Installation Manual Walter m601e, m601e-21[1]

INSTALLATION MANUAL

TURBOPROP ENGINE

MODELS

WALTER

M601E

WALTER

M601E-21

MANUAL PART No. 0982502

SECOND REVISED EDITION

ISSUED JUNE 26, 2001

This Manual was approved and signed in Czech by:

Oldřich Matoušek, Jan Beneš

Development Director Civil Aviation Authority

of the Czech Republic

Date: June 26, 2001 Date: July 10, 2001

WALTER a.s.

PRAHA 5 - JINONICE

CZECH REPUBLIC

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

RECORD OF REVISIONS

RECORD OF REVISIONS

The date on which new pages have been inserted into the Manual is affixed by the operator. The Bulletin No. is specified only if the revision has been issued as a Bulletin.

REVI-SION No. BULLETIN No. ISSUE DATE OF NEW PAGES NUMBERS OF AFFECTED PAGES DATE OF INSERTION AND SIGNATURE 1 Nov 15, 2004 Pages 0-1, 0-3, 0-4, 1-1, 2-1 to 2-8, 3-2, 3-10 to 3-38, 5-1, 5-7, 5-8, 8-3, 9-1, 9-2, 12-1, 12-2, 12-3, 12-6, 12-8, 13-1 2 May 11, 2006 Pages 0-1, 0-3, 0-4, 1-1, 4-2, 7-23, 12-6, App. 7

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INSTALLATION MANUAL

MANUAL PART No. 0982502

RECORD OF REVISIONS

REVI-SION No. BULLETIN No. ISSUE DATE OF NEW PAGES NUMBERS OF AFFECTED PAGES DATE OF INSERTION AND SIGNATURE

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INSTALLATION MANUAL

MANUAL PART No. 0982502

REVIEW OF EFFECTIVE PAGES - SHEETS

REVIEW OF EFFECTIVE PAGES - SHEETS

Section Page - Sheet Date

Title Sheet - in English Jun 26, 2001 - in Czech 26.6.2001 Record of 0-1 May 11, 2006 revisions 0-2 Jun 26, 2001 Review of effective 0-3 May 11, 2006

sheets - pages 0-4 May 11, 2006

Contents 0-5 Jun 26, 2001 0-6 blank Jun 26, 2001 Introduction 0-7 Jun 26, 2001 0-8 Jun 26, 2001 Section 1 1-1 Nov 15,2004 Engine 1-2 Jun 26, 2001 Description 1-3 Jun 26, 2001 1-4 Jun 26, 2001 Section 2 2-1 Nov 15,2004 Engine 2-2 Nov 15,2004

Operation Limits 2-3 Nov 15,2004

2-4 Nov 15,2004 2-5 Nov 15,2004 2-6 Nov 15,2004 2-7 Nov 15,2004 2-8 blank Nov 15, 2004 Section 3 3-1 Jun 26, 2001 Engine 3-2 Nov 15,2004 Performance 3-3 Jun 26, 2001 3-4 Jun 26, 2001 3-5 Jun 26, 2001 3-6 Jun 26, 2001 3-7 Jun 26, 2001 3-8 Jun 26, 2001 3-9 Jun 26, 2001 3-10 Nov 15,2004 3-11 Nov 15,2004 3-12 Nov 15,2004 3-13 Nov 15,2004 3-14 Nov 15,2004

Section Page - Sheet Date

Section 3 3-15 Nov 15,2004 (continued) 3-16 Nov 15,2004 3-17 Nov 15,2004 3-18 Nov 15,2004 3-19 Nov 15,2004 3-20 Nov 15,2004 3-21 Nov 15,2004 3-22 Nov 15,2004 3-23 Nov 15,2004 3-24 Nov 15,2004 3-25 Nov 15,2004 3-26 Nov 15,2004 3-27 Nov 15,2004 3-28 Nov 15,2004 3-29 Nov 15,2004 3-30 Nov 15,2004 3-31 Nov 15,2004 3-32 Nov 15,2004 3-33 Nov 15,2004 3-34 Nov 15,2004 3-35 Nov 15,2004 3-36 Nov 15,2004 3-37 Nov 15,2004 3-38...blank Nov 15,2004 Section 4 4-1 Jun 26, 2001 Engine 4-2 May 11, 2006 Mounting 4-3 Jun 26, 2001 4-4 Jun 26, 2001 4-5 Jun 26, 2001 4-6 Jun 26, 2001 4-7 Jun 26, 2001 4-8 Jun 26, 2001 4-9 Jun 26, 2001 4-10 blank Jun 26, 2001 Section 5 5-1 Nov 15,2004

Air Inlet System 5-2 Jun 26, 2001

5-3 Jun 26, 2001 5-4 Jun 26, 2001 5-5 Jun 26, 2001 5-6 Jun 26, 2001 5-7 Nov 15,2004 5-8 Nov 15,2004

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REVIEW OF EFFECTIVE PAGES - SHEETS

Section Page - Sheet Date

Section 6 6-1 Jun 26, 2001

Fuel System 6-2 Jun 26, 2001

Section 7

Electrical 7-1 Jun 26, 2001

System 7-2 Jun 26, 2001

and Monitoring 7-3 Jun 26, 2001

Instruments 7-4 Jun 26, 2001 7-5 Jun 26, 2001 7-6 Jun 26, 2001 7-7 Jun 26, 2001 7-8 Jun 26, 2001 7-9 Jun 26, 2001 7-10 Jun 26, 2001 7-11 Jun 26, 2001 7-12 Jun 26, 2001 7-13 Jun 26, 2001 7-14 Jun 26, 2001 7-15 Jun 26, 2001 7-16 Jun 26, 2001 7-17 Jun 26, 2001 7-18 Jun 26, 2001 7-19 Jun 26, 2001 7-20 Jun 26, 2001 7-21 Jun 26, 2001 7-22 Jun 26, 2001 7-23 May 11, 2006 7-24 Jun 26, 2001 Section 8 Lubrication 8-1 Jun 26, 2001 System 8-2 Jun 26, 2001 8-3 Nov 15,2004 8-4 blank Jun 26, 2001 Section 9 Cooling 9-1 Nov 15,2004 Requirements 9-2 Nov 15,2004 Section 10

Exhaust System 10-1 Jun 26, 2001

10-2 blank Jun 26, 2001 Section 11

Airbleed System 11-1 Jun 26, 2001

11-2 blank Jun 26, 2001

Section Page - Sheet Date

Section 12 Engine 12-1 Nov 15,2004 Accessories 12-2 Nov 15,2004 12-3 Nov 15,2004 12-4 Jun 26, 2001 12-5 Jun 26, 2001 12-6 May 11, 2006 12-7 Jun 26, 2001 12-8 Nov 15,2004 Section 13 Coolant 13-1 Nov 15,2004

Injection 13-2 blank Jun 26, 2001

Section 14

Propeller Unit 14-1 Jun 26, 2001

14-2 Jun 26, 2001

14-3 Jun 26, 2001

14-4 blank Jun 26, 2001 Section 15

Engine Controls 15-1 Jun 26, 2001

15-2 Jun 26, 2001 15-3 Jun 26, 2001 15-4 blank Jun 26, 2001 Appendix 1 Jun 26, 2001 2 Jun 26, 2001 3 Jun 26, 2001 4 Jun 26, 2001 5 Jun 26, 2001 6 Jun 26, 2001 7 May 11, 2006 8 Jun 26, 2001

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INSTALLATION MANUAL

MANUAL PART No. 0982502

CONTENTS

CONTENTS

Section Description Page No.

