationcomes on when the auxiliary fuel tank valve is open.

Unit indication (kg/lb).

NOTE Through configuration the amount of fuel can be

displayed in kg or lb.

The numeric indications display as maximum values:

-- 448 kg for the main tank (nominal max. quantity 452 kg*)

-- 48 kg for SPLY 1

-- 44 kg for SPLY 2

* 4 kg can not be displayed, because for crash safety the fuel sensors

must not contact the fuel cell ceiling.

Fuel System Monitoring CDS

The caution FUEL QTY FAIL comes on if one supply tank sensor orboth main tank sensors fail. The respective graph will reset to 0.

The caution FUEL QTY DEGR comes on if one of the main tanksensors fails.





CDS Fuel Indication


FUEL PRESSFUEL FILTF FILT CT

F QTY FAILFQTY DEGR


Bargraph Indica-tion

Numeric Indication

Unit Kg/LB

LOW Indication Light

FREE Indication LightMain Tank

XFER IndicationLight Aux. Tank





Fuel Quantity Indication CPDS

The display indicates the fuel quantity in the main tank and in bothsupply tanks. In addition to the symbolic display of the fuel contents inthe tanks, a numerical display of the tank contents is provided in theselected unit of measurement.

Fuel Flow Indication and Endurance Calculation

If a flow sensors is installed in each engine supply line, the actual fuelconsumption and the calculated endurance is displayed in the CAD.

Fuel System Monitoring CPDS

The caution FUEL QTY FAIL comes on if one supply tank sensor orboth main tank sensors fail. The respective graph will reset to 0.

The caution FUEL QTY DEGR comes on if one of the main tanksensors fails.

CPDS software version V2002 and higher:

The caution FUEL comes on after 15 sec if the indication of supply tank1 is below 40 kg or the indication of supply tank 2 is below 35 kg.





Fuel System Display CPDS

CAUTION/ADVISORY Half Page

Possible Fuel Indications

Numeric Indication

Bargraph Indication


F QTY FAILFQTY DEGRFUEL


Possible CAUTION Indications





Low Level Warning CDS/CPDS

General

The low level warning is an additional fuel level control to warn thepilot.The warning function can be checked with a test function. A visual and audio warning informs the pilot that:

-- There are approx. 28 kg (SYS I) and 24 kg (SYS II) of fuelremaining in the supply tank chambers.From SN 250 and up the position of the sensors has beenchanged. Therefore approx. 32 kg (SYS I) and 28 kg (SYSII) of fuel remaining in the supply tank chambers.

NOTE All configurations guarantee a minimum of 8

minutes remaining flight time.

Components

The low level warning mainly consists of:

-- One low level sensor in each supply tank chamber.

-- Red LOW caption at the warning unit for FUEL (SYS I andSYS II)

-- Circuit breaker FUEL--L--SYS I / SYSII

-- Circuit test function

Function

The sensors are fixed at a defined height to the fuel level transmitters.They are supplied by 28 VDC. As long as the sensors are cooled bythe fuel, their resistance is high resulting in a low current flow in thecircuit. If the resistors become free (level low), they will be heated up

by the current thus changing their resistance. As the resistors are“ntc--thermistors”, the resistance becomes low by the increasingtemperature, so the current in the circuit increases and hence activatethe LOW warning at the warning panel.

At the same time an audio warning is given through the head--phones: A gong every 3 seconds.

Power Supply

The low level warning is supplied with 28 V DC by the followingbusbars:

-- Essential Busbar PP 10E

-- Essential Busbar PP 20E





Fuel Supply -- Low Level Warning

PP 10E

TEST TEST

PP 20E

Electronics

ProcessingUnit

Electronics

ProcessingUnit

Circuit BreakerFUEL--L SYS I

Busbar PP 10E

Warning Unit

Low Level IndicationLOW FUEL SYSTEM 1

LOW LEVEL IndicationSYSTEM 2

Warning IndicationSYSTEM 2

Circuit BreakerFUEL--L SYS II

Busbar PP 20E

Supply Tank

Warning IndicationSYSTEM 1 (internallytriggered in the CDS)


F QTY FAIL

F QTY DEGRLow FuelSensor





Fuel Low Pressure Caution

General

The fuel low pressure caution indicates low pressure between theengine driven low-- and high pressure pumps.

Fuel Pressure Switch TM

The fuel pressure switch is attached to the fuel control unit. Thepressure is tapped between the fuel filter outlet and the high pressurepump inlet.

Whenever the fuel pressure drops below 1.3 bar, the pressure switchcloses and activates the caution FUEL PRESS in the CDS/CPDS.

Fuel Pressure Switch P&W

If installed, the fuel pressure switch is attached to the fuel managementmodule. The pressure is tapped between the low pressure pump outletand the fuel filter inlet.

Whenever the fuel pressure drops below 0.6 bar, the pressure switchcloses and activates the caution FUEL PRESS in the CDS/CPDS.

Fuel Filter Contamination Caution

General

The fuel filter contamination caution detects clogged filter elements.The indication is given at the CDS/CPDS.

The differential pressure switch is attached to the fuel management

module (P&W), resp. to the fuel control unit (TM). Differential pressureis tapped between the fuel filter inlet and outlet (valid for both enginetypes). When the filter element becomes dirty, the pressure differenceincreases. The switch closes reaching the pressure switch setting andthe caution FUEL FILT comes on the CDS/CPDS.

Circuit Monitoring

General

The electrical circuit of the fuel filter is automatically tested. If there isan interruption the caution F FLT CT will be displayed on theCDS/CPDS.





Fuel Supply -- Monitoring


TM P&W

FUEL PRESSFUEL FILTF FLT CT

FUEL PRESSFUEL FILTF FLT CT

CDS/CPDS

Fuel Pressure Switch Filter DifferentialPressure Switch

CDS/CPDS Test Function forF FLT CT Caution

Interface Helicopter -- Engine

1

2

2





Engine Turbomeca ARRIUS

General Description

Purpose

The EC 135 T utilizes two ARRIUS turboshaft engines to supply

energy (torque, bleed air, electrical power) to the helicopter systems.

General

The ARRIUS is a lightweight, free turbine, turboshaft engineincorporating a single stage centrifugal compressor driven by a singlestage compressor turbine and a single stage power turbine that drivesthe reduction gearbox and aircraft powertrain.

Metered fuel from the Fuel Control Unit is sprayed into a reverse flowannular combustion chamber through twelve fuel nozzles (10 mainplus 2 start nozzles) mounted around the gas generator case.

A high voltage ignition unit and dual spark igniters are used to startcombustion.

A single channel, Full Authority Digital Engine Control Unit (FADEC)system with a mechanical backup FMM ensures accurate control of the engine output speed and fast response to changes in powerdemand.An electrically operated stepper motor located within the FuelControl Unit works in conjunction with the FADEC and changes fuelflow as required.

Configuration

The engine is of modular design. Mainly it consists of:

-- The reduction gearbox module

-- Gas generator and power turbine module

-- Engine Subsystems





Engine Turbomeca ARRIUS

REDUCTION GEARBOX GAS GENERATOR AND POWER TURBINE

Accessory GeartrainCompressor Turbine (N1)

Compressor (N1) Power Turbine (N2)

Reduction Gear Train

Output Shaft

Air Intake

Combustion Chamber

Exhaust

Oil Tank





Reduction Gearbox Module

Purpose

The reduction gearbox module reduces the power turbine speed (N2)to a suitable speed for the main transmission input. A second geartrainreduces the gas producer turbine speed (N1) to a suitablespeed to turnall engine accessories.

Configuration

The reduction gearbox consists of a front and rear light alloy casing.The lower part of the reduction gearbox forms the engine oil tank. Awall, located around the reduction gearbox rear casing, sperates thereduction gearbox module from the gas producer/power turbinemodule. The output shaft is inclined upward to suit the maintransmission installation.

Operation

The reduction gearbox has a two stage helical and bevel gear typereduction geartrain which changes power turbine speed to output shaftspeed. The engine output shaft assembly is attached to the secondstage reduction gear by internal splines.

The accessory drive geartrain provides the appropriate speed

reduction to turn all engine accessories, which are:

-- Starter/Generator

-- Low pressure and high pressure fuel pump

-- Oil pump

-- Permanent magnet alternator





Reduction Gearbox

Second Stage Gear (Drive Shaft)5898 RPM

First Stage Gear10616 RPM

N2 Input Gear44038 RPM

N1 Input Gear54117 RPM

Intermediate Gear with Breather23984 RPM

Starter/Generator Gear12334 RPM

Fuel Pump Gear + N1 Phonic Wheel11992 RPM

Oil Pump and Alternator Gear12334 RPM

Idler Gear





Gas Producer / Power Turbine

General

This module provides the mechanical energy required to drive theaccessory drive and the reduction geartrain.

Configuration

The gas producer/power turbine module mainly consists of a singlestage centrifugal compressor driven by a single stage compressorturbine and a single stage free spool power turbine.

Function

Air enters the engine through a radial inlet plenum chamber, formedby the compressor inlet case where it is directed rearward to the

centrifugal impeller. The accelerated air from the impeller passesthrough diffusor tubes which turn the air 90û and converts velocity intostatic pressure. This high pressure air surrounds the combustionchamber liner.

The combustion liner has perforations which allow the pressurized airto enter. The airflow changes direction 180û and is mixed with fuel fromtwo starter nozzles and 10 main nozzles. The fuel/air mixture is ignitedand the resultant expanding gases are directed to the turbines.

The expanding gases from the combustion chamber pass through thecompressor turbine stator vanes to the single stage compressorturbine causing the turbine to rotate which drives the compressor. Thestill expanding gases continue rearward to the power turbine statorand turbine. The exhaust gas from the power turbine is directedthrough an annular exhaust plenum to the atmosphere.





Engine ARRIUS -- Operation





Oil Subsytem

General

The oil system ensures lubrication and cooling of the engine. All thecomponents are installed on the engine exept the cooling unit.

Lubrication Requirements

Lubrication is required for the following components:

-- Reduction gear train and accessory drive train (gears andbearings)

-- Centrifugal compressor front bearing

-- Compressor turbine rear bearings

-- Power turbine front bearing

ConfigurationThe oil system consists of:

-- Integral oil tank

-- Pressure system

-- Scavenge system

-- Breather system

-- Indication

Oil Tank

The oil tank is integral with the engine. It is formed by the lower sumpof the reduction gearbox. On the R/H and L/H front side of the gearboxhousing filler necks are provided (depending on the engine installationthe not used filler neck is plugged, the other one is equipped with a filler

cap). The R/H and L/H side of the oil tank is provided with an oil levelsight glass (depending on the engine installation, one of the sightglasses will be visible). On the lowest point of the oiltank a drain plugis installed.

Pressure System

The pressure pump draws oil from the tank and delivers it underpressure to the system. A pressure relief valve limits maximum

pressure by returning oil to the pump inlet.

The oil is then delivered, through the filter and a calibrated orifice, tothe engine sections which require lubrication.The oil is sprayed by jets onto the parts to be lubricated. It also suppliesa squeeze film for the gas generator front bearing and the powerturbine bearing.

Scavenge System After lubrication, the oil falls by gravity to the bottom of the sumps. Theoil is then immediately drawn away by the scavenge pumps andreturned to the tank through the cooling unit.Strainers protect the scavenge pump against any particles which maybe held in the lubrication oil.





Oil System ARRIUS

SUPPLY

SCAVENGE

BREATHING

AIR VENT

Filter Bypass Valve

Oilfilter

Pressure and Temperature TransmitterLow Oil Pressure Switch

Electrical Magnetic Plug

Pressure Pump withRelief Valve

Scavenge Pump (Two Stages)

Aircraft Mounted AirCooler with TemperatureBypass Valve

Electrical Magnetic PlugOil Tank withSight Glass

CAUTION ENG OIL P

CAUTION ENG CHP

CAUTION ENG CHP

Strainer

Strainer





Engine Fuel Subsystem

General

The ARRIUS turboshaft engine is equipped with a single channel Full Authority Digital Engine Control (FADEC) system. This integratedpowerplant control system incorporates all control units for completeautomatic and manual control of the engine.

The engine fuel subsystem delivers metered fuel to the engine. It isautomatically controlled by the engine control subsystem. The enginecontrol subsystem provides the capability to override the Full AuthorityDigital Engine Control (FADEC) function to provide for manualoperation of the fuel subsystem.

Components

The fuel and -- control subsystem mainly consists of the followingcomponents:

-- Fuel pump unit

-- Fuel filter

-- Fuel Metering Unit (FMU)

-- Valve assembly

-- Injection system

-- Electronic Engine Control Unit (FADEC)-- Indicating

Several sensors and electrical harnesses as well as cockpit discretescomplete the control system.

Fuel Pump Unit

Fuel is delivered to the fuel pump unit by the aircraft fuel distributionsubsystem.

The fuel pump unit supplies fuel under determined conditions of pressure and flow. It is mounted on the front face of the reduction

gearbox and driven by the N1 geartrain. The unit consists of a lowpressure centrifugal pump, and a high pressure gear type pump. Apressure relief valve in the HP pump outlet opens in case of excessivepressure and reliefs fuel to the inlet inlet port of the HP--pump.

Fuel Filter

The filter retains any particles that may be in the fuel in order to protectthe metering unit components. It is mounted on the front face of thereduction gearbox. In the system, the filter is between the low pressurepump outlet and the high pressure pump inlet. In case of filter clogginga bypass valve opens and unfiltered fuel is supplied to the FMU. Thefilter is differential pressure monitored by an impending bypass switchfor cockpit indication as well as by a mechanical blockage indicator atthe filter housing. In case of a defect low pressure pump, a lowpressure switch will close for cockpit indication.





