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  2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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. REVISION No. 1

2

BULLETIN No.

ISSUE DATE OF NEW PAGES

NUMBERS OF AFFECTED PAGES

DATE OF INSERTION AND SIGNATURE

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 May 11, 2006 Pages 0-1, 0-3, 0-4, 1-1, 4-2, 7-23, 12-6, App. 7

RECORD OF REVISIONS Page 0-1 May 11, 2006

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 REVISION No.

BULLETIN No.

ISSUE DATE OF NEW PAGES

RECORD OF REVISIONS Page 0-2 Jun 26, 2001

NUMBERS OF AFFECTED PAGES

DATE OF INSERTION AND SIGNATURE

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 REVIEW OF EFFECTIVE PAGES - SHEETS Section

Page - Sheet

Title Sheet - in English - in Czech

Section

Date Jun 26, 2001 26.6.2001

Record of revisions

0-1 0-2

May 11, 2006 Jun 26, 2001

Review of effective sheets - pages

0-3 0-4

May 11, 2006 May 11, 2006

Contents

0-7 0-8

Jun 26, 2001 Jun 26, 2001

Section 1 Engine Description

1-1 1-2 1-3 1-4

Nov 15,2004 Jun 26, 2001 Jun 26, 2001 Jun 26, 2001

2-1 2-2 2-3 2-4 2-5 2-6 2-7 2-8 blank

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3-1 3-2 3-3 3-4 3-5 3-6 3-7 3-8 3-9 3-10 3-11 3-12 3-13 3-14

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Section 3 Engine Performance

Date

Section 3 (continued)

3-15 3-16 3-17 3-18 3-19 3-20 3-21 3-22 3-23 3-24 3-25 3-26 3-27 3-28 3-29 3-30 3-31 3-32 3-33 3-34 3-35 3-36 3-37 3-38...blank

Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004 Nov 15,2004

Section 4 Engine Mounting

4-1 4-2 4-3 4-4 4-5 4-6 4-7 4-8 4-9 4-10 blank

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5-1 5-2 5-3 5-4 5-5 5-6 5-7 5-8

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Introduction

Section 2 Engine Operation Limits

Page - Sheet

Section 5 Air Inlet System

REVIEW OF EFFECTIVE PAGES - SHEETS Page 0-3 May 11, 2006

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Section Section 6 Fuel System Section 7 Electrical System and Monitoring Instruments

Section 8 Lubrication System

Section 9 Cooling Requirements Section 10 Exhaust System Section 11 Airbleed System

Page - Sheet

Date

6-1 6-2

Jun 26, 2001 Jun 26, 2001

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

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8-1 8-2 8-3 8-4 blank

Jun 26, 2001 Jun 26, 2001 Nov 15,2004 Jun 26, 2001

9-1 9-2

Nov 15,2004 Nov 15,2004

Section Section 12 Engine Accessories

Section 13 Coolant Injection Section 14 Propeller Unit

Section 15 Engine Controls

Appendix

10-1 Jun 26, 2001 10-2 blank Jun 26, 2001 11-1 Jun 26, 2001 11-2 blank Jun 26, 2001

REVIEW OF EFFECTIVE PAGES - SHEETS Page 0-4 May 11, 2006

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Date

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14-1 14-2 14-3 14-4 blank

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1 2 3 4 5 6 7 8

Jun 26, 2001 Jun 26, 2001 Jun 26, 2001 Jun 26, 2001 Jun 26, 2001 Jun 26, 2001 May 11, 2006 Jun 26, 2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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

CONTENTS Page 0-5 Jun 26, 2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502

CONTENTS Page 0-6 Jun 26, 2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.

INTRODUCTION Page 0-7 Jun 26, 2001


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.

INTRODUCTION Page 0-8 Jun 26, 2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.

Section 1 Page 1-1 May 11, 2006

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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. Fuel in the torch igniters is ignited by means of low voltage plugs.

Section 1 Page 1-2 Jun 26, 2001

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.



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“

Airbleed:

refer to Section 11


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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 % = 36,660 rpm)

100 %

Max. propeller speed:

2,080 rpm

Max. interturbine temperature:

735 °C

Max. torque of 2,725 Nm (2,010 lb.ft): (100 % = 2,570 Nm = 1,896 lb.ft)

106 %

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.

Section 2 Page 2-1 Nov 15,2004

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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. ITT [°C] INTERTURBINE TEMPERATURE

780

770

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

760

750

740

AREA „B“

730

AREA „A“ NO ACTIONS REQUIRED

720 5

0

10

15

20

30

25

TIME [sec] OVERTEMPERATURE LIMITS - STARTING CONDITIONS ONLY Fig. 2-1

Section 1 Page 2-2 Nov 15, 2004


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.