INTRODUCTION 0-7

1 ENGINE DESCRIPTION 1-1

2 ENGINE OPERATION LIMITS 2-1

3 ENGINE PERFORMANCE 3-1

4 ENGINE MOUNTING 4-1

5 AIR INLET SYSTEM 5-1

6 FUEL SYSTEM 6-1

7 ELECTRICAL SYSTEM AND MONITORING INSTRUMENTS 7-1

8 LUBRICATION SYSTEM 8-1 9 COOLING REQUIREMENTS 9-1 10 EXHAUST SYSTEM 10-1 11 AIRBLEED SYSTEM 11-1 12 ENGINE ACCESSORIES 12-1 13 COOLANT INJECTION 13-1 14 PROPELLER UNIT 14-1 15 ENGINE CONTROLS 15-1

Appendices: WALTER M601E/E-21 Engine Installation Drawing Appendix 1

Appendix 2 Appendix 3 Appendix 4 Appendix 5 Appendix 6 Appendix 7 Appendix 8

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INSTALLATION MANUAL

MANUAL PART No. 0982502

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

INTRODUCTION

INTRODUCTION

This Installation Manual summarizes the long-term experiences of the WALTER M601 engine manufacturer, i.e. WALTER a.s. It is intended for the airframe manufacturer. Engine installation in the airframe shall be approved by the engine manufacturer.

The Installation Manual includes the requirements, data, and documents approved by the Civil Aviation Authority of the Czech Republic for installation of the engine in the airframe. The Installation Manual summarizes the data on approved power ratings and operation limits, which relate to the installation in the airframe. Neither gas path characteristics, nor design particulars, which are of no significance for the above mentioned purpose, are provided. Should further data for engine installation and approval be needed by the airframe manufacturer these will be supplied on request without delay by the engine manufacturer, i.e. WALTER a.s.

Data compiled in individual sections of this manual are related to original equipment with LUN instruments. Installation of other instruments is aided by enclosed instrument characteristics. Application of other than original equipment shall be approved by the engine manufacturer.

Detailed information on periodic maintenance procedures, adjustment, and repairs is given in the „Maintenance Manual“.

CAUTION

INFORMATION DISCLOSED IN THIS MANUAL, AS WELL AS THE ENCLOSED DRAWINGS AND DIAGRAMS ARE INTENDED FOR DIRECT USE BY PERSONS AND ORGANIZATIONS, TO WHICH THIS MANUAL HAS BEEN CONVEYED EITHER BY THE MANUFACTURER, I.E. WALTER a.s., OR BY AUTHORIZED PERSONS OR ORGANIZATIONS.

REPRODUCTION OR DISCLOSURE OF THIS DOCUMENT, AS WELL AS TRANSFERRING TO FURTHER PERSONS OR ORGANIZATIONS IS NOT PERMITTED, EXCEPT BY WRITTEN PERMISSION FROM THE ENGINE MANUFACTURER.

INFORMATION INCLUDED IN THIS MANUAL AND/OR IN DOCUMENTS OBTAINED AT ADDITIONAL REQUEST BY THE MANUFACTURER OF THE AIRPLANE, MUST BE USED ONLY FOR PURPOSES, (E.G. WORKING OUT OF DESIGN MODIFICATIONS, PRODUCTION OF PARTS, PLACING OF AN ORDER, ETC.) FOR WHICH IT HAS BEEN INTENDED.

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INSTALLATION MANUAL

MANUAL PART No. 0982502

INTRODUCTION

NOTE: IF DRAWINGS/TABLES/DIAGRAMS SHOW DESIGNATION OF AN ENGINE MODEL THEN THEY APPLY TO THE RELEVANT ENGINE MODEL ONLY. IF ENGINE DESIGNATION IS NOT SHOWN THE DIAGRAM IS APPLICABLE TO ALL WALTER M601 ENGINE MODELS DESCRIBED IN THIS MANUAL.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 1

Section 1

ENGINE DESCRIPTION

Type Airworthiness Approval

The WALTER M601E and WALTER M601E-21 engines have been certified by EASA by means of Type Certificate No.: EASA E.070.

DESCRIPTION

The WALTER M601E/E-21 free turbine turboprop engine has been designed for normal, utility, and commuter category airplanes.

The engine is fitted with an injection system which can be utilized for coolant injection for improving flat rating capability of the engine or for compressor recovery washing.

The engine features two independent parts: the gas generator and the propulsor. The gas generator and free turbine shafts are arranged in a tandem layout.

Air enters the engine in the rear part, flows forward through the compressor, combustion chamber, and turbines and exits through exhaust nozzles near the front of the engine.

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INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 1

Gas Generator

Air enters the compressor in a radial direction via a protecting screen and annular plenum. The air is axially directed in front of the compressor. The compressor consists of two axial stages followed by one centrifugal stage. The combustion chamber is of an annular configuration. Part of the primary air enters the combustion compartment through perforations in the walls of the outer flame tube, the remainder passes via the hollow nozzle guide vanes of the gas generator turbine through inner flame tube. The fuel is atomized by a special ring rotating with the gas generator shaft. The one stage gas generator turbine drives the compressor via the compressor shaft. The interturbine temperature is measured by 9 thermocouples installed in the flow path at the gas generator turbine outlet.

Propulsor

The tip shrouded one stage axial-flow turbine drives the propeller via the two-stage countershaft reduction gearbox. The reduction gearbox embodies an integral torquemeter which provides for evaluation of the engine power. The exhaust gases from the free turbine pass through the annular plenum to the atmosphere via two opposed exhaust nozzles. Exhaust gases provide for an additional jet thrust.

Fuel System

The fuel system of the engine is a low pressure system with a gear fuel pump and fuel control unit. In case of the fuel control unit failure it is possible to use (by means of switching the electromagnetic valve on) an emergency circuit of the engine control. Fuel is atomized in the combustion chamber with aid of spray ring. The fuel in the combustion chamber is ignited by means of two torch igniters. Gear pump delivers fuel to the torch igniters.

Engine Starting

The engine is started by an electric starter/generator.