Fuel Metering Unit

The fuel metering unit is installed on the front face of the reductiongearbox. It is an hydromechanical unit which governs the fuel flowthrough the entire operational envelope of the engine. It uses EECUsignals or twist grip position as input parameters.

The FMU operates in two basic modes: the automatic mode, where the

required fuel flow is commanded by the EECU, and the manual modewhere the fuel flow is determined by the twist grip position. Pressurizedfuel from the fuel pump is routed to the Fuel Metering Valve and to thebypass valve which keeps a constant pressure differential across themetering valve.

-- Manual modeIn manual mode the metering valve is controlled by a inputlever, actuated by the collective lever mounted twist grip.

-- Automatic modeIn automatic mode the metering valve is controlled by astepper motor which is commanded by the EECU.

The FMU provides the following features:

-- Enables engine start and shutdown

-- Controls fuel flow as a function of power demand

-- Fail fixed with no power change during transition fromEECU mode to manual mode

-- Full power selection range available in manual mode aswell as in EECU mode

-- Limits the rate of acceleration/deceleration to preventengine surge/flame out during manual control mode





Fuel Metering Unit -- Basic Function ARRIUS (Simplified)

Stop--Valve

Fuel Outlet

Metering Valve

Fuel Inlet Fuel Return(to HP Pump Inlet)Stepper

Motor

Metering Valve

Positon Feedbackto EECU

Manual Inputfrom Twist Grip

Metering ValveControl Lever

Microswitch(Neutral Position)

ConstantnP Valve

Manual Control

ActuatorLP and HP

Pumps





Fuel Valve Assembly

The fuel valve assembly distributes metered fuel from the FCU to theinjection system. It is located on a support at the upper part of thecombustion chamber casing.

The valve assembly comprises the following valves:

-- Start electro-valve

The valve distributes fuel to the start injectors.-- Stop electro-valve

The valve controls fuel flow to the injection system ingeneral.

-- No preference injector valveThe valve closes fuel supply to the 9 main injectors duringa rapid fuel flow decrease.

Injection System

The injection system sprays fuel into the combustion chamber in orderto give stable and efficient combustion.

The injection system consists of 9 main injectors mounted around thecombustion chamber by two half-manifolds and 1 preference injector.

Forengine start two start injectors are additionally mounted at 1 o’clockand 9 o’clock position around the combustion chamber.

For engine starting only the start injectors deliver fuel to thecombustion chamber. At 50% N1, the start electro valve closes the fuelflow to the start injectors and opens the vent line to the outsideamosphere. Meanwhile the main injector valve is open and the maininjectors together with the preference injectors deliver fuel to thecombustion chamber.

If during normal operation the fuel flow is reduced significantly, themain injector valve closes but the preference injector still delivers fuelto the combustion chamber, in order to avoid an engine flame out.





Fuel Valve Assembly and Injectors ARRIUS

Left Half

Manifold

Right Half Manifold

PreferenceInjector

Fireproof Cover

Fuel Valve Assembly

Start Electro Valve

StopElectro Valve

Fuel Inlet

Main Injectors

Inlet from FMU

Stop Electro-Valve

Pressurizing Valve

Start Electro-Valve

Start Injectors

Manual Purge

Atmosphere

PreferenceInjector

Main Injectors

Main Injector Valve

Injection System

Injectors

Start Injector





Torque Indication

The torque measuring system measures the torque of the engineoutput shaft. An electromagnetic sensor with a confirmation box picksup the signal.

This signal is processed by the FADEC and sent to the analoginstrument and to the CDS or to the CPDS in thecockpit. The indicationis in % TQ.

The torque measuring system consists of two concentric shafts eachhaving a toothed wheel located at one end (phonic wheel) and a pulsepickup probe. The inner shaft (engine output shaft) is used to transmitengine torque and the outer acts as an unloaded reference shaft. Thetorsional deflection (twist) of the output shaft results in an angulardisplacement of the teeth between the loaded shaft and the referenceshaft.

The rotation of the phonic wheel formed by the teeth of each shaft, infront of the sensor produces a pulsed voltage in the sensor.

This voltage is sent to the FADEC which measures the displacementbetween the pulses and determines the engine torque for internal useand cockpit indication.

To compensate for material and manufacturing tolerances (no twoshafts will twist in the same manner) a torque conformation box is

installed on the engine. This box sends a trim value, which isdetermined on the test bench, to the FADEC. Since this value isspecific to a unique torque shaft, the trim module cannot be transferredto an other engine.

The pick--up is mounted in front of the reduction gear box near theoutput shaft.

The torque measuring system is supplied with power respective fromthe busbars PP10E / PP20 via the circuit breakers TRQ ENGI/II orCAD/VEMD ENG I/II. The sensors are supplied with power via therespective FADEC and circuit breakers.





Torque Indication ARRIUS

Pick--Up

ReferenceShaft

Output Shaft

PhonicWheel

Torque Indication

FADEC ConfirmationBox

CPDS (VEMD)

CDS Analog Instrument

CPDS (CAD)

Analog Torque Signal

Digital Torque Signal

Caution Backup Page





Gas Temperature Indication

This system provides an indication of the gas temperature at the gasgenerator turbine outlet. The gas temperature is an operatingparameter, particulary during engine starting for fuel flow control.

The four identically temperature sensors are located around the rearpart of the combustion chamber casing. Each sensor houses two

thermo elements.The four parallel--connected thermo elements supply their contactpotential to the indication system. A confirmation box allows acorrected temperature indication for a given turbine entry temperature.

The gas temperature indication is supplied with power respective fromthe busbars PP10E / PP20E via the circuit breakers TOT ENG I/II orVEMD ENG I/II.

CPDS Indication

The confirmed value appears in the FLI (Eng 1 via VEMD line 1 andEng 2 via VEMD lane 2). With one lane off the respective TOTindication in the FLI is lost. The non-confirmed TOT value is displayedin the SYSTEM STATUS page.

CDS Indication

The confirmed value is shown in the analog indicator. The digital valuecan be selected in the parameter field of the CDS.





Gas Temperature Indication ARRIUS

4 Double Sensors(Alumel/Chromel)at Position T4/5

Confirmation Box

FADEC

CDS Cockpit

CPDS Cockpit

T4T5

T4/5





Pratt & Whitney 206 B(2) Engine

General

The PW206B(2) is a lightweight, free turbine, turboshaft engineincorporating a single stage centrifugal compressor driven by a singlestage compressor turbine and a single stage power turbine that drivesthe reduction gearbox and aircraft power train.

Metered fuel from the Fuel Management Modul (FMM) is sprayed intoa reverse flow annular combustion chamber through twelve individualfuel nozzles mounted around the gas generator case.

A high voltage ignition unit and dual spark igniters are used to startcombustion.

A single channel, Full Authority Digital Engine Control Unit (FADEC)

system with a mechanical backup FMM ensures accurate control of the engine output speed and fast response to changes in powerdemand. An electrical torque motor located within the FMM works inconjunction with the Electronic Engine Control (EEC) and changes fuelflow as required.

Configuration

The engine is of modular design. Mainly it consists of:

-- The reduction gearbox module

-- Gas generator and power turbine module

-- Engine subsystems





Engine P&W 206(B)

REDUCTION GEARBOX TURBOMACHINERY

Accessory Geartrain

Compressor

Compressor Turbine (N1)

Output ShaftPower Turbine (N

2)





Reduction Gearbox Module

General

The reduction gearbox module reduces the power turbine speed (N2)to a suitable speed for themain transmission input. A second gear trainreduces the gas producer turbine speed (N1) to a suitablespeed to turnall engine accessories.

Configuration

The reduction gearbox consists of three machined aluminum casingswhich are the front and rear housings and the output shaft cover. Frontand rear housing are bolted together with the compressor inlet case.The rear face of the housing and the front face of the compressor inletcase form an integral oil tank. The output shaft cover supports theoutput shaft front bearings. It is bolted to the front housing.

The output shaft is inclined at 26û upward to suit the main transmissioninstallation.

Operation

The reduction gearbox has a two stage bevel gear type reduction geartrain which changes power turbine speed to output shaft speed. Theengine output shaft assembly is attached to the second stage

reduction gear by internal splines.The accessory drive geartrain provides the appropriate speedreduction to turn all engine accessories, which are:

-- Starter/generator

-- Permanent magnet alternator (PMA)

-- Fuel management module





Reduction Gearbox P&W

Starter/Generator Drive12,590 RPM

N2 Input Shaft39,130 RPM

Second Stage Gear5,928 RPM

First Stage Gear

Oil Pump Drive4,200 RPM

Fuel Pump Drive6,680 RPM

Permanent Magnetic Alternator24,667 RPM

N1 Input Shaft58,000 RPM





Gas Producer / Power Turbine

General

The turbomachinery module provides rotational drive to the reductiongearbox module.

Configuration

The turbomachinery module mainly consists of a single stagecentrifugal compressor driven by a single stage compressor turbineand a single stage free spool power turbine.

Function

Air enters the engine through a radial inlet plenum chamber, formedby the compressor inlet case where it is directed rearward to thecentrifugal impeller. The accelerated air from the impeller passes

through diffusor tubes which turn the air 90û and converts velocity intostatic pressure. This high pressure air surrounds the combustionchamber liner.

The combustion liner has perforations which allow the pressurized airto enter. The airflow changes direction 180û and is mixed with fuel from12 fuel nozzles. The fuel/air mixture is ignited and the resultantexpanding gases are directed to the turbines.

The expanding gases from the combustion chamber pass through thecompressor turbine stator vanes to the single stage compressorturbine causing the turbine to rotate which drives the compressor. Thestill expanding gases continue rearward to the power turbine statorand turbine. The exhaust gas from the power turbine is directedthrough an annular exhaust plenum to the atmosphere.





Engine P&W -- Operation





Oil Subsystem

Purpose

The engine oil system is a dry sump system. It supplies a flow of filteredoil to the engine in order to cool, lubricate, and clean the variouscomponents.

Configuration

The oil system consists of:

-- Integral oil tank

-- Pressure system

-- Scavenge system

-- Secondary air system

-- Indicating system

Oil Tank

The oil tank is integrated into the engine and is formed by the annularcavity created between the air inlet case and and the reductiongearbox rear case. A drain plug located at the bottom of the inlet casepermits drainage of the cavity. Oil level indication is provided by a sightglass.

Pressure SystemOil is drawn from the tank, through a protective screen, to the inlet of a gear type pressure pump. A cold start valve, located at the pressurepump outlet provides a safeguard against excessive pressure build updue to high oil viscosity at low temperatures (pressure above 13.5 bar(200 psi) is released to the gearbox).

A P3 operated shut off valve prevents oil supply to the lubricationpoints upon engine start up and shut down (below ¶ 40% N1). Thisensures that the most remote bearing cavities (No. 4 and No. 5) do notflood during motoring or rundown periods.

The main oil filter (not cleanable) traps particles picked up by the oil asit lubricates the engine components. The filter is equipped with abypass valve as a safeguard against filter blockage. An impendingbypass switch gives indication to the cockpit before the bypass valveopens.

A pressure regulating valve (not field adjustable) is used to set the oilsystem pressure to a predetermined value for a specified speed andoil temperature.

After passing a fuel heater, the pressurized oil is distributed to the

lubrication points in the gearbox and to the bearings No. 4 and No. 5.

Scavenge System

The scavenge system returns the oil to the gearbox. Approx. 80% of the used oil flows into the sump by gravity. Bearing No. 4 is scavengedby blowdown from lab seal bleed air. Bearing No. 5 is scavenged bya combination of a scavenge pump and blowdown. At high RPM’s anoil pump bypass valve opens allowing surplus oil to bypass thescavenge pump to the sump.

One scavenge pump stage draws the oil from the sump via a protectivescreen and a magnetic chip detector. The oil then flows through anairframe-mounted oil cooler before it is returned to the oil tank.





Oil System P&W

Pressurized Oil

Scavange Oil

Oil Filter

Shut--off Valve

Oil Pumps

Oil Pump Bypass

Magnetic Plug

Oil Tank

Strainer

Strainer

Air VentPressure RegulationValve

Oil Filter BlockageIndicator

Oil Cooler

Fuel Heater

Oil Press. Indication

Oil Temp. Indication

CAUTION ENG OIL P

CAUTION ENG CHIPCAUTION ENG O FILT





Engine Fuel Subsystem

General

The engine fuel subsystem delivers metered fuel to the engine. It isautomatically controlled by the engine control subsystem. The enginecontrol subsystem provides the capability to override the Full AuthorityDigital Engine Control (FADEC) function to provide for manualoperation of the fuel subsystem.

Components

The Fuel and -- control subsystem mainly consists of the followingcomponents:

-- Fuel Management Module (FMM)

-- Fuel manifold and nozzles

-- Fuel flow divider

-- Electronic Engine Control (EEC)

Fuel Management Module

The FMM is installed on the accessory gearbox of the engine. It is anelectro-mechanical unit which governs the fuel flow through the entireoperational envelope of the engine. It uses EEC signals, twist gripposition and ambient pressure as input parameters. The FMM has an

integral fuel pump which delivers high pressure fuel to the meteringportion of the unit.

Modes of Operation

The FMM operates in two basic modes:

-- the automatic mode, where the required fuel flow iscommanded by the EEC.

-- and the manual mode where the fuel flow is determined by

the twist grip position.

Pressurized fuel from the fuel pump is routed to the Fuel MeteringValve and to the bypass valve which keeps a contant pressuredifferential of 3.4 bar (50 psi) across the metering valve.

In manual mode the metering valve is controlled by a mechanical N 1

governor. The N1 governor setting is influenced by a input lever,actuated by the twist grip.

In automatic mode the metering valve is controlled by a torque motorwhich is commanded by the EEC.