ITT [°C] INTERTURBINE TEMPERATURE RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

800

780

AREA „B“ 760

AREA „A“ 740

720

NO ACTIONS REQUIRED

700 0

20

40

60

80

100

120 TIME [sec]

OVERTEMPERATURE LIMITS - ALL CONDITIONS EXCEPT STARTING (Not applicable for max. contingency ratings) Diagram 2-2a


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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. ITT [°C] INTERTURBINE TEMPERATURE RETURN THE ENGINE TO AN OVERHAUL FACILITY FOR INSPECTION/REPAIR ACC. TO OVERHAUL MANUAL

790

780

AREA „B“

770

760

750

AREA „A“ 740

730 NO ACTIONS REQUIRED

0

1

2

3

4

5

6

7

8

9

10

11

12

TIME [minutes] OVERTEMPERATURE LIMITS – ALL CONDITIONS EXCEPT STARTING (Valid for power ratings defined for the event of OEI flight) Fig. 2-2b


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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) PROPELLER TORQUE [%] 111 110

AREA „A“

109 108

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

107 106 105 104 103

AREA „A“

102 101 100 99

NO ACTIONS REQUIRED

98 0

1

2

3

4

OVERTORQUE LIMITS

5

6 TIME [minutes]

Diagram 2-3



Propeller speed

Measures

[rpm] 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


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 H [m] 6,000

H [ft]

Airbleed Envelope

20,000

18,000 5,000

16,000

14,000

In-Flight Start

4,000 12,000

3,000

10,000

8,000 2,000 6,000

4,000 1,000 2,000

0

0 0

0.1

0.2

0.3

0.4

0.5

0.6

Flight Mach Number ENGINE OPERATING ENVELOPE Diagram 2-5




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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 W

- square root of the temperature ratio (engine inlet air temperature to the standard one in Kelvins or RANKINE degrees) - 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


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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 oC (1.8 oF), 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 oC (1.8 oF) 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 oC (1.8 oF) the power decreases by 1 % and the fuel consumption by 0.66 %.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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  ∆P  causes the following than 90 % pressure loss in the air inlet system    P  IN changes in the engine parameters:

Relative power decrease:

∆NH  ∆P  = 1.95 ∗   NH  P  IN Relative fuel consumption decrease:

∆W F  ∆P  = 0.9 ∗   WF  P  IN Interturbine temperature increase:

 ∆P  ∆ITT = 100 ∗    P  IN b)

[° C]

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 oC (30.6 oF), power at the same time decreases by 6.4 % and fuel consumption by 4.4 %.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502  ∆P  and air temperature increase in the air inlet The relative pressure loss    P  IN system ∆TIN with the corrected gas generator speed nG / 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: 2

TC = TAM

1  V  + ∗  2009  3.6 

TIN = TC + ∆TIN

PC = PAM

T  ∗ C   TAM 

[K] [K]

3.5

[ kPa ]



  ∆P   PIN = PC ∗ 1 −      P  IN 

[ kPa ]

Where: V

= true air speed

[ km/hr ]

TAM

= outside air temperature

[K]

PAM

= ambient pressure

[ kPa]

TC

= total temperature

[K]

PC

= 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 ]

 ∆P  The values of ∆TIN and   in flight conditions can be determined by means of  P  IN Diagram 3-2 for relevant engine model. At zero flight speed at the first step we calculate the value of nG / Θ with TAM instead of TIN. Then we read in the Diagram 31 the value of ∆TIN and we can calculate TIN = TAM + ∆TIN. At the second step we calculate new value of nG/ 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  ∆P  of TIN and nG/ Θ and then the value of   can be read in the Diagram 3-1.  P  IN When calculated in the U.S. standard units, the above relations are to be written as follows: 2

TC = TAM


1  V  + ∗  1116 1.9438 

[ °R ]

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 TIN = TC + ∆TIN

PC = PAM

T  ∗ C   TAM 

[ °R ] 3.5

[ psi ]

  ∆P   PIN = PC ∗ 1 −      P  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

nG / Θ

In the metric system of units:

nG / Θ = nG

288 TIN

[%]

In the U.S. standard units:

n G / Θ = nG

519 TIN

[%]

[%]

For known values of nG / 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:



In the metric system of units:

TBL

n  = TIN + 273.8 ∗  G  100 

2

− 282

[ °C ]

Where TIN is in Kelvins, nG is in %. In the U.S. standard units:

TBL

n  = TIN + 492.8 ∗  G  100 

2

− 475.6

[ °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

 ∆WF     WF  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 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 oC (18 oF). After the coolant injection has been finished ITT returns to the initial level.



shaft power rating [kW] [SHP]

ESFC max. gas equivalent generator shaft speed power [kW] [g/kW/hr] GT [ESHP] [lb/ESHP/hr] [%]

take-off (5 min)

560

595

395

15 °C (59.0 °F)

751

798

0.6493

sea level static

560

595

23 °C (73.4 °F)

751

798

max. continuous

490

521

410

15 °C (59 °F)

657

699

0.674

sea level static

490

521

18 °C (64.4 °F)

657

699

560

595

751

751

560

595

-

98.6

propeller speed

torque

[rpm]

[N.m] [lb.ft]

2080

2570

max. interturbine temperature [°C]

710

1895 100

2080

2570

735

1895

-

96.5

97

1700 to 2080 1700 to 2080

2570

680

1895 2570

690

1895

take-off with water injection 300 l/hr (79 US gal/hr) (5 min) 97.325 kPa (14.12 psi)

-

100

2080

2570

735

1895

33 °C (91.4 °F) intermediate contingency sea level static

751

798

(10 min)

595

630

97.325 kPa (14.12 psi)

798

2570 -

100.5

2080

-

102

2080

1895

760

28 °C (82 °F) maximum contingency 845

2737

28 °C (82.4 °F)

NOTE: gas generator speed 100 % 2,080 propeller rpm

= 36,660 rpm = 31,023 power turbine rpm

ENGINE POWER RATINGS ACC. TO JAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES WALTER M601E Table 3-1


2019

780




take-off (5 min)

560

595

389

15 °C (59.0 °F)

751

798

0.64

sea level static

560

595

28 °C (82.4 °F)

751

798

max. continuous

490

521

405.9

15 °C (59 °F)

657

699

0.667

sea level static

490

521

21 °C (69.8 °F)

657

699

560

595

751

751

560

595

-

98.1

propeller speed

torque

[rpm]

[N.m] [lb.ft]

2080

2570


690

1895 100

2080

2570

735

1895

-

96.2

97

1700 to 2080 1700 to 2080

2570

660

1895 2570

690

1895

take-off with water injection 300 l/hr (79 US gal/hr) (5 min) sea level static

-

100

2080

2570

735

1895

42 °C (107.6 °F) intermediate contingency sea level static

751

798

(10 min)

595

630

97.325 kPa (14.12 psi)

798

2570 -

100.5

2080

-

102

2080

1895

760

32 °C (90 °F) maximum contingency 845

2737 2019

780

31.5 °C (88.7 °F)



ENGINE POWER RATINGS ACC. TO JAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES WALTER M601E-21 Table 3-2





take-off (5 min)

560

595

395

15 °C (59.0 °F)

751

798

0.6493

sea level static

560

595

23 °C (73.4 °F)

751

798

climb and max. cruise

490

521

410

15 °C (59 °F)

657

699

0.674

sea level static

490

521

18 °C (64.4 °F)

657

699

560

595

751

751

560

595

-

98.6

propeller speed

torque

[rpm]

[N.m] [lb.ft]

2080

2570


710

1895 100

2080

2570

735

1895

-

96.5

97

1700 to 2080 1700 to 2080

2570

680

1895 2570

690

1895

take-off with water injection 300 l/hr (79 US gal/hr) (5 min) 97.325 kPa (14.12 psi)

-

100

2080

2570

735

1895

33 °C (91.4 °F) max. continous sea level static

751

798

(5 min)

595

630

97.325 kPa (14.12 psi)

798

2570 -

100.5

2080

-

102

2080

1895

760

28 °C (82 °F) Max. take-off 845

2737

28 °C (82.4 °F)



ENGINE POWER RATINGS ACC. TO FAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES WALTER M601E Diagram 3-3


2019

780




take-off (5 min)

560

595

389

15 °C (59.0 °F)

751

798

0.64

sea level static

560

595

28 °C (82.4 °F)

751

798

Climb and max. cruise

490

521

405.9

15 °C (59 °F)

657

699

0.667

sea level static

490

521

21 °C (69.8 °F)

657

699

560

595

751

798

560

595

-

98.1

propeller speed

torque

[rpm]

[N.m] [lb.ft]