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Section 1

Oil System

The oil system is a circulatory, pressure fed system with an integral oil tank incorporated in the accessory gearbox. The oil system provides lubrication for all areas of the engine; further, pressure oil for the torquemeter and propeller speed governor.

System of Limiters

The engine is equipped with a two-level limiter system. The system of limiters evaluates

the magnitude of monitored engine parameters (nG, nV, torque, ITT and at engine starting

dITT/dt) and optically indicates on the signalling panel the exceeding of these parameters. By limitation of fuel supply the system prevents from exceeding of parameters that can cause the damage of the engine.

Engine Controls

Power plant is controlled by three levers. Engine operation mode at forward-thrust ratings and at BETA range is selected by means of engine control lever. Shut-off valve is controlled by the shut-off valve actuating lever and when the emergency circuit is switched on the lever controls engine operation. Controlled propeller speed and propeller feathering (in front extreme position) is selected by means of propeller control lever.

Engine Mounting

The engine is mounted to the engine mount bed ring by three elastically supported pins which are located on the centrifugal compressor casing.

Engine Accessories

List of all instruments and accessories (including nonstandard equipment) is given in the Section 12, description of the electrical system and operation of individual instruments is given in the Section 7.

Propeller

For propeller specifications refer to Section 14, Propeller Unit. Basic engine equipment provides for emergency feathering by means of the propeller control lever; an airplane can be fitted with system of manual propeller feathering (for all power ratings), actuated by a push button in the cockpit and with autofeathering system which operates in case of engine shutdown at higher power rating.

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Section 1

LEADING PARTICULARS

Engine type: free-turbine, two-shaft tractor turboprop

Performance: refer to Section 3, Table 3-1

Sense of rotation:

(looking forward) gas generator - CCW

propeller turbine - CCW

propeller shaft - CW

Dimensions: max. height 650 mm (25.6 in)

max. width

(exhaust nozzles removed) 590 mm (23.2 in)

max. length (app.) 1,675 mm (66 in)

Weights: refer to Section 4, Table 4-1

Fuel: For approved fuels see „Operation Manual“

Oil: For approved oils see „Operation Manual“

Oil consumption: 0.1 l/hr (0.025 US gal/hr)

Coolant: For approved injected liquids see „Operation Manual“

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Section 2

SECTION 2

WALTER M601E1/E-21 ENGINE OPERATION LIMITS

Operation Limits:

Atmospheric temperature: -50 °C (-58 °F) to +50 °C (104 °F)

Starting at low ambient temperature:

- without preheating:

at oil temperature higher than -20 °C - with preheating:

at oil temperature lower than -20 °C

The preheating can be stopped when oil temperature reaches +5 °C.

Max. gas generator speed: 100 %

(100 % = 36,660 rpm)

Max. propeller speed: 2,080 rpm

Max. interturbine temperature: 735 °C

Max. torque of 2,725 Nm (2,010 lb.ft): 106 %

(100 % = 2,570 Nm = 1,896 lb.ft)

Max. oil temperature: 85 °C

Acceptable gravitational load factors: longitudinal nx = ± 1.5 g

vertical ny = + 4.15 g (+ 5 g short-period at landing, short-period at gust load)

- 2.15 g

lateral nz = ± 1.5 g

ωy = ± 0.45 rad/sec

ωz = ± 0.60 rad/sec

Permitted loads at flight manoeuvres are shown in Fig 4-3, Page 4-9 (for coordinate system refer to Page 4-3).

For further operation limits see the „Operation Manual“.

The engine is approved to operate in severe ice-forming conditions.

If interturbine temperature (ITT), torque or propeller speed exceed their maximum permitted values, it is necessary to evaluate them and proceed in accordance with the applicable Diagrams 2-1, 2-2, 2-3 and Table 2-1 when taking further measures for engine operation.

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Section 1

Area „A“ - 1. Check the condition of the power source (board storage batteries or ground power source).

NOTE: If the fuel is ignited with delay (due to weak battery) the system of limiters cannot prevent the overtemperature when the accumulated fuel is burning.

2. Check the proper function of the limiter system.

Area „B“ - 1. Put the record of the interturbine temperature and the interval of its exceeding in the „Engine Log Book“.

2. Carry out the checks presented as 1. and 2. for Area „A“.

3. Check whether the instructions for starting given in the „Operation Manual“ were respected.

OVERTEMPERATURE LIMITS - STARTING CONDITIONS ONLY Fig. 2-1 770 780 750 760 730 720 740 30 5 0 10 15 20 25

ITT [°C] INTERTURBINE TEMPERATURE

RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

TIME [sec] NO ACTIONS REQUIRED

AREA „A“

AREA „B“

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Section 3

Area „A“ - 1) Enter ITT and time of overtemperature into Engine Log Book. 2) Check total time of overtemperature - it must not exceed 200 min. 3) Find out the fault and rectify cause of overtemperature.

Area „B“ - 1) Enter ITT and time of overtemperature into Engine Log Book. 2) Check total time of overtemperature - it must not exceed 30 min. 3) Find out the fault and rectify cause of overtemperature.

OVERTEMPERATURE LIMITS - ALL CONDITIONS EXCEPT STARTING (Not applicable for max. contingency ratings)

Diagram 2-2a TIME [sec] 780 800 740 760 700 720

ITT [°C] INTERTURBINE TEMPERATURE

RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

NO ACTIONS REQUIRED

0 20 40 60 80 100 120

AREA „A“

AREA „B“

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Section 1

Area „A“ - The use of this ITT is allowed solely in the case of one engine inoperative (OEI) flight at the intermediate contingency rating. The time of its use is limited by the time necessary for finishing the flight. Enter the indicated ITT and time of overtemperature in the Engine Log Bock. Total time in this area must not exceed 200 min during TBO.

Area „B“ - The use of this ITT is allowed solely at the maximum contingency power rating to reach the safe altitude when one engine becomes inoperative at take-off or at aborted landing. Enter the indicated ITT and time of overtemperature in the Engine Log Bock. Total time in this area must not exceed 30 min during TBO.

OVERTEMPERATURE LIMITS – ALL CONDITIONS EXCEPT STARTING (Valid for power ratings defined for the event of OEI flight)

Fig. 2-2b ITT [°C] INTERTURBINE TEMPERATURE

0 1 2 3 4 5 6 7 8 9 10 11 12 TIME [minutes] 780 790 760 770 740 750

RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

NO ACTIONS REQUIRED

AREA „A“

AREA „B“

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Section 3

Area „A“ - Value and time interval of overtorque have to be put in the „Engine Log Book“. Determine the cause and rectify the failure.