The FMM provides the following features:

-- Enables engine start and shutdown

-- Controls fuel flow as a function of power demand

-- Fail fixed with limited power change during transition from

EEC mode to manual mode-- Full power selection range available in manual mode as

well as in EEC mode

-- Limits the rate of acceleration to prevent engine surgeduring manual control mode

-- Operates in speed govening modes





Engine Fuel System P&W

Fuel Tank with Prime Pump

Regenerative Fuel Pump

Impending Bypass Switch

By--pass Valve

Low Pressure Filter

Jet Pump Gear Pump

Drive from Engine Gearbox

Main Air Blast Fuel Nozzles

Hybrid Fuel Nozzles

Fuel Drain

Fuel Flow Divider

Fuel Metering Unit with Fuel Metering Valve

CAUTION FUEL FILT

CAUTION FUEL PRESS (if installed)





Fuel Pump

Fuel is delivered to the fuel pump by the aircraft fuel distributionsubsystem. The engine driven low pressure fuel pump takes fuel fromthe supply tank, increases the pressure and pumps the fuel throughthe fuel filter.

The filter is differential-pressure monitored for cockpit indication by animpending bypass switch. A bypass valve will open if the filter becomes

clogged.

An engine oil heated fuel heater is installed in the line between thepump outlet and the filter inlet. At fuel temperatures below 43ûC atemperature controlled bypass valve is closed and the fuel has to flowthrough the heater (At fuel temperatures above 57ûC the valve is fullyopen and the heater is by-passed.

In case of a defect low pressure pump a low pressure switch will close

for cockpit indication.

From the fuel filter, fuel is routed to the second stage of the fuel pump(high pressure stage) and delivered to the fuel metering module forflow control.

Fuel Flow Divider

The flow divider distributes the metered fuel flow from the FMM to theprimary and secondary side of the fuel manifold. During enginestart-up, the flow divider routes fuel flow to the primary nozzles only. As the engine accelerates, fuel pressure increases and the flow dividerroutes fuel to the secondary nozzles too.

The fuel flow divider also prevents of fuel accumulation in the

combustion chamber after engine shut-down. For this, residual fuel iskept in an accumulator. At the next start as the fuel pressure increases,the accumulator piston forces the fuel towards the manifold.

Fuel Manifold

The fuel manifold distributes primary (start) and secondary (main) fuelto the engine combustion chamber. The fuel manifold is located on theengine gas generator case and consists of one inlet fuel nozzle, sixsecondary fuel nozzles and five primary (hybrid) fuel nozzles. The fuelnozzles are equally spaced around the combustion chamber for evenfuel flow.

Primary fuel flow from the primary nozzles remains constant duringstart-up and engine operation.





Fuel Metering Module, Fuel Manifold

Fuel Filter

Torque Motor

Low Fuel Pressure Switch

Fuel Filter Impending Bypass Switch1 Air Blast Nozzle2 Hybrid Nozzle

1

2

1

21

2

1

2

Primary Fuel Manifold

Fuel Flow Driver

Rear View of Engine





Temperature Indication

Exhaust Gas Thermocouple (T6)

The exhaust gas temperature measuring system consists of twiidentical harnesses with four thermocouple elements in parallelconnected to an engine mounted terminal assembly. From the two junction boxes the signals are lead to the inlet temperature sensor box

where they are paralleled and integrated with the signal from the T1sensor. The T6 system provides an output signal which is proportionalto the arithmetic average of the of the exhaust temperatures to whichthe eight thermocouples are exposed.

Inlet Temperature Sensor (T1)

The T1 sensor incorporates a platinium resistance temperatureelement together with a cold junction for the T6 thermocouples. The

active portion of the sensor is located near the inlet to the compressorinlet scroll, therby giving a signal proportional to engine air inlettemperature (T1). The signals provided by the T1 and T6 system arelead to the FADEC where the measured gas temperature (MGT) iscomputed.

CPDS Indication

The digital TOT value is displayed on the FLI (Eng 1 via VEMD lane

1, Eng 2 via VEMD lane 2) In case of one lane off, the respectiveanalog back-up value is displayed via the CAD.

CDS Indication

The value is shown in the analog indicator. The digital value can beselected in the parameter field of the CDS.





Temperature Sensors P&W

2 Sensors (Alumel/Chromel Cold Junctions) and1 Temperature Element (Platinum) at Position T1

Terminal Box

Trim Box

FADECCDS Cockpit

8 Sensors(Alumel/Chromel)

at Position T6

T6T1

Analog Signal

Digital Signal





Engine Control T2/P2

General

The helicopter is equipped with an electronic engine control systemFADEC (Full Authority Digital Engine Control), that facilitatesautomatic control of both engines for all RPM and power ranges. Theengine power parameters of the EC 135 are optimized with the aid of

the electronic engine controls, i.e. engine power is adjusted tooptimally fit flight profile and/or maneuver while simultaneouslykeeping fuel consumption to a minimum.

In case of failure of the FADEC the pilot has the possibility of manualengine control.

System Components

The engine control system consists of:-- Electronic power control FADEC

-- Emergency engine control





Engine Control System ARRIUS





Engine Control System P&W





Electronic Power Control T2/P2

General

Electronic power control ensure automatic operation of all engine--related hydro--mechanical and electrical components. A FADEC box(Turbomeca call it EECU, Pratt&Whitney call it EEC) per engine servesas central processor. The digital control unit is mounted in thehelicopter and connected to the engines by wiring harnesses.





Functional Schematic -- Engine Control

Starting

--Ignition/Starter--Fuel injection--Acceleration--Idling speed control

(N2 /NRO)

Flight

--Rotor Speed Governing N2 /NRO

--Acceleretion/Deceleration viaN1 /Fluel Flow Regulation

--Overspeed Protection

--Training Mode--CAT A Mode--Topping Selection

Digital Engine Control System

TM (EECU) P&W (EEC)

Ground Operation

--Idling speed control--N2 /NRO

--Flight (Flat Pitch) SpeedControl N2 /NRO

CDS/CPDS

Manual EmergencyEngine Control(Twist Grip)

Indicator NRO





FADEC--Box

The digital control unit controls the fuel supply and monitors the wholeengine functions with sensors. The packages provide for ambientcondition sensing, signal conditioning and excitation for externalsensors, analog and frequency to digital conversion, and serial datatransmission and reception.

Location FADEC--Boxes

The FADEC--Boxes are mounted with angle brackets and vibrationdampers below the engine deck in the middle section of the fuselagebetween frame 5 and frame 6. They are positioned respectively to theleft and to the right in front of the engines.

There is a connection flange respectively for the control lines from thehelicopter to the engines.

FunctionThe digital control units FADEC for engine 1 and engine 2 provide thefollowing functions:

-- Automatic start-up of engines.

-- Fuel supply depending on N1 gas generator RPM duringstarting of engines as well as in ground idle and flight RPMrange (IDLE/FLIGHT).

-- Automatic engine control in all RPM and power ranges.

-- Monitoring of engine and power parameters.

-- Limitation of the fuel flow after topping parameters havebeen reached.





Engine Control -- Digital Control Unit FADEC

FWD

Plugs

FADEC (TM)

Engine Deck

ConnectorPlate

Wiring Harness to Engine

Identification Plate

Port for Connectionwith the Helicopter

Port for Connectionwith the Engine Mounting for

Ground Strap

P0--SensorInput

Bracket

FADEC (P&W)

Port for Connectionwith the Engine

Port for Connectionwith the Helicopter





Power Supply FADECPower supply to the FADECs is made by the helicopters 28 VDC--system and by engine mounted alternators.

During engine starting (up to 35--40 % N1) or if there is a failure of analternator the helicopter DC--system supplies the FADEC.

During normal operation the alternator (AC) supplies the FADEC, afailure of the helicopter DC--system has no influence to the function.

On ground with the engines not running power supply is made from theESSENTIAL busbars 1/2 (PP 10E and PP 20E) via the circuit breakersFADEC 1/2 and the switches FADEC 1/2 located on the engine controlpanel.

NOTE As long as the alternater delivers AC power the

FADEC remains operative even when the FADEC

switch in the engine control panel is in the offposition. In this case the FADEC is disconnected

from the H/C DC system only.

The caution REDUND comes on in the CDS/CPDS

(TM only).

LocationEngine TM:The alternator is mounted to the front of pump--filter support block. Apermanent magnet rotor with eight poles is mounted on the oil pumpdrife shaft.

Engine P&W:The alternator is an integral part of the reduction gearbox having its

rotor mounted directly onto the accessory drive gearshaft and thestator mounted into the reduction gearbox.





Alternator


Rotor with Internal Magnet8 Poles

Oil Pump Drive Shaft

Alternator Body

Alternator Body


Engine TMEngine P&W





Ignition System -- Location

Ignition Unit

Ignitor Plug (TM)

Circuit Breaker IGN ENG I

Circuit Breaker IGN ENG II

TM P&W





Control Unit (Overhead Panel)General

In order to control the engine over the complete operating range, theFADEC modulates the fuel flow for each particular operating condition.On the engine Mode selector panel, there are several operatingconditions selectable.

-- Switch NORM/MAN ENGI / ENG II

With the engine control switch in position NORM the automatic powermanagement is engaged. Switching into position MAN the engine canbe controlled manually with the twist grip.

-- Switch VENT/OFF/STARTMAN ENG I/II:

The starter/generator could be activated in the switch position STARTMAN for a manually starting of the engines. (Function inactive and not

certified).If the switch VENT/START MAN ENG I/II is set from OFF to VENT, thestarter/generator will be powered through the electrical master boxwhich controls the required operating voltage to the starter/generator.The starter/generator will begin to run up the gas generator assemblyto approx. 20% N1, while the starting relay remains de--energized andthe engine ignition system is deactivated. The fuel flow remainsshut--off.

-- Potentiometer N2 Adjust (Engine TM)

After installation of a new engine or a FADEC the N2 speed has to beadjusted by the potentiometer N2 ADJUST. The potentiometer isinstalled in the control panel ENGINE MODE SELECT in the overheadpanel.

-- Dip switch N2 Adjust (Engine P&W)

The dip switche N2 ADJUST is installed in the control panel ENGINEMODE SELECT in the overhead panel.

EC





Engine Mode Selector

Switch NORM/MAN ENG I / ENG II

Switch START MAN/OFF/VENT ENG I / ENG II

Dip Switch/PotentiometerN2 ADJUST

EC 135





Engine Control Panel -- Automatic Engine StartingGeneral

The FADEC controls the complete starting procedure including theincrease of RPMs, fuel flow and thereby the increase of the TOT. Thepilot has only to monitor the engine indicating system in order to abortthe start in case of malfunction.

The engine 1 starting cycle is described in the following. The engine2 is started in the same way.

Automatic Engine Starting

With the switch FADEC in ON position the electronic control is suppliedwith power. After the internal self test is passed the caution FADECFAIL on the CDS/CPDS disappears.

With the ENGINE CONTROL SWITCH I in position IDLE the starter,

the engine ignition system and the automatic regulation of the fuel flowis activated and the caution STARTER appears on the CDS/CPDS insystemI. At 50 % N1 the selfsustaining RPM is reached and the starteris switched automatically to the generator mode.

At the same time the cautions STARTER, ENG FAIL and GENDISCON disappear and the caution ENG IDLE comes on.

The N1 RPM continues the acceleration until the N2 /NRO reaches

approx. 70 %. This value will be regulated by the FADEC and is calledGROUND IDLE.

After a successful start of engine 2 both ENGINE CONTROLSWITCHES have to be set into the FLIGHT position. Thereby the N 1

of both engines accelerate until the N2 /NRO reaches approx. 98 % andthe cautions ENG IDLE disappear.

When the collective is raised and the helicopter takes off the N 2 /NRO

will increase automatically to 100% (Pitch Compensation).

Quick Start Procedure

The pilot may preselect both engines the same time with the ENGINECONTROL SWITCHES in FLIGHT position. The first engine

accelerates until the N2 /NRO reaches 98 %. When passing 50 % N1,the second engine will be activated automatically. Starting bothengines the same time is not possible.

Manual Start

The manual start is not certified and deactivated.

EC 135





Engine Starting

STARTERGEN DISCON

TRAIN SEL

OFFFADEC

ONFLIGHT ON

IDLE

OFF

ENG

OFFFADEC

ENG

OFF

IDLE

FLIGHT ARM

O

FF

ENG CONTROL

ENG FAIL

EC 135





Power Sharing of the Power Turbines N2The switch position FLIGHT provides power turbine speed at thenominal governing speed for normal flight (100 % N2). With densityaltitude between from 4000 ft to 9000 ft the N2 speed is automaticallyincreased between 100% and 104%. The electronic control governsand optimizes the performance of the engines and adjusts theperformance characteristic of both engines to each other. If there is aspeed or torque difference between the two engines, the pilot is ableto adjust the torque with the switch ENG TRIM. The adjustment iscontrolled by the engine indications. With a constant main rotor speed,there are two different trim operations possible:

-- Increase power engine 1 (L+) and decrease powerengine 2 (R--)

-- Increase power engine 2 (R+) and decrease powerengine 1 (L--)

The operation of the beep switch ENG TRIM is processed in theFADEC and routed to the N2 control unit of the engine. If the speed of one of the engines is increased a little, at the same time the speed of the other is decreased by the same value.

Droop Compensation

When there are control inputs from the collective lever (pitch) or from

the tail rotor control (yaw) there is an additional load to the engines.The result is a decrease of the speed of the power turbine N2

respective the main rotor speed (NRO). To maintain the N2 speedindependent from the engine load, the required change of load isadjusted automatically by the N2 control unit of the engine. This settingis realized by the FADEC, which takes input signals form the lineartransducers.

Linear Transducer

The linear transducer is a position sensor which transforms amechanical deflection into a electrical signal. Both the lineartransducers are located side by side to the right below the front cabinfloor. They are connected to the collective shaft. Each movement of theshaft is transfered to the inner guiding cylinder in form of a lift. The

operation of the tail rotor control moves a linkage to the outer guidingcylinder and causes a lift.