2080

2570


690

1895 100

2080

2570

735

1895

-

96.2

97

1700 to 2080 1700 to 2080

2570

660

1895 2570

690

1895

take-off with water injection 300 l/hr (79 US gal/hr) (5 min) sea level static

-

100

2080

2570

735

1895

42 °C (107.6 °F) Max. continous sea level static

751

798

(5 min)

595

630

97.325 kPa (14.12 psi)

798

2570 -

100.5

2080

-

102

2080

1895

760

32 °C (90 °F) Max take-off 845

2737 2019

780

31.5 °C (88.7 °F)



ENGINE POWER RATINGS ACC. TO FAR V = 0 km/hr ( 0 kt ), NO INSTALLATION LOSSES WALTER M601E-21 Table 3-4



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. Diagram 3-1



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.

pBL = 14 696/pIN [psi]

INSTALLATION MANUAL MANUAL PART No. 0982502

AIR PRESSURE AT THE AIR BLEED MANIFOLD DELIVERY FLANGE. VARIATION WITH THE AIR BLEED FLOW RATE AND GAS GENERATOR SPEED. Diagram 3-3



REL. FUEL FLOW RATE AND INTERTURBINE TEMPERATURE INCREASE VARIATION WITH AIR BLEED, EQUIVALENT THROTTLING ORIFICE DIAMETER Diagram 3-4



TAKE-OFF RATING SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS. WALTER M601E

Diagram 3-5



TAKE-OFF RATING NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS. WALTER M601E

Diagram 3-6



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

Diagram 3-7



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

Diagram 3-8



TAKE-OFF RATING SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 8 °C CONDITIONS. WALTER M601E

Diagram 3-9



TAKE-OFF RATING NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 8 °C CONDITIONS. WALTER M601E

Diagram 3-10



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

Diagram 3-11



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

Diagram 3-12



TAKE-OFF RATING WITH COOLANT INJECTION SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 20.2 °C CONDITIONS. WALTER M601E

Diagram 3-13



TAKE-OFF RATING WITH COOLANT INJECTION NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 20.2 °C CONDITIONS. WALTER M601E

Diagram 3-14



TAKE-OFF RATING SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA CONDITIONS. WALTER M601E-21

Diagram 3-15



TAKE-OFF RATING NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA CONDITIONS. WALTER M601E-21

Diagram 3-16



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

Diagram 3-17



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

Diagram 3-18



TAKE-OFF RATING SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 13 °C CONDITIONS. WALTER M601E-21

Diagram 3-19



TAKE-OFF RATING NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 13 °C CONDITIONS. WALTER M601E-21

Diagram 3-20



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

Diagram 3-21



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

Diagram 3-22



TAKE-OFF RATING WITH COOLANT INJECTION SHAFT POWER, FUEL CONSUMPTION, GENERATOR SPEED. NO INSTALLATION LOSSES - ISA + 27 °C CONDITIONS. WALTER M601E-21

Diagram 3-23



TAKE-OFF RATING WITH COOLANT INJECTION NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 27 °C CONDITIONS. WALTER M601E-21

Diagram 3-24




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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 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 Mass of oil charge

207 7

kg kg

456 lb 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.



Polar mass moment of inertia about axis x

3.4 kgm2

11,618 lb in2

Mass moment of inertia about axis y

38

kgm2

130,000 lb in2

Mass moment of inertia about axis z

38

kgm2

130,000 lb in2

THE ENGINE MASS MOMENTS OF INERTIA Table 4-2

Coordinate system of the engine (right handed) and positive sense of rotation y + ωy; εy

Axis of the engine Direction of flight + ωz; εz

+ ωx; εx

x

z

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



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 Test bed

Propeller test bed

Sn 3

VEF

VRMS

VEF

VRMS

(mm/sec)

(in/sec)

(mm/sec)

(in/sec)

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.



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))

C = 29,250

[N]

... ultimate load (FAR 33.23 (b)(2))



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.



SCHEME SHOWING LOCATION OF ENGINE MOUNT PADS AND COORDINATE SYSTEM WITH FORCES AND COUPLES ACTING IN POSITIVE DIRECTION Fig. 4-1



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



-2

ny

ny

5

5

4

4

3

3

2

2

1

1

1

-1

2

nx

1

-1

-1

-1

-2

-2

Landing (engine idling) nz = ± 1.5 ωy = 0 ωz = 0 εx = ± 3 εy = ± 10 εz = ± 10

nx

Flight 1

Flight 2

(idling to take-off power) nz = ± 1.5 ωy = ± 0.3 ωz = ± 0.6 εx = ± 4 εy = ± 6 εz = ± 6

(idling to take-off power) nz = ± 1.5 ωy = ± 0.45 ωz = ± 0.6 εx = ± 4 εy = ± 6 εz = ± 6

(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




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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 in2) meets this condition.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 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.