NOTE: 100 % torque = 2,570 Nm (1,896 lb.ft) OVERTORQUE LIMITS Diagram 2-3 103 104 109 108 107 106 105 101 102 99 98 100 6 1 0 2 3 4 5 PROPELLER TORQUE[%]

RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

TIME [minutes] 111 110

AREA „A“

AREA „A“

NO ACTIONS REQUIRED

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Section 2

Propeller speed [rpm]

Measures

up to 2,220 No action required

2,220 to 2,300 Overspeed not longer than 20 sec:

Record the rpm in the Engine Log Book. If occurrence of these overspeeds exceeds number 10, the engine must be returned to an overhaul facility for inspection/repair acc. to Overhaul Manual. Overspeed longer than 20 sec:

Ref. to the Propeller Operation Manual

2,300 to 2,400 1) Record the rpm in the Engine Log Book. If occurrence of these

overspeeds exceeds number 2, the engine must be returned to an overhaul facility for inspection/repair acc. to Overhaul Manual. 2) Inspect chip detectors and oil filter cartridge for contamination with

metal chips. Refer to Section 79.10.00, Maintenance Manual. 3) After engine shut-down turn-by propeller manually. Check for

symptoms of power turbine blades seizing (unusual noise). This repeat at 10 min and at 20 min after engine shut-down.

4) Record the results of the check (Item 2) in the Engine Log Book. 5) If the propeller can be manually turned in all three checks without any

symptoms of seizing, the engine can continue in operation for remaining T.B.O. without any limitation.

6) If in one check of these three checks the power turbine blades are in contact with the turbine stator, the engine must be returned to an overhaul facility for inspection/repair acc. to Overhaul Manual.

above 2,400 Return the engine to overhaul facility for inspection/repair acc. to

Overhaul Manual.

PROPELLER OVERSPEED LIMITS Table 2

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Section 2

ENGINE OPERATING ENVELOPE Diagram 2-5 H [m] H [ft] 18,000 20,000 14,000 16,000 10,000 12,000 6,000 8,000 6,000 5,000 4,000 3,000 2,000 1,000 2,000 0 0 0 0.1 0.2 0.3 0.4 0.5 0.6 4,000 Airbleed Envelope

Flight Mach Number In-Flight Start

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Section 3

Section 3

ENGINE PERFORMANCE

This section contains basic information about engine performance. Table 3-1 presents basic power ratings and relevant performance parameters in static conditions, sea level. The following diagrams present engine performance - i.e. shaft power, fuel consumption, gas generator speed, additional thrust and ITT vs altitude and true air speed at take off and maximum continuous rating.

All mentioned values were calculated under the following conditions: - No installation losses but considering ram effect

- No air bleed

- No power off-take from accessory gearbox - Fuel of min. LHV = 42,915 kJ/kg (18,458 BTU/lb)

- Flight altitude corresponds with pressure altitude defined by ISA

Mentioned diagrams were calculated for both ISA conditions and increased OAT to which the engine is flat rated (refer to Table 3-1). Additional diagrams showing take-off engine

parameters with 3rd _{stage of coolant injection applied are presented for the OAT to which}

take-off power is flat rated at sea level. This applies for engine model for which coolant injection is assumed. The OAT is then assumed higher by relevant temperature difference within the whole range of altitudes.

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Section 3

The following nomenclature is used in this section: JAR - Joint Aviation Requirements

FAR - Federal Aviation Regulation

F - jet thrust

ISA - International Standard Atmosphere

(same as the ICAO Standard Atmosphere Conditions)

ITT - interturbine temperature

nG - gas generator speed

LHV - lower heating value OAT - outside air temperature

P - absolute pressure

T - absolute temperature

NH - shaft power

V - velocity relative to undisturbed ambient air (true air speed)

∆ - difference

Θ

- square root of the temperature ratio (engine inlet air temperature to the standard

one in Kelvins or RANKINE degrees)

W - mass flow

Subscripts:

AM - ambient static values of undisturbed air

BL - bleed

F - fuel

G - generator

IN - inlet duct

N - net

TAS - true air speed

C - total

R - corrected

Temperature Conversion Formulae: °F = 1.8 (°C + 40) - 40 °R = °F + 459.67 K = °C + 273.15

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Section 3

Correction on Ambient Temperature Which is Different from ISA Conditions

At an ambient temperature lower than the ISA one, it is always possible to get the shaft power specified for ISA conditions; this is effected at decreased gas generator speed.

At constant gas generator speed the decrease in atmospheric temperature by 1 o_C

(1.8 o_{F), results in the increase of the power by app. 1 % and of the fuel consumption}

by app. 0.7 %. This way increased power can be utilized until limits shown in the Table of Engine Operation Limits are reached (refer to the Operation Manual).

Atmospheric temperature variation shows also an adverse effect: at constant gas

generator speed the ambient temperature increase by 1 o_{C (1.8} o_{F) results in decrease}

of the power by 0.9 % and of the fuel consumption by 0.6 %. The interturbine temperature increases a bit at the same time. When ITT limit for relevant rating is reached it is necessary to slightly decrease gas generator speed. Influence of the atmospheric temperature on the engine parameters is then more apparent. When the

atmospheric temperature increases by 1 o_{C (1.8} o_{F) the power decreases by 1 % and}

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Section 3

Correction on Installation Losses

Engine installation into the aircraft effects the engine performance in two ways: first, the power decreases due to pressure losses in the inlet duct; second, the power decreases due to leakage of exhaust gas in the compressor intake. Entrainment of hot air/gas from the engine compartment results in some warming of air entering the compressor. At reverse engine layout when the inlet duct is situated under the engine the elimination of this effect is very difficult.

a) Pressure Loss Influence:

At power ratings which are characterised by the gas generator speed nG higher

than 90 % pressure loss in the air inlet system ∆P

P _IN

 



 causes the following

changes in the engine parameters: Relative power decrease:

∆N ∆ N 1.95 P P H H = ∗ _IN    

Relative fuel consumption decrease:

∆W ∆ W 0.9 P P F F = ∗ _IN    

Interturbine temperature increase:

[ ]

∆

ITT 100

∆

P

P

_IN

C

=

∗









°

b) Air Warming Influence:

Warming of air at the compressor inlet ∆TIN results in the same effect as the

atmospheric temperature increase. It means that for ∆TIN = 1 °C (1.8 °F) the power

decreases by 0.9 % and the fuel consumption by 0.6 %.

Interturbine temperature increase must be considered if the ITT limit for given power rating can be exceeded. In this case it is necessary to decrease the gas

generator speed. When the gas generator speed nG decreases by 1 % the ITT

decreases app. by 17 o_{C (30.6} o_{F), power at the same time decreases by 6.4 %}

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Section 3

The relative pressure loss ∆P

P _IN  



 and air temperature increase in the air inlet

system ∆TIN with the corrected gas generator speed n_G /

Θ

, as measured on

the L410 aircraft, are shown in the Diagram 3-1. In Diagram 3-2 the changes are plotted in these parameters with the true air speed V. As the engine will be effected by the installation like that, the relative pressure loss and air temperature increase can assumed to be near to the values presented in these diagrams. With the known actual values of the pressure loss and the air temperature increase in the inlet duct for a different layout of the air inlet duct, the influence on the engine parameters can be calculated according to the above presented equations or according to Diagram 5-3, in the Section 5 „Air Inlet System“.