Crosstalk Capability

Due to the extended crosstalk capability between the FADEC boxesthe following features are available:

Detection of:

-- Manual Mode-- OEI Situation

-- Training Mode activation

-- Automatic Bleed Air Shut Off during OEI Situation

Automatic torque matching.

NOTE If an OEI situation is detected, the remaining

engine accellerates slightly to stabilize the rotorRPM according the AEO curve in the diagram

below. In the earlier versions T1/P1 the droopcompensation of the stopped or idling engine is

lost and the rotor RPM is regulated according the

OEI curve.

EC 135





Torque Trim -- Beep Switch ENG TRIM

Beep SwitchENG TRIM

Linkage

CollectiveLever

Collective Shaft

Switch Unit

View from Side

View from Top

Linear Transducer

Pedals Linear Transducer

Speed Control Loop

Pedals

Front Cabin Floor

R o t o r r p m

[ % ]

Density Altitude Z! [ft]

Variable Rotor Speed Automatically Controlled(Simplyfied Diagramm)

only P1/T1



EC 135



Training ManualPower Plant


Torque Limiter Concept

ToppingSelect Switch

Collective Grip

CAT A Switch

CAT A Switch

EC 135





Training Mode (Dual Engine) A training mode is implemented to perform realistic OEI training. Thistraining mode is based on a twin engine training concept featuring aso called TRAINING and a TRAINING IDLE engine. The training modeis designated in a way that the engine acceleration/deceleration andNRO governing mirror a real OEI situation.

The combined power of both engines in training mode will not exceed

the maximum power of the 30’’/2’ rating as long as the pilot operateswithin a “normal” NRO range. The load will be equally distributedbetween both engines.

The training function is pre-selectable. If only the training pre-selectionis activated, the engines stay in normal AEO mode. Training mode canonly be entered if training has been pre-selected and confirmed bysetting the main engine switch of one engine to position IDLE.

Training pre-selection is achieved by putting the training selectorswitch to ARM. Pre-selection of training is indicated by an advisoryTRAIN ARM on the CAD.

The TRAINING IDLE engine is chosen by switching the respectiveengine main selector to IDLE.

If the FADEC has successfully entered training mode, the cautionsTRAIN IDLE and TRAINING are indicated on the CAD to indicate thestatus of the TRAINING IDLE and the TRAINING engine, respectively.

NOTE All ENG EXCCED cautions triggered during thetraining mode situation will be deleted

automatically when the training mode is left.

Indication Variants

Engine TM: On the FLI, TRAIN will be indicated instead of TRAINING.

Engine P&W: On the FLI, TRAIN will be indicated instead of TRAINING and IDLE instead of TRAINING IDLE (FLI constraints).

Additionally a yellow inverted triangle appears next to the countdown

timer. Cockpit indication logic for the 30’’/2’ OEI indication in twinengine training mirrors the indication of the real 30’’/2’ OEI rating. Notethat while the FLI simulates OEI, the real engine parameters areavilable as digital values on both sides of the FLI gauge.

EC 135T i i M l





FLI Indication in Training (Example)

TRAIN TRAIN IDLE

T

IDLE

OEI LO

P2 Indication

53.3

730 720

100.1100.1

50.0

1 57

TRAIN ARM

TRAIN SEL

OFFFADEC

ONFLIGHT ON

IDLE

OFF

ENG

OFFFADEC

ENG

OFF

IDLE

FLIGHT ARM

OFF

ENG CONTROL

TRAINING TRAIN IDLE

Training Select Switch






Training Torque LimitationEngine TM: If the engine is operated in training and NRO is within thenormal range necessary for training, the torque topping is set at aconstant value of 128%. However, in case of excessive rotor speeddrop which could e.g. occur if the pilot would pull too much collectivepitch, torque is increased. Thethreshold forthe start of torqueincreaseis90%NRO. The increase in torque is 1.5% per 1% decrease in N2 andthus guarantees constant power.

Engine P&W: If the engine is operated in training and NRO is within thenormal range necessary for training, the torque topping is set at aconstant value of 128%. If the rotor speed drops below 92%, trainingis aborted:

-- training indications disappear (TRAIN IDLE becomes IDLEas long as the ENG CONTROL switch is in idle position)

-- FLI reverts to real AEO mode

-- switches have to be set back to normal position

Training idle engine in flight mode and train selector switch in OFFposition.

NOTE When the training mode is left due to RPM drop, for

safety reason the training idle engine does not

decellerate to a real idle speed although the

selector switch of the respective engine is in idleposition.






Overspeed ProtectionEngine ARRIUS

The threshold for the activation of the over-speed fuel shut-off is114%!1% N2. If the free turbine speed exceeds this value, the EECUcommands the shut-off valve to close which causes an engine shutdown.

Once the over-speed fuel shut-off device has been activated and one

of the engines has been shut down, activation of the over-speed fuelshut-off device of the second engine is inhibited to avoid twin engineshut down during flight. The over-speed event is stored in the FADEC.

Two three position switches (TEST--OFF--REARM) are installed in theoverhead panelof the H/C to trigger the test of theover-speed fuel shutoff device and rearm the system after the test.

Static Test

At power-up, the EECU performs a static test of parts of the overspeedprotection chain. If a failure is detected by the EECU, a caution OVSPappears on the CPDS. Putting the cockpit switch for the the overspeedprotection into its REARM position has no effect on this signal and thecaution OVSP will remain. As long as this failure does only effect theover-speed protection system, engine start will remain possible.

Dynamic TestThis test is initiated by pushing the cockpit switch for the overspeedfuel shut-off device into its TEST position. The test is performed on

ground after the CPDS and EECU auto-tests have been successfullycompleted. The EECU only accepts the test if N2 is lower than 25%.This avoids unintended engine shut down.

If the switch is pushed into TEST position, it simulates an N2 > 114%and therfore triggers an over-speed detection. If the test is successfullycompleted and the system works properly, the caution OVSP is

indicated on the CPDS. The pilot can then rearm the system andextinguish the caution by pushing the three position switch into theREARM position.

If, however, a failure is detected during the test, the signal for theOVSPcaution remains regardless of the position of the three position switch.Nevertheless, engine start is possible as long as the failure within theoverspeed protection system does not affect other systems andprohibits normal engine operation.

The correct function of the O/S inhibition is tested by first triggering theO/S test for one engine and then triggering the O/S test for the otherengine without rearming the system of the first engine. If the O/Sinhibition of the first engine is operational, there will be no O/Sindication for the second engine.

An OVSP caution is indicated on the CPDS if an actual over-speedevent has been detected, if a failure of the system has been detected

during normal operation or test or if the over-speed fuel shut-off hasbeen successfully tested.






Engine Pratt&WhitneyThe PW206B2 control system features an independant over-speedlimiting system to avoid overspeed of the engine output shaft. Thesystem is part of the inertial EEC and available as long as the EEC ispowered, i.e. in auto and in manual mode. The over-speed limitingsystem uses the data provided by the engine’s standard speedsensors. After detecting an over-speed event (N2 > 112.9%) thesystem reduces engine fuel flow to a minimum in a controlled manner.

As soon as the reset RPM is reached, the fuel flow returns to astandard fuel flow control.

The test of the overspeed protection is started using a single threeposition switch (OvspTestEng1--OFF--OvspTestEng2) which isinstalled in the overhead panel. As the pilot can monitor the correctfunction on his indicators for N2, no additional cockpit indication isinstalled.

The usage of the over-speed test function is inhibited by the controlsystem if N2 > 81.4%. The test procedure instructs the pilot to set theengine to IDLE and then to activate the over-speed test function. Thislowers the threshold of the over-speed protection to a value of 72% N2. As the nominal value for ground idle is 74% N2, the pilot isable tocheckthe correct function of the over-speed protection by observing an dropor oscillation of N2 on his cockpit indicators.



Power Plant


FADEC Malfunction Indications (Engine TM)

Manual Mode

Mode DatumFADEC Inputs

Metering Valve Position

FADEC Self Test Results

FailureMode

FADEC InputsN1 Trim

Control ModeN2 Demand

Control of “FAIL” Ind.

RS 422 Data Bus

Detection,Isolation,

Writing and Transmissionof a Fault Report

and Recording

Control of Stepper Motor

Freeze

REDUND

DEGRADE

FADEC FAIL

REDUND REDUND

DEGRADE

FADEC FAIL

REDUND

EC 135Training ManualP Pl t



Power Plant


Engine Emergency ControlGeneral

In case of failure of the electronic engine control the engine emergencycontrol provides manual control of the fuel metering valve. Theemergency control is designed to facilitate individual power control of each engine by the pilot.

A twist grip to control each engine is installed on the collective lever.

Engine shut-down can only be performed by the pilot. The electricalshut-off valve located in the engine is not automatically controlled bythe digital control unit.

Components

The engine emergency control mainly consists of:

-- Twist grip for engine 1 and 2

-- Two red push-buttons STOP MIN FUEL

-- Gear box at the collective lever

-- Flex ball control cables

-- Connection flex ball cables for dual flight control

-- Connecting mechanism

Twist Grip

There are two twist grips installed on the upper section of the collectivecontrol lever. The upper twist grip controls engine 1, the lower onecontrols engine 2. Both twist grips are coupled with torsion tubes,routed downwards inside the collective lever. To control the powermanually the twist grips can be rotated independently from the neutral

detent “N” to the “MAX” and “MIN” positions. The twist angle of theversion TM is +/--45°, of the version P&W is +/--55°.

Push Button STOP MIN FUEL

To prevent of an inadvertant engine shut down during manualoperation the twist grips are equipped with a red “STOP” button (Pilot’sside only). When pressed, it releases a travel stop at the “MIN”position, enabling the pilot to rotate the twist grip further in order toclose a shut off valve in the FMM. This cuts the fuel supply to therespective engine.

Gear Box

The gear box is bolted to the lower end of the collective control lever.It converts the circular motion of the torsion tubes inside the collective

control lever into a longitudinal movement for controlling the ballbearing controls. The gearbox also gives the NEUTRAL position of thetwist grips. The force necessary to turn the twist grips out of the neutralposition can be adjusted at the gear box.

Flex Ball Control Cables

To transmit the twist grip movement to the engine input levers, flex ballcontrol cables are installed. They are routed from the gear box

(collective lever) towards the R/H side panel. In the vicinity of frame 4they are routed upwards to the main transmission deck and to engine1 and engine 2. There they are connected to the emergency controlinput levers of the N1 fuel control unit.




Power Plant


Engine Emergency Control

Twist Grip ENG 1

Gear Boxat the Lower Endof Collective Lever

Twist Grip ENG 2

Push--Button STOP MIN FUEL

Push--ButtonSTOP MIN FUEL

Stop Ring for Neutral Position

Outer Torsion Shaft

Inner Torsion Shaft

Stop Ring forNeutral Position

Connections to Flexball Cables

Collective Stick




Power Plant


Connection Flex Ball CablesThe gearboxes at the collective levers are interconnected by flex ballcontrol cables. It is possible to control both engines by both the twistgrips.

Connecting Mechanism Flexball to Engine

The ball bearing control ends facing the engine are furnished with aconnecting mechanism, comprising of the following:

-- Bracket angle (different for TM and P&W)

-- Bracket hinge

-- Boot with guiding sleeves

-- Ball joint

Function

The function of emergency control for engine 1 is shown:To manually control power of engine 1 the operating mode selectorswitch NORM/MAN ENG I must be switched from the NORM positionto the MAN position. Thereby the actual position of the fuel meteringcable is frozen. The indication on the CDS/CPDS appears:

-- ENG MANUAL

From this moment on the pilot takes charge of the power control byhand directly with the metering valve (TM) or a mechanical backup N1governor (PW). Twisting the grip out of the NORM position in thedirection MIN or MAX a warning indication is displayed on theCDS/CPDS:

-- TWIST GRIP

If the engine is still operating in the NORM MODE the rotor RPM willbe governed by engine 2 and a change of the power setting at engine1 will result in a torque split only as long as engine 2 is able tocompensate the changes in power demand.

NOTE With both engines in MANUAL mode there is no

automatic power sharing (N2--power turbine) and

no automatic droop compensation possible. The

N1--/N2 speed and therby the rotor RPM must beregulated by manual control.

D TM: If there is unintentional movement of the twist grip withthe operating mode selector switch in position NORM thecaution TWIST GRIP comes up. In this case the electroniccontrol detects the unintended use of the twist grip (MixedMode) and compensates the pilot’s inputs over the entire

range of the stepper motor. After turning back the twist gripto the neutral position the engine runs again in theautomatic mode. There is no reset necessary.

D PW: Any movement of the twist grip out of the neutralposition results in an immediate switch over to the manualmode.The indication TWIST GRIP and MANUAL MODE on

CDS/CPDS comes up.For a reset back to the automatic mode the pilot has to turnthe twist grip slowly back to the neutral position.




Power Plant


Engine Emergency Control

Engine TM

FWD

Flexball Cable Feedthrough Pocket

Bracket AngleRH Engine

Bracket AngleLH Engine

Boot withGuiding Sleeves

Bracket Hinge

Ball Joint

Engine P&W

FWD




Power Plant


Major Differences between P2/T2 and P1/T1 Versions

1. For EC135 T1 and P1 the following features are not available:

-- Extended cross talk capability

-- Dual engine training mode

-- CAT--A mode

-- Topping threshold selection

-- 30’’ and 2’ OEI power (only 2.5’ OEI and transient torque(20 sec)

Limiting values depend on engine version:

P1: PW206B

T1: Arrius 2B1; 2B1A; 2B1A_1

2. PW Engine 206B (P1) only:

After the manual mode has been entered by turning the twist gripduring normal flight condition out of the neutral position, the pilot hasto turn back into the neutral position and to perform a reset at theengine mode selector switch in the overhead panel to return to thenorm mode.

3. If the training mode is installed in P1 or T1 (2B1 engine only)versions only a single engine mode is available:

The training engine will be topped at a lower level (AEO Power) andthe idle engine decouples completely (no power sharing) but idles witha high idle speed (92% instead of 70% N2). Thus the rotor RPM canbe recovered earlier in case of training engine failure.





Oil Cooling System

General

Both engines as well as the main transmission of the helicopter areequipped with internal, independent oil circuits. These ensurepermanent lubrication and cooling of highly stressed componentsunder all operating conditions. To keep the oil temperature within

limits, a oil cooling system is installed in the helicopter.Independant cooling circuits are availble for the:

-- LH engine

-- RH engine


Components

The oil cooling system consists of the following:

-- 2 cooling fans

-- 2 inlet airducts

-- 2 outlet airducts

-- 2 dual section oil coolers (engine / main transmission)

-- 2 thermal controlled bypass valves in the engine circuits

-- several hoses and connectorsCooling Fans

The cooling fans are mounted on the front side of the maintransmission RH and LH. They are driven by the main transmissiongeartrain. (12665 RPM at 100%)

Oil Cooler

The oil coolers are mounted to the RH and LH side of the maintransmission. They are split into two sections. The smaller section of each cooler, which is connected to the main transmission by bushingsdirectly, serves for cooling the main transmission oil (50% each side).

The larger section of each cooler is connected to the associatedengine by oil hoses. This section serves for cooling the engine oil.

For optimizing cold-start characteristics a thermal controlled bypassvalve is installed in each engine oil cooling circuit.

At temperatures below approx. 85 ûC the bypass valve is open andallows the oil to bypass the oil cooler.

Cooling Air Flow Ambient air which enters the air intakes is drawn by the cooling fansand forced through the oil coolers via the inlet air ducts. From there theair is directed overboard by the outlet ducts.





Oil Cooling System -- Functional Scheme

Oil Cooler with Fan

Engine 1 withSensors (TM)

Temperature BypassValve System

Temperature BypassValve System

Oil Temperature IndicationOil Pressure Indication(CDS, TM)

Main Transmission

Front Firewalls Engine 2 withSensors (P&W)

Oil Cooler with Fan


Oil Temperature IndicationOil Pressure Indication(CDS, P&W)

ENG OIL PENG CHIPENG O FILT

ENG OIL PENG CHIPENG O FILT

VEMD IndicationCDS/CPDS





Oil Cooling System -- General ArrangementHose Arrangementto Engine TM

Thermal Bypass Valve

Firewall

Inlet Duct

Outlet Duct

Oil Cooler

To/FromEngine

To/From MainTransmission

FWD

Inspection Door

Hose Arrangementto Engine P&W





Engine Mounts

General

The Engine mounts attach each engine to the helicopter structure.Each engine has an inboard and outboard forward mount thatattaches the engine reduction gearbox to the fuselage fittings and arear mount strut that attaches the turbine section to a fuselage fitting.They are designed to retain the engines in the event of a crash landingdownward with a load factor of 20 g and forward with a load factor of 16 g.

The engines are installed on the engine deck behind the maintransmission. They are tilted at a V--angle of approximately 9° to thelongitudinal axis of the helicopter and of 2° to the horizontal axis.

Inboard Mounts

On the inboard side the engines are mounted by bearing blocks viaspherical bearings. On each side the spherical bearing are attachedto a mounting bracket on the tail boom mounting cone.

Outboard Mounts

On the outboard side the engines are attached to “V”--shaped lateralstruts, bolted to the engine compartment floor. A spherical bearing isinstalled on the upper side of the “V”--strut.

NOTE The inboard and outboard mounting points form an

axis, around which the engine can tilt.

Rear Mount Strut

The rear mount is the third engine attachment point. The strut isadjustable on the lower end for engine alignment (Engine output shaftmust be aligned to the main transmission input shaft).





Engine Mounts TM

Outer MountingBlock

Airframe Bracket

FWD

FWD

Inboard Mount

SphericalBearing

Rear Mount Strut

SphericalBearing

Rear Fitting on Engine

V--Strut

SphericalBearing





Engine Alignment

To ensure that the engines are properly aligned with the maintransmission input flange, an engine alignment check is required,whenever:

-- Replacement of an engine.

-- Replacement or adjustment of the rear mounting strut.

Engine alignment is performed with the alignment fixture. The

alignment fixture is installed between the engine stub shaft flange andthe transmission input flange, substituting the transmission shaft.

NOTE An alignment of the engine is not necessary if thesame engine is installed and the lengh of the rear

strut remains unchanged.

Installation of the Alignment Device

For Turbomeca engine an adapter has to be set on output flange of theengine stub shaft.

The mandrel of the alignment device must be shifted back, after thatthe device must be attached to the output flange of the engine stubshaft. Now the centering disk is screwed to the flange of freewheelshaft. The connecting flange of freewheel shaft must be rotated untilthe marking line on the centering disk is in horizontal position. The

mandrel must be extended until the tip of it almost contacts thecentering disk. If the tip exactly points to the marking line, the engineis correctly aligned to the main gearbox.

If a deviation in downward or upward direction is evident, thealignmentmust be corrected by adjusting the Z--strut.




Firewalls

General

To prevent fire from spreading, in the event that one of the enginesstarts burning, the firewalls constitute a complete fire resistant cellaround each engine. The firewalls are made of titanium because of itshigh melting point by low weight.

Configuration

The firewalls are divided into several subassemblies:

-- Airframe fixed firewalls

-- Engine fixed firewalls

To provide minimum effort during maintenance, certain parts of thefirewalls are installed by camlock fasteners. All edges facing to the

engine cowlings are provided with fire resistant seals.

Airframe Fixed Firewalls

The airframe fixed firewalls are divided into several subassemblies:

-- Foreward firewall assembly

-- Aft firewall assembly

-- Exhaust ejectors

Foreward Firewall Assembly

The fwd firewall assembly separates the engine compartment fromthetransmission compartment. It is designed with several holes, throughwhich the drive shaft, the engine oil lines as well as the engineemergency control cables are routed. It houses also the generatorcooling air inlet.

The inner sheets of the forward firewalls isolate the engine air intakezone and separate the engine compartments from each other.





FWD Firewall Assembly

Emergency Control Cable Bellows

Fixed Sheet Frame 5

LH Outer Sheet

Bellow

Inner Sheet

Air Inlet Sheet

Cover Plate

Air Wall

Drive Shaft Fairing





Aft Firewall Assembly

The aft firewall assembly separates the engine compartments from theequipment deck. It is designed with several holes, through which theexhaust gas ducts and the tail rotor drive is routed.

Exhaust Ejectors

The exhaust gases from each engine are routed rearward andoverboard through exhaust ejector tubes.

Air from the engine compartments is drawn by the exhaust gasesentering the ejector tubes. This serves for engine compartmentventilation and engine hot section cooling. Additional the engine noiseis muffled by this.





Aft Firewall Assembly

Lower Firewall Sheet

Upper Firewall Sheet

Stiffening Angle

Center Stiffening Sheet Assy

FWD

Ejector





Engine Fixed Firewalls

The engine fixed firewalls form a fire and debris protection seal for theair inlet plenum.

The firewalls are made of titanium sheet material. They are bolted tothe engine. An access door is provided for engine compressorinspection.

NOTE The general arrangement of the engine fixed

firewalls is identical with the P and T versions, butthe parts are of different design.

NOTE After reinstallation of the fire walls, check all bolts

and camlock fasteners are fixed and tightened.


E i Fi d Fi ll




Engine Fixed Firewalls

Side FirewallSheet

Upper FirewallSheet

AccessDoor

Lower FirewallSheet

Engine FirewallSheet

FWD

Engine Firewall Sheets

Engine TM

EngineP&W


Fi W i S t




Fire Warning System

General

Each engine is equipped with its own independantfire warning system.The systems consist of 2 thermal switches per engine, installed in thedesignated fire zones, and visual and audio warning devices in thecockpit. System function can be checked for continuity by test switchesin the overhead panel.

Components

The fire warning system consists of the following:

-- 2 fire detectors per engine

-- 2 fire warning lights (combined with the EMER OFF SW I/II)

-- Test switches FIRE E/W 1, FIRE E/W 2

-- Circuit breaker FIRE--D ENG I, FIRE--D ENG 2Locations

The fire detectors are located beneath the starter--generator andbeneath the combustion chamber casing.

Trigger Temperatures

TM P&W

Reduction Gearbox 210°C 204°C

Power Turbine 271°C 260°C


Fire Detectors Locations




Fire Detectors -- Locations

Fire Detector

FWDFWD

Engine TMEngine P&W


Function




Function

The following describes the functioning of the no.1 engine fire warningsystem. The no. 2 engine fire warning system functions in the sameway.

The test switch FIRE E/W 1 is set to OFF and circuit breaker FIRE--DENG I is depressed.

The no.1 engine electrical fire warning logic circuitry located in the

warning unit is supplied with 28 VDC power from theESSENTIAL--Busbar PP 10E.

If overheating is detected on the engine, the respective fire detectorcompletes the circuit to ground via test switch (OFF--Position) and thefire warning logic circuitry. The warning caption

-- FIRE I

on the pushbutton indicator EMER OFF SW 1 in the warning unit

illuminates. At the same time, the circuit to the AUDIO control unit is completed andan alarm bell sounds in the pilot’s headsets.


Fire Warning System




Fire Warning System

Engine 2Engine 1 Test Switch FIRE E/W 1/2

Indicator FIRE II

TEST

1 FIRE E/W 2

EXT

EXT / WARN WARNUNIT

OFF CDS SYS 1

HYDSYS 2

NORM

Indicator FIRE I

Fire Detector Fire Detector

Firewarning Bell




Two conditions are necessary to activate the fire extinguishing system: System Test




-- Fire warning caption FIRE on (signal from fire detector)

-- N1 of the respective engine < 50 %

When the two conditions are fulfilled a switch controlled by the N1 RPMcontrol unit will be closed, causing activation of the fire extinguishingbottle by means of a explosive cartridge separately for L/H or R/H side.The extinguishing agent will be released to the respective enginecompartment via tubes and nozzles.

As a result FIRE EXT caution will illuminate on CDS/CPDS cautiondisplay SYS I/II to inform the crew that the fire extinguisher was usedand the bottle is empty.

If one of the conditions is not fulfilled only the fuel shut--off valve closeswhen the switch FIRE is released.

y

Two switches installed in the overhead panel allow to test the firewarning system as well as the fire extinguisher system forserviceability. The switches are 3--position toggle switches.

The following positions and functions are availble:

-- OFF:No test function, fire warning and extinguisher system isarmed

-- EXT:Fire extinguisher system will be tested. CDS/CPDS cautionFIRE EXT will come on together with MASTER CAUTION

-- EXT/WARNCDS/CPDS caution FIRE E TST comes on, FIRE EXTremains on. Additionally the fire warning circuit will betested. Respective FIRE caption will come on together with

the audio warning BELL.

The switches are spring loaded between the positions EXT andEXT/WARN. They must be switched back to the OFF positionmanually.

NOTE The weight of the bottle must be checked every 12

month.




Engine Drain Lines




General

The engine drain lines ensure the necessary draining and disposal of minor fuel and lubricant leakage from the respective engine. Additionally the amount of sampled liquids in the drain bottles is forleakage detection of the system.

ComponentsThe engine drain lines comprises the following components:

-- Drain line -- fuel pump seal

-- Drain line -- combusting chamber

-- Drain line -- output shaft sealing (TM)

-- Drain line -- starter/generator output drive (P&W)

-- Drain line -- fuel starting injectors


Engine Drain Lines (Engine TM)




FWD

Drain Bottle

Fuel Pump Drain Line

Drain LineFuel Starting Injectors

Drain LineOutput Shaft Sealing

Drain Line

CombustingChamber

Drain LineOutput ShaftSealing




Engine Drain Lines (Engine P&W)




Fuel Pump Seal Drain Line

A

A

Starter/GeneratorOutput Drive Drain Line

Drain Bottle

CombustionChamberDrain Tube


Fuselage Drain Lines

f f f




The fuselage drain lines are made of transparent hoses leading fromthe engine deck in the rear right and left side shell downward. Theoutlets of the hoses are located respective in the left and right rearlower shell.


Fuselage Drain Lines




FWD

EC 135Pilot’s ManualStandard Equipment

Table of Contents

QUIT



Table of Contents

Standard Equipment Chapter P07

Not Applicable for pilotstraining manual 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EC 135Pilot’s ManualStandard Equipment

QUIT



STANDARD EQUIPMENT

NOT APPLICABLE FOR PILOTSTRAINING MANUAL

EC 135Pilot’s ManualOptional Equipment

Table of Contents

QUIT



Table of Contents

Optional Equipment Chapter P08

Informations aboutoptionals to betaken from special documents and added here 2. . . . . . . . . .

EC 135Pilot’s ManualOptional Equipment

OPTIONAL EQUIPMENT

QUIT



OPTIONAL EQUIPMENT

INFORMATIONS ABOUT OPTIONALS TO BE TAKENFROM SPECIAL DOCUMENTS AND ADDED HERE

EC 135Training ManualElectrical System




Electrical System


Table of Contents




Electrical Power Supply 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

DC Power Generation 6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Starter/Generator 10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Electrical Master Box 12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Battery System 18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Switches GEN I, GEN II, BAT MSTR 26. . . . . . . . . . . . . . . . . . . .

CDS/CPDS Indication 26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Warning Unit Indication 26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

External Power Receptacle 28. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

DC Power Distribution 32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Overhead Panel 32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Circuit Breaker Console 1 and 2 36. . . . . . . . . . . . . . . . . . . . . . . .