Correction on Airbleed at Compressor Outlet:

Diagram 3-3 shows the airflow rate, that can be bled for the aircraft needs from the compressor of the relevant engine model. The influence of air bleed on the engine proper can be simulated by an equivalent throttling orifice, installed at the airbleed manifold delivery flange where the bled air is let to flow free in the ambient atmosphere. Diagram 3-3 shows pressure available at the airbleed manifold delivery flange with the

gas generator speed and given airbleed flow rate WBL. All values are presented for ISA

sea level, static conditions. Nevertheless it is possible to recalculate them for any inlet conditions in the usual way.

First, calculate the total pressure PIN and the total temperature TIN at the compressor

inlet. In metric units, temperature in Kelvins, and flight speed in km/hr, the inlet total temperature is obtained:

T

T

V

3.6

1

2009

C AM 2

=

+









∗

[ K ] T_IN = T_C + ∆T_IN [ K ]

P

P

T

T

C AM C AM 3.5

=

∗













[ kPa ]

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Section 3

P P 1 P P IN C IN = ∗ −           ∆ [ kPa ] Where:

V = true air speed [ km/hr ]

TAM = outside air temperature [ K ]

PAM = ambient pressure [ kPa]

T_C = total temperature [ K ]

P

C = total pressure [ kPa ]

∆TIN = air temperature increase in the air inlet system [ K ]

∆P P _IN  



 = relative pressure loss in the air inlet system

TIN = total temperature at the compressor inlet [ K ]

PIN = total pressure at the compressor inlet [ kPa ]

The values of ∆TIN and ∆P

P _IN  



 in flight conditions can be determined by means of

Diagram 3-2 for relevant engine model. At zero flight speed at the first step we calculate the value of n_G /

Θ

with TAM instead of TIN. Then we read in the Diagram

3-1 the value of ∆TIN and we can calculate TIN = TAM + ∆TIN.

At the second step we calculate new value of n_G/

Θ

using this TIN. Then we can

read in the Diagram 3-1 new value of ∆TIN. With this value we calculate the final values

of TIN and n_G/

Θ

and then the value of ∆P

P _IN  



 can be read in the Diagram 3-1.

When calculated in the U.S. standard units, the above relations are to be written as follows:

T

T

V

1.9438

1

1116

C AM 2

=

+









∗

[ °R ]

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

T_IN = T_C + ∆T_IN [ °R ]

P

P

T

T

C AM C AM 3.5

=

∗













[ psi ] P P 1 P P IN C IN = ∗ −           ∆ [ psi ] Where:

V = true air speed [ kt ]

TAM = outside air temperature [ °R ]

PAM = ambient pressure [ psi ]

TC = total temperature [ °R ]

PC = total pressure [ psi ]

∆TIN = air temperature increase in the air inlet system [ °F ]

∆P P _IN  



 = relative pressure loss in the air inlet system

TIN = total temperature in the compressor inlet [ °R ]

PIN = total pressure in the compressor inlet [ psi ]

Calculate the corrected gas generator speed n_G /

Θ

[ % ]

In the metric system of units: n_G /

Θ

= n_G 288

T_IN [ % ]

In the U.S. standard units: n_G /

Θ

= n_G 519

T_IN [ % ]

For known values of n_G /

Θ

, PIN, TIN, the demanded airbleed flow rate WBL, the

pressure of air at the airbleed delivery flange and the equivalent diaphragm dia can be found in Diagram 3-3. Temperature of bled air can be estimated as follows:

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

In the metric system of units:

T

T

273.8

n

100

282

BL IN G 2

=

+

∗









−

[ °C ] Where TIN is in Kelvins, nG is in %.

In the U.S. standard units:

T T 492.8 n 100 475.6 BL IN G 2 = + ∗     − [ °F ]

Where TIN is in Rankine degrees, nG is in %.

The engine response to airbleed is as follows:

At constant gas generator speed nG the shaft power does not change but there is an

increase in fuel consumption and ITT. This increase can be determined from the curves in Diagram 3-4, where the difference in fuel consumption with airbleed open/closed

∆W W F F _BL       

 and the corresponding ITT increase (∆ITT)BL are shown as a function of

equivalent diaphragm dia. When the ITT exceeds the limit for the given rating it is necessary to decrease the gas generator speed. The relation between the 1 %

decrease in gas generator speed nG and the ITT drop by app. 17 oC (30.6 oF), power

decrease by 6.4 % and fuel consumption by 4.4 % holds true.

Correction on Power Off-Take for Electric Generator.

Engine response to electric generator loading is an ITT increase. At ratings defined by

gas generator speed within nG = 90 to 100 % the ITT increases by 5 oC for each

100 A of generator loading. Should the ITT increase exceed the limit for the given rating, it is necessary to decrease the gas generator speed.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

Correction on Coolant Injection

This paragraph applies only to engine models with coolant injection. For procedure refer to Para 6, Standard Practices, Operation Manual. Further information is presented in Section 13, Installation Manual and in Section 82.00.00, Maintenance Manual.

Coolant injection into compressor inlet can be used for short time power augmentation at take-off rating. Quantity of injected coolant into compressor inlet should be in accordance with the required of constant shaft power at higher temperatures. In fact it is possible to select one of three available rates of coolant injection according to the ambient temperature. The selection of the coolant injection rate is specified in the „Operation Manual“ (see Fig. 2-2) in accordance with the ambient temperature and pressure.

At the highest coolant flow rate, at constant gas generator speed nG the shaft power

increases by at least 10 %. At the same time the fuel consumption increases by app.

10 % and the interturbine temperature decreases by app. 10 o_{C (18} o_{F). After the}

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

rating shaft power equivalent shaft power

ESFC max. gas generator speed propeller speed torque max. interturbine temperature [kW] [SHP] [ESHP] [kW] [lb/ESHP/hr][g/kW/hr] GT [%] [rpm] [N.m] [lb.ft] [°C] take-off (5 min) 15 °C (59.0 °F) 560 751 595 798 395 0.6493 98.6 2080 2570 1895 710 sea level static

23 °C (73.4 °F) 560 751 595 798 - 100 2080 2570 1895 735 max. continuous 15 °C (59 °F) 490 657 521 699 410 0.674 96.5 1700 to 2080 2570 1895 680 sea level static

18 °C (64.4 °F) 490 657 521 699 - 97 1700 to 2080 2570 1895 690 take-off with water injection

300 l/hr (79 US gal/hr) (5 min) 97.325 kPa (14.12 psi) 33 °C (91.4 °F) 560 751 595 751 - 100 2080 2570 1895 735 intermediate contingency sea level static