Function -- Complete System 38. . . . . . . . . . . . . . . . . . . . . . . . . . .

Operation with Battery (Emergency Operation) 38. . . . . . . . . . .

Automatic Engine Starting 40. . . . . . . . . . . . . . . . . . . . . . . . . . . . .Operation with One Generator 42. . . . . . . . . . . . . . . . . . . . . . . . .

Operation with Generators Connected in Parallel 44. . . . . . . . .

Operation with Separated Generators 46. . . . . . . . . . . . . . . . . . .

Operation with External Power Unit 48. . . . . . . . . . . . . . . . . . . . .

Connection of Shedding Busbar 1 and 2 50. . . . . . . . . . . . . . . . .

Fault Reactions 52. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

AC Power System 54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .







Electrical Power Supply

G




General

The electrical power supply systems generate and distribute power foroperation and control of the helicopter systems. The EC 135 electricalsystems operate on 28 V DC, when supplied by the battery, theyoperate on 24 V.

An AC system is installed additionally.

Components

The electrical power supply consists of:

-- Power generation

-- External power receptacle

-- Power distribution

-- AC power systemPower Generation

The power generation consists of two generators, a battery and thecorresponding master boxes.

External Power Receptacle

It is possible to supply the electrical power system with DC power by

an external power unit. The voltage of the EPU operates between 24and 28 V DC. The voltage of the EPU must be higher than the voltageof the battery (UEPU > UBATT).

Power Distribution

The power distribution consists of the following components:

-- Two master boxes

-- Battery master box

-- Two circuit breaker panels

-- Overhead panel-- DC receptacle

-- Terminal junctions

Several busbars are installed in the master boxes, the overhead paneland both circuit breaker panels, to which all electrical consumers of thehelicopter are connected by means of circuit breakers.

AC Power SystemThe AC power system generates two different AC voltages (26 V AC,115 V AC) out of 28 V DC. The AC voltages are distributed to theconsumers (navigation instruments) via modules and busbars.


Electrical Power Supply -- Locations

Generator 1




Generator 1FADEC 1

Overhead Panel

Electrical Master Box 1

Inverter II

InstrumentConsole

Battery

BatteryMaster Box

EPU ReceptacleHigh Load Bus 2

High Load Bus 1

Generator 2

FADEC 2

El. Master Box 2

2 3 4 4a 5 6 7 81Frame


DC Power Generation

General Generator System 1/2




General

The DC power generation supplies direct current by means of two DCgenerators and a battery.

Components

The power generation consists of:

-- Starter/generator, engine 1/2, with temperature switch andelectrical master box 1/2

-- Battery with temperature switch, battery master box andfuse

-- Switches (GEN I, GEN II, BAT MSTR)

-- Central Panel Display System (CPDS)

-- Warning unit

-- Bonding system

Generator System 1/2

The generator system 1/2 consists of the following components:

-- Starter/generator 1/2

-- Electric master box 1/2

-- PRIMARY busbar 1/2

-- SHEDDING busbar 1/2-- Fuses

-- Relays

Battery System

The battery system comprises the following components:

-- Battery with temperature sensor


-- Battery busbar

-- Fuses

-- ESSENTIAL BUS relay

-- Battery relay


DC Power Generation




PP10E PP20EPP10S

SBC1

GC1SC1

HLC1 BTC1

PP20H

BATC

EBC2EBC1

SC2GC2

SBC2

HLC2HPC2

GPUC

PP20SPP10H

Battery

ElectricalMaster Box 1

ElectricalMaster Box 2

PRIMARYBusbar 1

Receptecalfor EPU

BATTERY--Busbar

PRIMARYBusbar 2

Battery Master Box

Generator1

Generator2

HPC1

F

BTC2

F

Abbreviations:BATC Battery ContactorBTC Bus Tie ContactorEBC Essential Bus ContactorGC Generator Contactor

HLC High Load Bus ContactorHPC High Power ContactorSBC Shedding Bus ContactorSC Starter ContactorGPUC Ground Power Unit ContactorF High Powert Consumer

SHED. Bus

HIGH L. BusESS. BUS


General Description of the DC Power Supply

The battery is connected via the relays BATC, BTC1 and BTC2 to thePRIMARY busbar

For the pilot there are three switches on the switch unit of the

instrument console:The switch BAT/MSTR in position ON closes the relay BATC and the




PRIMARY busbar.

The SHEDDING busbars are connected via two relays SBC1 andSBC2 to the PRIMARY busbar.

Both the generators G1 and G2 are connected via two relays GC1 andGC2 parallel to the PRIMARY busbar.

The EPU supplies the PRIMARY busbar via the relay GPUC and thetwo relays BTC1 and BTC2.

When the EPU is connected to the helicopter’s electrical system (BATMASTER SW in position ON), both the relays BATC, GC1 and GC2 areopened. By means of this automatically function the generators areinsulated from the EPU. The relays SBC1 and SBC2 are automaticallyclosed in the following configuration:

-- Power supply with an EPU

-- Power supply with an active generator

The switch BAT/MSTR in position ON closes the relay BATC and therelays EBC1 and EBC2. Aditionally the relays BTC1 and BTC2 areclosed, if the switch BUS TIE is in position NORM.

The push button position RES engages the relay BATC after a failureagain, if the coupling conditions are not fullfilled.

The two switches GEN I and GEN II with their positions ON/OFF/RES

closes the two relays GC1 and GC2. The push button position RES isused for engaging again a disconnected generator after a failure.


Leading Particulars DC Power Supply

Engine EC 135 T EC 135 P

G t i ht 7 60 k 9 5 k

Leading Particulars DC Power Distribution

Voltage range 26 -- 30 V

P i t f l ti POR 28 V DC"0 1 V




Generator weight 7.60 kg 9.5 kg

Nominal data 30 V DC, 160 A 30 V DC, 200 A

Speed range 8,400--12,100 RPM 7,050--12,000 RPM

Max. speed (5 min) 14,000 RPM 14,000 RPM

Temperature switch 205!5.5 °C 205!5.5 °C

Number of Batterys 1

Voltage 24 V

Capacity 17Ah, 25 Ah, 26 Ah, 40 Ah

Assembly 20 cells series connected

Temperature switch 70!3 °C

Point of regulation POR 28 V DC"0.1 V

Fuses

--in the masterboxes Blowout fuses 50 A, 80, 100A(150A)

--in the overhead panel Circuit breakers, different values

Total weight of the masterboxes

17 kg

Leading Particulars AC Power Supply

Number of systems 1 or 2 optional

Input voltage 24--28 V DC

Output voltage and power 26 V AC, 400 Hz, 150 VA115 V AC, 400 Hz, 350 VA

Max. current 15 A DC input


Starter/Generator

General

Th DC ti b t i d i d t l l t i l

Generator Mode

I th t d th t t / t li th l t i l




The DC power generation subsystem is designed to supply electricalenergy from several sources. Depending on the operation mode allthree master boxes determine the source from which the energy canbe taken and which busbars are supplied. For this the three masterboxes are connected to each other.

Starter/Generator

The starter/generator can be used in two modes:

-- Starter mode

-- Generator mode

Starter Mode

In the starter mode the starter/generator is used to start the engines.

The starter input is supplied with current by means of an externalpower unit or the installed battery. The starter drives the engine gasgenerator assembly by means of the drive shaft.

In the generator mode the starter/generator supplies the electricalsystem and loads the battery.

The generator mode is only available when the engine is running, asthe armature is driven by n1 geartrain of the engine.

In the generator mode a magnetic field is built up in the armature via

the exication input. If the armature is driven, voltage is induced. Thebrushes collect the induced voltage from the collector coil and transmitthis voltage to the connectors of the generator. A compensating coil isconnected in series to the armature to compensate for arcing. A fancools the generator during operation.

In the generator mode the generator supplies the PRIMARY busbar inthe associated master box with current.

The engines are equipped with the following starter/generators:

-- Engine T 160 A (200 A optional)

-- Engine P 200 A (160 A optional)

Temperature Switch

The temperature switch monitors the temperature of thestarter/generator cooling air and opens contact when the temperatureis higher than approx. 205 °C.

The caution GEN OVHT is displayed on the CAD.


Starter/Generator

Exiting Winding D (Balance)

T t S it h




CompensationWinding

Starter Winding

E-- (Ground)

(Starter Input) C+

GeneratorOutput B+

(Excitation Input) A+

Temperature Switch

Engine T

Fan

Commutator

Brush Holder

Exciting Winding

Armature

Drive Shaft


Electrical Master Box

Electrical Master Box System 1

The electrical master box controls the function of the DC system 1 and

Location

The electrical master box 1 is installed behind the LH interior paneling




The electrical master box controls the function of the DC system 1 andregulates the voltage of the system to 28 " 0.1 V.

The current supplied by starter/generator 1 is distributed depending onthe operating mode to the other busbars via the PRIMARY busbar.Systems with a high current flow such as starter/generator 1 aredirectly connected to the PRIMARY busbar 1. The connections to

systems and busbars are protected by fuses and controlled by severalcontactors.

A control circuit disconnects the primary busbar 1 from the remainingsystem if a short-circuit occurs. The installed generator control unitcontrols and monitors the operation of starter/generator 1 andswitches it off if a failure occurs.

The electrical master box is electrically connected to the battery

master box and the electrical master box 2. The BUS TIE functionconnects the PRIMARY busbar 1 to the PRIMARY busbar 2 and theBATTERY busbar.

Operating conditions of the system 1 are indicated by the electricalmaster box 1 on the CAD.

Test Function

A built-in test function may be activated by a switch located at themaster box housing (after removing the inner lining) and indicatespossible failures in the electrical master box 1 by means of indicatorlights.

The electrical master box 1 is installed behind the LH interior panelingnear frame 5.

Electrical Masterbox System 2

The construction and function of the electrical master box 2 is similarto that of the electrical master box 1. The function of the electrical

master box 2, however, may be extended by inserting an additionalprinted circuit board and a connector. This board controls and monitorsthe operation of an external power unit which is connected to theexternal power receptacle.

Location

The electrical master box 2 is installed behind the RH interior panelingnear frame 5.


Electrical Master Boxes




Power Supply Cable

BondingJumper

Boards

Plug

Connector for EPU(Masterbox 2 only)

Connectors for HighPower Consumers

Frame 5

FWD







Electrical Master Boxes 1 / 2

Electrical Master Box 1 Electrical Master Box 2




D Bus TieF2 Ess. BusB GeneratorF1 Shed. Bus

A StarterE High Load BusF High Power

(Box1 Ext. Hoist,Box 2 A/C)

H EPU Connector

Fuse

Board

Z300

Z200

Z100

Z300

Z200Z100


Built-In Test

The built-in test enables during maintenance on ground to check thefunctions of the master box. The following conditions are necessary:

Failure Indications

The following failures can be indicated by the corresponding lettersand numbers:




-- The master box must be supplied by the EPU

-- The generators are standing still

-- The generator switch is in position NORM

-- The switch BAT MST is in position ON

-- The start relay is open

-- The switch SHED BUS is in position NORM

-- If high power consumers (e.g. external hoist, airconditioning syst.) are installed the systems have to beswitched on

Test Procedure

The TEST push button must be pressed for the duration of the test run.

Minimum for 10 seconds. During the built-in test running the red LED“r” is illuminated. If the test was successful, the green LED “o” isilluminated for a short time. If there is a failure detected in themasterbox, a red LED of the corresponding failure and the red LED “f”comes on.

F not used

E not used

D not used

C Fuses of internal supply of Z 500 and Z 600

B Distributing fuses (Essential bus, Shedding bus, High Loadbus, high power consumers)

A not used

9 Bus tie relay

8 not used

7 Shedding bus relay

6 High load bus relay

5 High power relay4 GEN relay

3 Test and supply board Z 300

2 Logic and guard board Z 200

1 Generator control board Z 100

f Test failed (red)

o Test successful (green)

r Test is running (red)


Electrical Master Box -- Built--In Test Indication




TEST 28V 0V

POR

Red LED: Test is running

Green LED: Test O.K.

Receptacle forVoltage Measuring

Test Switch

+

F E D C B A 9 8 7 6 5 4 3 2 1 of r

Potentiometer forGenerator OutputVoltage

Red LED: Test failed


Battery System

General

Th b tt li t f l f ti




The battery supplies current for several functions:

-- Starting the engines

-- Supplying the vital electrical systems, if both generators fail

-- On ground when the engines are not running

Components

The battery system comprises the following components:

-- Battery with temperature switch


-- Blowout fuses

-- Battery bus

-- Essential bus relay EBC1/2 and battery relay BATC

Battery

The battery consists of 20 nickel-cadmium cells installed in a housingwhich is ventilated/vented by two openings. A temperature switch isinstalled in the housing which closes contact at a temperature of 70!3 °C and thus activates the indication BAT TEMP in the warningunit. The battery is electrically connected to the DC power system via

a power connector. The temperature switch has an individualconnector which is connected to the warning unit.


BatteryCover




NC Cell

VentilationConnector for Temperature Switch

Power ConnectorHousing

Mounting Bolt

Fixing Nut with Pin

Mounting FrameBattery Housing

Equipment Deck

TemperatureSwitch


Battery Master Box

The battery master box controls the operation of the battery.The battery is charged, if at least one of the generators suppliescurrent If the battery is used for power supply the battery busbar

Blowout Fuse

A fuse (325A) located in the battery bonding line melts when thecurrent flow is excessive and thus prevents the system from beingdamaged.




current. If the battery is used for power supply, the battery busbardelivers current to the system.

With connections from the battery busbar to both the ESSENTIALbusbars 1 and 2 the supply is done in case of failure of both thegenerators. The connections are fused by blowout fuses.

By occuring failures the battery and the battery busbar are isolatedautomatically from the PRIMARY busbar.

During operation the actual current or voltage provided by thegenerators or the battery is displayed by the VEMD. If the batteryoperates as the power source, it will be discharded. The warningdisplay BAT DISCH illuminates at the warning unit.

NOTE During long time opteration on ground with EPU itis recommended to disconnect the battery in order

to avoid any discharging via the ESSENTIAL BUSor the power consumption in the battery master

box. As the battery relay is open, the battery can

not be charged by the EPU or vice versa. (EPU

voltage < battery voltage)

The battery master box is installed in the lower part of the aft fuselagesection below the battery.

The fuse is mounted next to the battery master box to the fuselage.


Battery Master Box




Blowout Fuse325 A

Plug

SupportPower Supply

Lines

BondingJumper





Built-in Test

The built-in test enables during maintenance on ground to check thefunctions of the master box. The following conditions are necessary:

-- The battery master box must be supplied by the battery

Failure Indications

The following failures can be indicated by the corresponding lettersand numbers:

1 Stabilizing board Z 100




-- The generators are standing still

-- The switch BAT MST is on position ON

-- The start relay is open

Test ProcedureThe TEST push button must be pressed for the duration of the test run.Minimum for 10 seconds. During the built-in test running the red LED“r” is illuminated. If the test was successful, the green LED “o” isilluminated. If there is a failure detected in the battery master box, a redLED of the corresponding failure comes on.

g

2 Power supply board Z 200

3 Internal supply fuses Z 500

4 Bonding fuse

5 Essential distribution fuses

6 BAT Relay circuit

7 not used

f Test failed (red)

o Test successful (green)

r Test running (red)


Battery Master Box -- Built-In Test Indication




DIST 1 DIST 2 DIST 3 DIST 4 DIST 5 DIR BAT

TEST

LED Indication

Battery Master Box

Test Switch

Circuit Breakers (Optional)

7 6 5 4 3 2 1 f o r


Switches GEN I, GEN II, BAT MSTR

GeneralThe switches GEN I and GEN II are three position toggle switches withthe positions:

If a generator or the battery should be engaged after a failure, therespective switch must be set to the position RESET. This provides areset of failure indications and of the protective functions.Subsequently the switch can be set to the position NORM




NORM--OFF--RESET.

The position RESET is spring loaded to the position OFF.

The switch BAT MSTR is a three position toggle switch with thepositions:

ON--OFF--RESET.

The position RESET is spring loaded to the position OFF.

Location

The switches GEN I, GEN II and BAT MSTR are mounted to theswitching unit in the middle part of the instrument console.

FunctionThe position NORM of the switch GEN I/II activates the generator bythe corresponding master box reaching the n1 speed of 50 %.

In position OFF, the generator is disconnected from the power supplysystem.

The position ON of the switch BAT MSTR connects the battery or theEPU via the battery master box to the power supply system.

The position OFF disconnects the battery/EPU from the power supplysystem.

Subsequently the switch can be set to the position NORM.

CDS/CPDS Indication

The voltage and the current of the generators and the battery areindicated on the CDS/CPDS. If there is a generator isolated from thehelicopter’s power supply (with the electrical system is active), thecaution GEN DISCON will be displayed in the SYSI / SYS II field of theCDS/CPDS.

In case of overtemperature the caution GEN OVHT will be displayedin the SYSI / SYSII field of the CDS/CPDS.

Warning Unit Indication

The warning indications BAT TEMP and BAT DISCH are integrated inthe warning unit display.In case of battery overtemperature (> 70 °C)the indication BAT TEMP comes up at the warning unit.

If the battery operates as the power source, it will be discharged. At acurrent of more than approx. 2 A the warning BAT DISCH comes up

at the warning unit display.


Power Supply -- Switches and Indications

Warning Unit




TRAIN SELFADEC

ONFLIGHT

ONIDLE

OFF

ENG

OFF

FADEC ENG

ARM

OFF

ENG CONTROL

OFF

OFF

OFF

NORMONNORM

RESET RESET RESET

GEN I BAT MSTR GEN II

DC POWER CONTROL

Voltage and Current Indication

Switch Unit

Temperature Warning of Battery

Temp > 70 °C

FLIGHT

IDLE

OFFOFF

CDS

Switches for DCPower Supply

Discharge Warning of Battery

I > 2 Ampere

CPDS



General

An external power receptacle is installed in the helicopter to connectan external power unit (EPU) It is protected by a cover The external

Power Connector

A mechanical safety-device prevents the socket from being insertedincorrectly The negative pin of the power connector is connected to




an external power unit (EPU). It is protected by a cover. The externalpower unit should supply at least 24 V DC. The external powerreceptacle is designed to a (short-time) current flow of up to 700 A.

The external power receptacle is installed on the RH side of the

helicopter beyond the lower maintenance step.

Components

The external power receptacle consists of:

-- Power connector

-- Intercom socket

-- Circuit breaker EXT PWR

-- Switch EPU DOOR-- CDS/CPDS Indication

incorrectly. The negative pin of the power connector is connected tothe bonding point E1 (connection to the bonding system) via aconductor rail. The two large pins are used for the negative andpositive poles. The shorter pin (positive, +1) is used for engaging thebattery master box. The current flows over the two large pins, until thecontacts are closed savely.






Switch

Circuit Breaker

Receptacle

GroundConnection

--

Switch EPU DOOR

Receptacle

Circuit Breaker

Intercom Socket

+1+

E1


Intercom Socket

An intercom may be connected to the aircraft intercommunicationsystem. It enables the maintenance personnel to communicate withpersons in the cockpit even during excessive noise levels (e.g. whenthe engines are running).

Power Supply on Ground

If power supply on ground is ensured by an external power unit, bothstarters/generators and the battery are disconnected (generator relay1/2 and battery relay are open) from the PRIMARY busbars. Theycannot be connected to these busbars together with the external

i




Circuit Breaker

By means of the circuit breaker the control line for the external powerreceptacle is activated. When the circuit breaker is pressed, the

electrical master box 2 disconnects the battery and bothstarter/generators from the PRIMARY busbars. On the CDS/CPDS thecautions BAT DISCON (MISC), GEN DISCON (SYSI/II) are displayed.

CDS/CPDS Display

The display EXT PWR indicates that an external power unit isconnected and activated. The display is controlled by the electricalmaster box 2. The display EPU DOOR indicates that the cover at the

external power receptacle is open. It is activated by the EPU DOORswitch. Both displays are integrated in the CDS/CPDS and areindicated in the MISC area.

Function of the Ext. Power Receptacle

The connection of an external power unit to the helicopter’s powersupply system is controlled by the electrical master box 2. Thefollowing modes are available:

-- Power supply on ground

-- Starting the engines

NOTE Charging the battery with the EPU is not possible.

power unit.

Starting the Engines

If starting of the engines is effectedby means of an external power unit,

both starter/generators serve as starter for the engines, however, theyare disconnected from the helicopter’s power supply system as soonas the engines are running and the starter/generators operate asgenerators, i. e. supply current. For starting the external power unitshould supply currents of 500 -- 600 A at a nearly constant voltagelevel.


External Power Receptacle -- Function




EXTPWR

EPU DOOR

Electrical MasterBox 2

EPU PowerConnector

Switch Position: Door Open

+1 +

CDS/CPDS

GEN DISCON GEN DISCONEXT PWREPU DOORBAT DISCON


DC Power Distribution

General

The DC power distribution system routes the direct current supplied bythe battery, generators or the external power unit to the individual

Busbars

The following busbars route the current to the individual consumers:

ESSENTIAL busbar 1 (PP10E)




y, g ppower consumers via several busbars.

Overhead Panel

General

Busbars and circuit breakers supplying the consumers with current areintegrated in the overhead panel. Several systems are activated andcontrolled at the overhead panel.

Assembly

The overhead console consists of two component brackets and thefront panel containing the components and the busbars on the rear.

All circuit breakers, switches and rheostats are mounted on the frontpanel. The relays, fixed resistors and all other components aremounted on the component brackets.

The front panel consists of three parts which each have backgroundlighting and bear the decals of the installed circuit breakers, switches

and rheostats.

-- ESSENTIAL busbar 1 (PP10E)

-- ESSENTIAL busbar 2 (PP20 E)

-- SHEDDING busbar 1 (PP10S)

-- SHEDDING busbar 2 (PP20S)

Additionally, the following busbars are available at the overhead panelfor AC voltage:

-- AC busbar 1

-- AC busbar 2

The essential consumers are connected to the two ESSENTIALbusbars. Further DC power consumers are connected to the

SHEDDING busbars. Consumers which require AC voltage areconnected to the AC busbars.

The overhead panel is supplied with DC voltage by the PRIMARYbusbars 1 and 2 or the BATTERY busbar via the blocking diodes. TheBATTERY busbar supplies the ESSENTIAL busbars 1 and 2. Furtherlines coming from the master boxes 1 and 2 supply the SHEDDINGbusbars 1 and 2.


Overhead Panel

Switch SHEDDING BUS

Switch BUS TIE I




AC BUS I

SHEDDINGBUS I

ESSENTIAL BUS I

AC BUS II

SHEDDINGBUS II

ESSENTIALBUS II

Switch BUS TIE II


Switch SHED BUS

The switch SHED BUS is a two position switch with the positionsNORM / EMER. The NORM position is protected by a safety guardwhich has to be opened before switching to the EMER position.

In position NORM the relays SBC1 and SBC2 are closed, as soon thefi li h




first generator supplies power to the system.

In position EMER the relays SBC1 and SBC2 are re-closed. Thisswitch position is selected, if both generators should fail or if the

system should be supplied by the battery.Switches BUS TIE I / II

The switches BUS TIE I / II are three position toggle switches with thepositions NORM / OFF / RES. The switches are protected by a safetyguard, which positions the switch in the NORM position.

The switches allow the coupling or decoupling of the PRIMARYbusbars 1 / 2 with the relays BTC1 and BTC2.

In position NORM the respective bus tie relay is closed. The positionOFF opens the respective bus tie relay. The position RES allows aftera system failure again to close the respective bus tie relay.


Overhead Panel -- Switches




EMER

O

F

F

M

A

X

NORM


Circuit Breaker Console 1 and 2

General

The HIGH LOAD busbar 1 is installed in the circuit breaker panel 1, theHIGH LOAD busbar 2 is installed in the circuit breaker panel 2. Allcircuit breakers which are connected to one of both HIGH LOAD




busbars are installed in the respective circuit breaker panel.Consumers with high energy demand are connected to both HIGHLOAD busbars.

Circuit Breaker Console 1

Circuit breaker panel 1 contains the HIGH LOAD busbar which isdirectly supplied with DC voltage from PRIMARY busbar 1 in theelectrical master box 1. It is also equipped with the 28V DC receptacleand a connector for the “Inflight Track & Balance” system.

Circuit Breaker Console 2

Circuit breaker panel 2 contains the HIGH LOAD busbar 2 which isdirectly supplied with DC voltage from PRIMARY busbar 2 in theelectrical master box 2.

Locations

The circuit breaker consoles are installed on the LH side and on theRH side of the cargo bay, respectively.


Circuit Breaker Panel 1 and 2

DC RECEPT

DC Receptacle

Functional Schematic




DCRECEPT

TR&BALINFLT

3MJA

19VVA

10 5

10APP 10H

Bonding Connector 100 VV

DC Receptacle

Circuit Breaker

Circuit Breaker Panel 1 Circuit Breaker Panel 2

20 5


Function -- Complete System

General

The following operating modes are possible in the DC power system:

-- Operation with battery (emergency function)

Switch Positions

The switches must be set to the following positions:

BAT MSTR ON




-- Automatic engine starting

-- One generator working

-- Generators working in parallel (normal function)

-- Generators working individually

-- Operation with external power unit (EPU)

-- Connection of SHEDDING busbars 1 and 2

-- System reactions due to malfunctions

Operation with Battery (Emergency Operation)

The battery supplies the BATTERY busbar with current. BothESSENTIAL and PRIMARY busbars are supplied by this busbar. TheHIGH LOAD busbars 1 and 2 and the SHEDDING busbars 1 and 2 arenot supplied with current.

The warning BAT DISCH is illuminated on the warning panel.

GEN I NORM/OFF in case of emer-gency operation

GEN II NORM/OFF in case of emer-

gency operation

SHED BUS NORM

BUS TIE I NORM

BUS TIE II NORM

CDS/CPDS Cautions

The following cautions are displayed on the CDS/CPDS:

SYS I MISC SYS II


The following electrical values are displayed on the CDS/CPDS:

SYS I SYS II

DC VOLT 24 24

GEN AMPS 0 0

BAT AMPS current load


Operation with Battery

Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2




PP10E PP20E

1 2

PP10S

SBC1

HLC1BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20SPP10H

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

O N O F F

R E S

E T

N O R M

O F F

R E S

E T

N O R M

O F F

R E S

E T

R E S

E T

O F F

N O R M

E M

E R

O N

N O

R M

R E S

E T

O F F

N O R M

BAT MSTR GEN I GEN II SHED BUS BUS TIE I BUS TIE IISwitch Position

HPC2

F F

Ext. Hoist (opt.)

A/C (opt.)

HPC1



The engines can be started by means of the battery or an externalpower unit (refer to operation by means of an external power unit).

The battery supplies the PRIMARY busbars 1 and 2 and theESSENTIAL busbars 1 and 2 with current via the BATTERY busbar.To start the engines the starter/generator 1 is supplied with current

CDS/CPDS CautionsThe following cautions for the respective engine during the start-up aredisplayed on the CDS/CPDS:

SYS I MISC SYS II




g g ppfrom the PRIMARY busbar 1, the starter/generator 2 from thePRIMARY busbar 2. The engines can only be started successively.When n1 exceeds 50%, the battery master box disconnects the battery

from the power supply circuit and the generator of the started enginesupplies current to the electrical system.