28 °C (82 °F) 560 751 595 798 - 100.5 2080 2570 1895 760 maximum contingency (10 min) 97.325 kPa (14.12 psi) 28 °C (82.4 °F) 595 798 630 845 - 102 2080 2737 2019 780

NOTE: gas generator speed 100 % = 36,660 rpm

2,080 propeller rpm = 31,023 power turbine rpm

ENGINE POWER RATINGS ACC. TO JAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

rating shaft power equivalent shaft power

ESFC max. gas generator speed propeller speed torque max. interturbine temperature [kW] [SHP] [ESHP] [kW] [lb/ESHP/hr][g/kW/hr] [%] GT [rpm] [N.m] [lb.ft] [°C] take-off (5 min) 15 °C (59.0 °F) 560 751 595 798 389 0.64 98.1 2080 2570 1895 690 sea level static

28 °C (82.4 °F) 560 751 595 798 - 100 2080 2570 1895 735 max. continuous 15 °C (59 °F) 490 657 521 699 405.9 0.667 96.2 1700 to 2080 2570 1895 660 sea level static

21 °C (69.8 °F) 490 657 521 699 - 97 1700 to 2080 2570 1895 690 take-off with water injection

300 l/hr (79 US gal/hr) (5 min)

sea level static 42 °C (107.6 °F) 560 751 595 751 - 100 2080 2570 1895 735 intermediate contingency sea level static

32 °C (90 °F) 560 751 595 798 - 100.5 2080 2570 1895 760 maximum contingency (10 min) 97.325 kPa (14.12 psi) 31.5 °C (88.7 °F) 595 798 630 845 - 102 2080 2737 2019 780

NOTE: gas generator speed 100 % = 36,660 rpm

2,080 propeller rpm = 31,023 power turbine rpm

ENGINE POWER RATINGS ACC. TO JAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

rating shaft power equivalent shaft power

ESFC max. gas generator speed propeller speed torque max. interturbine temperature [kW] [SHP] [ESHP] [kW] [lb/ESHP/hr][g/kW/hr] GT [%] [rpm] [N.m] [lb.ft] [°C] take-off (5 min) 15 °C (59.0 °F) 560 751 595 798 395 0.6493 98.6 2080 2570 1895 710 sea level static

23 °C (73.4 °F) 560 751 595 798 - 100 2080 2570 1895 735 climb and max. cruise

15 °C (59 °F) 490 657 521 699 410 0.674 96.5 1700 to 2080 2570 1895 680 sea level static

18 °C (64.4 °F) 490 657 521 699 - 97 1700 to 2080 2570 1895 690 take-off with water injection

300 l/hr (79 US gal/hr) (5 min) 97.325 kPa (14.12 psi) 33 °C (91.4 °F) 560 751 595 751 - 100 2080 2570 1895 735 max. continous sea level static 28 °C (82 °F) 560 751 595 798 - 100.5 2080 2570 1895 760 Max. take-off (5 min) 97.325 kPa (14.12 psi) 28 °C (82.4 °F) 595 798 630 845 - 102 2080 2737 2019 780

NOTE: gas generator speed 100 % = 36,660 rpm

2,080 propeller rpm = 31,023 power turbine rpm

ENGINE POWER RATINGS ACC. TO FAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

rating shaft power equivalent shaft power

ESFC max. gas generator speed propeller speed torque max. interturbine temperature [kW] [SHP] [ESHP] [kW] [lb/ESHP/hr][g/kW/hr] [%] GT [rpm] [N.m] [lb.ft] [°C] take-off (5 min) 15 °C (59.0 °F) 560 751 595 798 389 0.64 98.1 2080 2570 1895 690 sea level static

28 °C (82.4 °F) 560 751 595 798 - 100 2080 2570 1895 735 Climb and max. cruise

15 °C (59 °F) 490 657 521 699 405.9 0.667 96.2 1700 to 2080 2570 1895 660 sea level static

21 °C (69.8 °F) 490 657 521 699 - 97 1700 to 2080 2570 1895 690 take-off with water injection

300 l/hr (79 US gal/hr) (5 min)

sea level static 42 °C (107.6 °F) 560 751 595 798 - 100 2080 2570 1895 735 Max. continous sea level static 32 °C (90 °F) 560 751 595 798 - 100.5 2080 2570 1895 760 Max take-off (5 min) 97.325 kPa (14.12 psi) 31.5 °C (88.7 °F) 595 798 630 845 - 102 2080 2737 2019 780

NOTE: gas generator speed 100 % = 36,660 rpm

2,080 propeller rpm = 31,023 power turbine rpm

ENGINE POWER RATINGS ACC. TO FAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TEMPERATURE INCREASE AND RELATIVE PRESSURE LOSS IN THE AIR INLET SYSTEM.

STATIC GROUND OPERATION OF AN ENGINE INSTALLED IN THE L 410 UVP AIRPLANE NACELLE.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TEMPERATURE INCREASE AND RELATIVE PRESSURE LOSS IN THE AIR INLET SYSTEM.

IN-FLIGHT MEASUREMENT. ENGINE INSTALLED IN THE L410 UVP AIRPLANE. Diagram 3-2

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

AIR PRESSURE AT THE AIR BLEED MANIFOLD DELIVERY FLANGE. VARIATION WITH THE AIR BLEED FLOW RATE AND GAS GENERATOR SPEED.

Diagram 3-3

pBL

= 14 696/p

IN

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

REL. FUEL FLOW RATE AND INTERTURBINE TEMPERATURE INCREASE VARIATION WITH AIR BLEED, EQUIVALENT THROTTLING ORIFICE DIAMETER

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 8 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 8 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 3 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 3 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING WITH COOLANT INJECTION

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 20.2 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING WITH COOLANT INJECTION NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 20.2 °C CONDITIONS.

WALTER M601E

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 13 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 13 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 6 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

MAX. CONTINUOUS RATING RATING acc. to JAR CLIMB AND MAX CRUISE acc. to FAR

NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 6 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING WITH COOLANT INJECTION

SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 27 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 3

TAKE-OFF RATING WITH COOLANT INJECTION NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 27 °C CONDITIONS.

WALTER M601E-21

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

SECTION 4

ENGINE MOUNTING

General

Conditions for mounting the engine to the airframe structure as described within Section 4 should be strictly followed.

Mounting System Arrangement

The engine is supported by three mounts located on the centrifugal compressor case. Mounting pad locations are shown in Figure 4-1 (schematic). Each mounting pad is numbered for the purpose of easy interpretation.

Reactions from engine mount bed must be taken in point R (refer to Figure 4-2). The perpendicular distance between this point and mounting pad plane must not exceed 45 mm (1.77 in).

Maximum bending moments acting upon the pads must not exceed 900 Nm (7,820 lb.in).

Mounting pad stud nut tightening torque must not exceed 26.5 Nm (235 lb.in). Centering shoulder of mount body must be engaged into the mounting pad at least 2 mm in depth (0.08 in) - refer to Figure 4-2.