Switch Positions


BAT MSTR ON

GEN I NORM

GEN II NORM

SHED BUS NORM

BUS TIE I NORM

BUS TIE II NORM

In addition:

FADEC ON

ENG CONTROL ENG I IDLE/FLIGHT


STARTER STARTER

CDS/CPDS Indications


SYS I SYS II

DC VOLT 24 24

GEN AMPS 0 0

BAT AMPS current load

The warning BAT DISCH illuminates on the warning unit.



Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2




PP10E PP20E

1 2

PP10S

SBC1

HLC1BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20SPP10H

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

O N O F F

R E S

E T

N O R

M

O F F

R E S

E T

N O R

M

O F F

R E S

E T

R E S

E T

O F F

N O R

M

E M

E R

O N

N O

R M

R E S

E T

O F F

N O R

M

BAT MSTR GEN I GEN II SHED BUS BUS TIE I BUS TIE II

Ext. PowerReceptacle

Switch Position

HPC2

F F

Ext. Hoist(opt.)

A/C (opt.)

HPC1


Operation with One Generator

The HIGH LOAD busbars 1 and 2 are disconnected from the system.The battery is again charged via the BATTERY busbar. Generator 1supplies PRIMARY busbar 1 and, via the BUS TIE connectionPRIMARY busbar 2 with current. The SHEDDING busbars 1 and 2 andthe ESSENTIAL busbars 1 and 2 are supplied with current by thePRIMARY busbar 2

Switch Position


BAT MSTR ON

GEN I NORM




PRIMARY busbar 2.

If the defective generator 2 is operative, it can be connected again (setGEN II switch first to position RESET, then to NORM).

Automatic deactivation of the HIGH--LOAD busbars and (optional)high-current consumers (except Ext. Hoist) prevents overload of thegenerator still in operation.

GEN II NORM/OFF/RESET

SHED BUS NORM

BUS TIE I NORMBUS TIE II NORM

CDS/CPDS Cautions


SYS I MISC SYS IIGEN DISCON



SYS I SYS II

DC VOLT 28 28

GEN AMPS current load 0

BAT AMPS charging current, if provided (negative)


Operation with One Generator

Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2




PP10S PP10E PP20E

1 2

SBC1

HLC1 BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20SPP10H

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

HPC1

HPC2

F F

Ext. Hoist(opt.)

A/C (opt)

O N O F F

R E S

E T

N O R

M

O F F

R E S

E T

N O R

M

O F F

R E S

E T

R E S

E T

O F F

N O R

M

E M E

R

O N

N O R M

R E S

E T

O F F

N O R

M




Operation with Generators Connected in Parallel

Both starter/generators operate as power sources and supply currentto their respective PRIMARY busbars, which in turn supply all the otherbusbars with current. The battery is charged via the BATTERY busbar.The system load is shared equally by both generators due to theconnection of PRIMARY busbar 1 to PRIMARY busbar 2, i. e., theBUS TIE I and II switches are set to NORM.

Switch Position


BAT MSTR ON

GEN I NORM




GEN II NORM

SHED BUS NORM

BUS TIE I NORMBUS TIE II NORM



SYS I SYS IIDC VOLT 28 28

GEN AMPS current load current load

BAT AMPS charging current, if provided (negative)

The current load on generator 1 and generator 2 is identical.


Operation with Parallel Connected Generators

Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2




PP10HPP10S PP10E PP20E

SBC1

HLC1 BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20S

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

O N O F F

R E S

E T

N O R

M

O F F

R E S

E T

N O R

M

O F F

R E S

E T

R E S

E T

O F F

N O R

M

E M E R

O N

N O R M

R E S

E T

O F F

N O R

M



HPC2

F F

Ext. Hoist(opt.)

A/C (opt.)

HPC1


Operation with Separated Generators

With the BUS TIE I in the OFF position both PRIMARY busbars aredisconnected. Each generator supplies the respective PRIMARYbusbar only and the generator load will be different. Generator 2additionally charges the battery.

The HIGH LOAD busbars 1 and 2 are disconnectedfrom thehelicopterpower supply system

CDS/CPDS Cautions


SYS I MISC SYS II

BUSTIE OPN




power supply system.

Switch Position


BAT MSTR ON

GEN I NORM

GEN II NORM

SHED BUS NORM

BUS TIE I OFF

BUS TIE II NORM

In position NORM the BUS TIE switches are protected by means of acover against unintended operation.



SYS I SYS II

DC VOLT 28 28

GEN AMPS current load current load

BAT AMPS Charging current, if provided (negative)


Operation with Separated Generators

Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2





SBC1

BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20S

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

O N O F F

R E S

E T

N O R

M

O F F

R E S

E T

N O R

M

O F F

R E S

E T

R E S

E T

O F F

N O R

M

E M

E R

O N

N O

R M

R E S

E T

O F F

N O R

M



HLC1

HPC2

F F

Ext. Hoist(opt.)

A/C (opt.)

HPC1


Operation with External Power Unit

The electrical master box 2 connects the external power unit to the

PRIMARY busbar 2. If the BUS TIE I and BUS TIE II switches are setto NORM, the PRIMARY busbar 1 is again supplied with current. Allother busbars, except the BATTERY busbar, are supplied with currentby both PRIMARY busbars. The BATTERY busbar is only connectedto the battery and both ESSENTIAL busbars and disconnected fromth i i l t l th t l it

CDS/CPDS Cautions


SYS I MISC SYS II

GEN DISCON BAT DISCONEXT POWER

GEN DISCON




the remaining power supply system as long as the external power unitis connected. The battery cannot be recharged by means of theexternal power unit.

Both starter/generators are also disconnected from the power supplysystem, as long as the external power unit supplies current. They cannot be connected.

Switch Position


BAT MSTR ONGEN I NORM/OFF

GEN II NORM/OFF

SHED BUS NORM

BUS TIE I NORM

BUS TIE II NORM

In addition, the circuit breaker on the external power receptacle mustbe activated to enable the external power supply to be connectedthrough the electrical master box 2.

EPU DOOR



SYS I SYS II

DC VOLT 28 28

GEN AMPS

BAT AMPS 0

There is no load indication of the EPU.


Operation with External Power Unit

Generator1

Generator2

GC1SC1 SC2GC2

El. MasterBox 1

El. MasterBox 2





SBC1

HLC1 BTC1

PP20H

BATC

EBC2EBC1

SBC2

HLC2

BTC2 GPUC

PP20S

Battery

PRIMARY-Busbar 1

BATTERY--Busbar

PRIMARY-Busbar 2

Battery Master Box

O N O F F

R E S E T

N O

R M

O F F

R E S E T

N O

R M

O F F

R E S E T

R E S E T

O F F

N O

R M

E M

E R

O N

N O

R M

R E S E T

O F F

N O

R M



HPC2

F F

Ext. Hoist(opt.)

A/C (opt.)

HPC1




AC Power System

General

The AC power system generates 26 V and 115 V AC voltage with 400Hz each out of 28 V DC voltage. The helicopter is equipped with onesystem (SYS 2) or two systems (SYS 2 and SYS 1). The AC voltagesare distributed via busbars and modules.




The alterning voltages are used for navigation instruments and for theStability Augmentation System (SAS).

Components

The system 2 of the AC power system consists of the following:

-- Static inverter

-- Circuit breaker INV 2

-- Switch INV 2

-- AC busbar-- Modules

-- CDS/CPDS as display unit


AC (400 Hz) Power System

AC--Busbar 2 AC--Busbar 1

Switch AC BUS SEL




E N GI

ENGII

OFF

M AX

Static

Inverter

Plug

Circuit Breaker INV 2

Switch INV 2Switch INV 1

Circuit Breaker INV 1


Static Inverter

The static inverter 2 collects DC voltage from the ESSENTIAL busbar

by pushing the circuit breaker INV 2 PWR and closing the INV 2 switch.It converts the supplied 28 V DC into two AC voltages 26 V and 115 Vwith 400 Hz each. The voltages are then stabilized in the staticinverter 2. They are distributed to the consumers via modules and the AC busbar 2.

The static inverter 2 is installed on the RH side behind the interior

CDS/CPDS Cautions

If the static inverter 2 is defective, INVERTER is displayed in theSYS II

area of the CDS/CPDS. If there is 28 V DC at the CDS/CPDS input,the caution will disappear.

The following conditions at the signal output of the inverter arepossible:

-- 28 V DC: CDS/CPDS Caution off




The static inverter 2 is installed on the RH side behind the interiorpaneling behind frame 4.

Circuit Breaker

The circuit breaker INV 2 is installed in the overhead panel.

Switch

The switch INV 2 is installed in the overhead panel.

AC Busbar

The AC busbar 1 and 2 are integrated in the overhead panel. Theydistribute the AC voltage to their consumers as long as the inverterselect switch is in the NORM position (2 inverters installed). After afailure of one inverter the remaining inverter can be selected for thecomplete AC system by switching to position INV1/INV2.

Modules

The modules for AC high/low are installed in the cabin roof.

-- Open circuit: CDS/CPDS Caution on


AC (400 Hz) Power System -- Functional Schematic

26VAC

INV 2NORM

INV 1




ESSENTIAL Busbar 2

115VAC

400 Hz

26VAC

400 Hz

115VAC400 Hz

26VAC

400 HzINV 1 PWR

INV 2 PWR INV 2

INV 1

INV SEL

AC BUS SPLY

INV 1

26VAC

400 Hz

ESSENTIAL Busbar 1

Inverter 2

Inverter 1

ModulesCDS/CPDS

115VAC

400 Hz

115VAC

400 Hz

26VAC

400 Hz

INVERTER INVERTER

EC 135Training ManualInspections

Inspections





Table of Contents

Types of Inspections 3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Scheduled Checks and Inspections 4. . . . . . . . . . . . . . . . . . . . .




Scheduled Checks and Inspections

General

To guarantee the airworthiness of the EC 135 helicopter, checks andinspections have to be carried out according to chapter 05 of the AMM.

The EC 135 inspection system in general is split into:

-- ChecksT b i d t b th il t h i ith t th

Preflight Check

The preflight check isto be performed by the latest prior to the first flightof the day.

The checklist is included in the flight manual, resp. pilots checklist andcan be carried out by the pilot or a mechanic. Only “on the job” trainingis neccessary.




To be carried out by the pilot or a mechanic without the

need of an inspector.-- Inspections.

To be carried out by a mechanic and signed by aninspector.

Types of Checks and Inspections

The following checks and inspections have to be carried out accordingto the maintenance manual/flight manual:

-- Preflight check (O--level)

-- Complementary check 50 Fh (O--level)

-- Complementary check 100 Fh (O--level)

-- Intermediate inspection 400 Fh (I--level)

-- Periodical inspection 800 Fh or 2 years (O/I--level)

-- Supplementary inspections acc. to operating time

-- Inspections after operation under special environmentalconditions

-- Special inspections after maintenance activities

-- Ground run / functional check flight

is neccessary.

Complementary Checks A) Every 50 flight hours a complementary check has to be performed.The time limit of 50 h may be exceeded by up to 10 flight hours.

The complementary check 50 Fh can be carried out by the pilot or amechanic. Only “on the job” training is neccessary.

B) Every 100 flight hours a complementary check has to be performed.The time limit of 100 h may be exceeded by up to 10 flight hours.

The complementary check 100 Fh can be carried out by the pilot or amechanic. Only “on the job” training is neccessary.

Intermediate Inspection

An intermediate inspection has to be performed:

-- After 400 flight hours TSN (time since new)

-- then 400 flight hours after due time of a periodicalinspection

The time limit 400 h may be exceeded by up to 80 flight hours. If performed at the same due time, the intermediate inspection isreplaced by the periodical inspection.


Inspections

SCHEDULED INSPECTIONS

Preflight Check




05--21--00, 6--2 Complementary Check 50 Fh

05--21--00, 6--3 Complementary Check 100 Fh

05--22--00 Intermediate Inspection 400 Fh

05--24--00 Periodical Inspection 800 Fh or every 2 Years

05--23--0012--MonthInspection

05--25--00 Supplementary Inspections acc. to Operating Time


12--Month Inspection

An12--month inspection is to be performed acc. to AMM 05--23--00

page 601:

The time limit of 12 month may be exceeded by up to 3 month.

If performed at the same due time, the 12--month inspection isreplaced by the periodical inspection.

Periodical Inspection

Conditional Inspections after Operational Incidents

These inspections have to be performed after specific operational

incidents either prior to the next flight or at specified time intervals.

The inspections ensure that airworthiness will be maintained or maybe restored as a result of specific maintenance activities.

Ground Run and Functional Check Flight

Section 05--60--00 contains the procedures for ground check run andfunctional check flight




Aperiodical inspection is to be performed

-- After 800 flight hours TSN or two years TSN, whichevercomes first

-- then 800 flight hours or every 2 years, wichever occurs first

Supplementary Inspection acc. to Operating Time

Supplementary inspections are to be performed. The given time limitmay be exceeded by 10% of the resp. interval.

Conditional Inspections after Maintenance Activities

Conditional Inspections have to be performed, due to performance of a maintenance measure after time limits of parts and componentshave been reached. The given time limit may be exceeded by 10% of the resp. interval.

functional check flight.

The description for both helicopter models is provided in forms andarranged as a test report, that may be equally used for performing andrecording purposes.

The scope of ground check run and functional check flight may berestricted depending on maintenance measures performed. Possiblerestrictions are listed in front of test reports.





CONDITIONAL INSPECTIONS

GROUND RUN AND FUNCTIONAL FLIGHT

05--51--00 Conditional Inspections after Operational Incidents

05--52--00 Conditional Inspections after Maintenance Activities

05--60--00 Ground Run and Functional Check Flight

MANUAL DE VUELO EC 135

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Transcript of MANUAL DE VUELO EC 135