Distance between propeller centre of gravity and propeller flange must not exceed 110 mm (4.33 in).

Engine vibrations are insulated from aircraft structure by elastic supports. Mount Loads

Forces acting upon individual mounts must be defined before the installation of the engine to the airframe. These load factors must be included in the analysis:

1. Gravitational and inertial forces from the engine including accessories and propeller. 2. Aerodynamic forces from the propeller.

3. Gyroscopic moments induced by angular movements. 4. Acceleration moments of rotating parts.

Gyroscopic moments induced by angular movements of the gas generator and power turbine rotors can be neglected with respect to the gyroscopic moment of the propeller. Allowable loads during take-off, flight, and landing are shown in Figure 4-3.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

Engine Mass and Dynamic Characteristics

Engine mass and mass moments of inertia about the center of gravity are listed in Tables 4-1 and 4-2. The location of the center of gravity is shown on Sheet 1, Sector E8. Accessory mounting pads and the center of gravity projection for respective accessories are shown on the same drawing. The center of gravity projection together with the mass of the accessories causes overhang moments acting upon the pads of casing. The values of overhang moments of accessories listed in Section 12 are max. allowable values. In case of using the accessories exceeding the limits, the relevant details should be submitted to engine manufacturer WALTER a.s. and asked for approval.

Engine dry mass 207 kg 456 lb

Mass of oil charge 7 kg 15.4 lb

Table 4-1

NOTE:

The following standard equipment delivered with the engine is included in the mass: Fuel pump, fuel control unit, starter/generator, ignition unit, integrated speed transmitters, oil temperature transmitter, interturbine transmitter, oil pressure transmitter, fuel pressure transmitter, three torquemeter transmitters, min. oil pressure switch, min. oil quantity signaller, engine mounts (three pieces).

Additional equipment mass:

Exhaust nozzles L.H. (M601-418.7) and R.H. (M601-419.7) 2.8 kg (6.2 lb)

Alternator gearbox 2.5 kg (5.5 lb)

LUN 2102 Alternator 9.8 kg (21.6 lb)

Engine mounting ring 5.16 kg (13.5 lb)

Actual engine mass and relevant C.G. position of the dry engine with equipment as required by customer are shown in the Engine Log Book.

The actual engine mass and C.G. position do not respect weights of oil and fuel pressure transmitters.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

Polar mass moment of inertia about axis x

3.4 kgm2 _{11,618 lb in}2

Mass moment of inertia about axis y

38 kgm2 _{130,000 lb in}2

Mass moment of inertia about axis z

38 kgm2 _{130,000 lb in}2

THE ENGINE MASS MOMENTS OF INERTIA Table 4-2

Coordinate system of the engine (right handed) and positive sense of rotation

Propeller mass and dynamic characteristics are shown in the Section 14 Propeller Unit. y

x

z

+ ωy; εy

+ ωz;εz + ωx;εx

Axis of the engine Direction of flight

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

Vibration

Vibration load of the engine is evaluated by effective vibration speed VEF (mm/s) or

VRMS (in/s). Each engine is submitted to the check on vibration in the manufacturing

plant. Vibration is measured at steady-state condition after 2 to 3 minutes of engine operation. Vibration sensors are located in a radial direction as follows (Refer to Figure 4-1):

a) on the reduction gearbox (Designation Sn 1), where the first harmonic component of power-turbine speed is measured.

b) on the accessory gearbox (Designation Sn 3), where the first harmonic component of generator speed is measured.

SENSOR Sn 1 Sn 3 Test bed _V EF (mm/sec) VRMS (in/sec) VEF (mm/sec) VRMS (in/sec) Propeller test bed 10 0.4 5 0.2

VIBRATION LOAD LIMITS Table 4-3

Mentioned check of the vibration limits after engine installation and during operation is not necessary.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

Checking Mount Pad Load

Forces and couples acting on the engine's center of gravity at flight manoeuvres where some parameters exceed the limit given in Fig. 4-3 must satisfy the following relationships:

Mount Pad No. 1

Mount Pad No. 2

Mount Pad No. 3

Where

X, Y, Z [N] ... forces actuating on the engine center of gravity

(see Installation Drawing, Sheet 1, Sector E8) Mxy, Myz, Mzx [Nm] ... couples acting on coordinate planes

R [m] ... pitch circle radius of hinged pin holes

TM ... engine center of gravity

XTM [m] ... distance of the engine center of gravity from spigot entry axis

in mount pads

(see Installation Drawing, Sheet 1, Sector E8)

C = 19,500 [N] ... limit load (FAR 33.23 (b)(1))

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

For the U.S. standard units the following values shall apply

X, Y, Z [lb] ... forces acting on the engine center of gravity

(see Installation Drawing, Sheet 1, Sector E8) Mxy, Myz, Mzx [lb.in] ... couples acting on coordinate planes

R [in] ... pitch circle radius of hinged pin holes

TM ... engine center of gravity

XTM [in] ... distance of the engine center of gravity from spigot entry

axis in mount pads

(see Installation Drawing, Sheet 1, Sector E8)

C = 4,384 [lb] ... limit load (FAR 33.23 (b)(1))

C = 6,576 [lb] ... ultimate load (FAR 33.23 (b)(2))

The number of mount pads and the coordinate system with forces acting in positive direction are shown diagrammatically in Fig. 4-1.

An excessive limit value obtained for any mount pad must be approved by the engine manufacturer.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

SCHEME SHOWING LOCATION OF ENGINE MOUNT PADS AND COORDINATE SYSTEM WITH FORCES AND COUPLES ACTING IN POSITIVE DIRECTION

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

FITTING THE ENGINE MOUNT BODY TO THE MOUNTING PAD (SCHEMATIC DIAGRAM) Fig. 4-2

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 4

(Coordinate system - Page 4-3)

(a) nx, ny, nz [g] gravitational load factors

ωx, ωy, ωz [rad/sec] angular velocities

εx, εy, εz [rad/sec2] angular accelerations

(b) Permitted loads are related to engine center of gravity

PERMITTED LOADS AT FLIGHT MANOEUVRES Fig. 4-3

n

y

n

y 2 1 5 4 3 -1 -2 -1 -2 1 2

n

x 2 1 5 4 3 -1 -2 -1 1

n

x

Landing

(engine idling) n_{z = ± 1.5} ω_y = 0 ω_{z = 0} ε_x = ± 3 ε_{y = ± 10} ε_{z = ± 10}

Flight 1

(idling to take-off power) n_{z = ± 1.5} ω_{y = ± 0.3} ω_{z = ± 0.6} ε_{x = ± 4} ε_{y = ± 6} ε_{z = ± 6}

Flight 2

(idling to take-off power) n_{z = ± 1.5} ω_{y = ± 0.45} ω_{z = ± 0.6} ε_{x = ± 4} ε_{y = ± 6} ε_{z = ± 6}

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

SECTION 5

AIR INLET SYSTEM

The purpose of the air inlet system is to deliver air into the compressor with both minimum pressure loss and minimum warning up over a wide range of operational conditions. This is required to obtain required shaft power levels at specified consumption.

The WALTER M601E model engine has been substantiated to the surge and stall requirements of FAR 33.65, induction system icing requirements of FAR 33.68, and the foreign object ingestion of FAR 33.77 fitted with the B 062350 inlet duct. This induction system meets mentioned requirements for all WALTER M601 engine models.

The B 062350 inlet duct is a part of the engine nacelle designed by airframe manufacturer LZ Aeron. Industries, inc. Installations powered by this engine model must incorporate this duct or an equivalent inlet duct with an integral protective device. This equivalent inlet design may not introduce distortion in excess of that of the B 062350 inlet duct design. An inlet plenum chamber must be tight enongh to prevent from warm air leaking from engine compartment. This must be ensured within the whole operating period of the installed engine. Engine manufacturer should be consulted regarding the details of the aerodynamic and structural requirements of the inlet duct system.

Description

The main components of the air inlet system are the inlet lip and throat area, the diffuser duct, and the plenum.

Air enters the compressor from a plenum chamber through a protecting air inlet screen. The plenum-type intake allows considerable versatility in the position and orientation of the air inlet system.

Inlet Lip and Throat Area

It is necessary to extend the inlet lip as far forward as possible in order to obtain uniform pressure and velocity distribution of air. It is recommended to ensure that the entering air is not disturbed by the nacelle boundary layer. The inlet lip should be located in the relatively uniform stream outside the macroturbulent zone at the propeller blade roots. Great attention should be paid to the geometric design of the lip to prevent undesired flow separation. The velocity of entering air at take-off condition should not

exceed 40 m/sec (130 ft/sec). Throat flow area greater than 0.078 m2 _{(121 in}2_{) meets}

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INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

Inlet Diffuser Duct

The diffuser transports air from the inlet lip to the inlet plenum and decreases the velocity of the stream. Mean velocity of the stream before the inlet plenum should not

exceed 25 m/sec (82 ft/sec). Corresponding outlet area of the diffuser is app. 0.120 m2

(186 in2_).

Inlet Plenum Chamber

The inlet plenum chamber completely surrounds the outside of the air inlet screen. The compartment of the inlet plenum is bound in axial direction by front and rear air bulkheads. It is recommended that the plenum (nacelle) wall should have a minimum

clearance of 60 mm (2.3 in) at a point of 180o _{opposite to the transfer duct entrance.}

This clearance should gradually increase to a minimum of 120 mm (4.7 in) for 90o _on

either side of this point towards the diffuser duct entrance. Air Inlet Screen

Air enters the compressor through a protecting air inlet screen which protects the engine from ingestion of large particles. As the screen area is very large and entering air stream velocity low, air inlet screen pressure loss is very small.

Air Inlet Obstructions

Various lines and struts located in the inlet system might disturbe the inlet flow pattern and influence adversely the performance of the engine. It is necessary to minimize adverse effects of wakes caused by various structure components. These should be kept as thin a practical and with suitable profile aligned with the direction of flow. Air coolers, fluid pumps, and fluid lines should not be installed in the engine inlet air stream. Such equipment increases the possibility of toxic contamination of bleed air. Engine Anti-Icing Protection

In order to provide the aircraft with an all-weather capability, the power plant installation must include an anti-icing system (see Fig. 5-1) The leading edge of the inlet lip and, if necessary further components of the air inlet system structure, which might considerably affect the stream of air entering the compressor, should incorporate suitable means to prevent icing. In order to prevent ingestion ice particles and/or other foreign objects it is recommended that the design of the air inlet system should be based on a sudden turn in the air-stream. Ice particles, due to their greater momentum cannot follow the sudden turn of the air-stream. Therefore the probability of their ingestion into the compressor is considerably decreased.

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INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

In order to improve the particle separation effect, the air inlet duct should be equipped with a vane. The extended vane deflects the air-stream and in addition increases its velocity. The resultant increase in particle momentum decreases the probability of their ingestion. A by-pass vane allowing the air with foreign objects to be discharged overboard should be opened at the same time. The system of vanes is put into operation only when icing conditions occure. Small particles of ice which could pass to the compressor inlet screen cannot considerably affect the performance of the engine. Installation in Agricultural Aircraft

Operation in agriculture requires more efficient protection of the compressor against ingestion of foreign objects. In addition to the dust particles the separation system must ensure the separation of sprayed agent aerosols. The quantity of fine dust particles is such that the protection based only on the sudden turn in the air-stream is not sufficient. The inlet air system should be provided with an efficient filtering device. If this filtering device cannot capture the aerosol, the air inlet system should be fitted with a special filter for this purpose. The most advantageous location for the latter is the inlet to the plenum chamber (Refer to Fig. 5-2).

Influence of Air Inlet System on Performance

The air inlet system can effect the performance of the engine by two basic factors: warming of entering air from hot parts of the engine and admission of leaking hot air; pressure loss of air inlet system.

Warming results in increased temperature at the compressor inlet. Thus the performance of the engine should be defined with respect to this temperature. For procedure refer to the Section 3, Correction on Installation Losses.

Direct influence of air inlet system pressure loss on power, fuel consumption and interturbine temperature can be seen from curves on Fig. 5-3.

WALTER a.s.

INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

AIR INLET SYSTEM Fig. 5-1 1 - De-icing vane 2 - Vane 3 - Vane 4 - Screen 5 - Oil cooler a - air inlet

b - air path (icing mode)

c - air path at normal operation (non icing mode) d - engine air inlet compartment

e - air flow into oil cooler f - ice particles

h - air outlet from the oil cooler

k - vane positions at normal operation (non icing mode) m - vane positions (icing mode)

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INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

AGRICULTURAL AIRPLANE ENGINE AIR INLET SYSTEM Fig. 5-2

1 - Air Filter

2 - Aerosol Separator (Demister)

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INSTALLATION MANUAL

MANUAL PART No. 0982502

Section 5

ENGINE SHAFT POWER, FUEL FLOW RATE AND INTERTURBINE TEMPERATURE AT CONSTANT GAS GENERATOR SPEED.

VARIATION WITH AIR INLET SYSTEM PRESSURE LOSS Fig. 5-3

∆ difference

WF fuel flow rate

NH shaft power

∆ITT interturbine temperature increase (°C)

∆P difference between P0 - P1

P0 ambient total pressure

P1 total pressure at compressor protecting screen inlet

1 W W F F + ∆ 1 N N H H + ∆ ∆P P₀ IN         1.00 0.98 0.96 0.94 0.01 0.02 0.03 0 ∆ITT (°C) 1.00 0.98 0.96 4 2 0