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)



AGRICULTURAL AIRPLANE ENGINE AIR INLET SYSTEM Fig. 5-2 1 - Air Filter 2 - Aerosol Separator (Demister) 3 - Oil Cooler (outside of air inlet system)


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 4 ∆ITT (°C) 2

0

1+

1.00 ∆WF 1+ WF 0.98

∆NH NH 1.00

0.96

0.98

0.96

0.94 0

0.01

0.02

 ∆P     P0  IN

0.03

ENGINE SHAFT POWER, FUEL FLOW RATE AND INTERTURBINE TEMPERATURE AT CONSTANT GAS GENERATOR SPEED. VARIATION WITH AIR INLET SYSTEM PRESSURE LOSS Fig. 5-3 ∆ WF NH ∆ITT ∆P P0 P1

difference fuel flow rate shaft power interturbine temperature increase (°C) difference between P0 - P1 ambient total pressure total pressure at compressor protecting screen inlet



∆p/pSTR 100 (%)

12

∆p/pAV x 100 [%] - ratio of max. acceptable inlet total pressure difference and average inlet total pressure nGR=nG(288.15/T)0:5 nGR - corrected gas generator speed [%] nG - gas generator speed [%] T - outside air temperature [K]

11

10

9

8

7

6

5

4

3

2

1

0 60

70

80

90

nGR (%)

100

INLET AIR DISTORTION LIMIT Diagram 5-4




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 6 FUEL SYSTEM Fuel supply to the engine fuel pump must be ensured by the airframe fuel system under all specified operating conditions and must attain values as follows. The fuel entering LUN 6290.04-8 engine fuel pump must flow through the LUN 7691.01-8 filter. Instruments is not installed on the engine and thus it should be located in the nacelle. By-pass valve of the LUN 7691.01-8 fuel filter starts to open at fuel filter pressure loss of 0.045 to 0.055 MPa (6.5 to 8.0 psi), full fuel flow of 270 l/hr is reached at pressure loss of 0.057 to 0.067 MPa (8.26 to 9.7 psi). Fuel is delivered to the LUN 6290.04-8 engine fuel pump through the flow adapter shown in the engine Installation Drawing, Sheet 3, Sector E4.

Airframe Fuel System Airframe fuel system shall provide engine fuel supply of 370 litre/hr (98 US gal/hr) within all operating conditions. Fuel flow entering fuel filter shall be larger by fuel flow by-passed back to fuel tank. Fuel pressure at the LUN 7691.01-8 fuel filter inlet shall be maintained in the range of 0.15 to 0.3 MPa abs. (21.8 to 43.5 psia). The system shall provide for both easy fuel cut-off upstream of the fuel filter and possibility of regular mud discharge. Fuel entering filter cartridge of the LUN 7691.01-8 fuel filter must not contain free water, air and fuel vapours. The flow adapter on the LUN 7691.01-8 fuel filter fitted with a nozzle of 1.5 mm dia provides for deaerating of the airframe fuel system. Outlet of the pipe from this adapter must be situated below fuel level in the tank to prevent from system aerating in case that the booster pump is off.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Cleanlines of the fuel at the LUN 7691.01-8 fuel filter inlet has to meet the NAS 1638 class 8 requirement as minimum. Total allowed quantity of impuriteis is max. 1 mg/liter. Fuel additives may be used acc. to the principles mentioned in the engine Operation Manual. To prevent from filter cartridge clogging with ice crystals at low temperatures it is possible to fit the fuel system with an oil/fuel heat exchanger upstream of the LUN 7691.01-8 fuel filter.

Drainage In the case of unsuccessful starting, unburnt fuel mostly accumulates in the lowest part of the combustion chamber jacket; some fuel passes then further through turbines and accumulates in the lowest part of the outlet channel from where it flows out through a hole in the outlet channel wall into the outlet casing. Certain amount of unburnt fuel comes into outlet casing directly via the push fit collar of the outlet channel in the power turbine nozzle guide vane ring. From the combustion chamber jacket bottom, the accumulated fuel is brought via the pipe union directly to the drain valve and, from there, to the common drainage outlet. Fuel accumulated in the outlet casing is directed via a pipe to the drain valve body where it joints the fuel from the combustion chamber jacket and flows through the above pipe to the common drainage outlet. Then the fuel should be handled by the airframe fuel system. Max. 0.9 litre/hr (0.24 US gal/hr) of fuel leaks during normal engine operation from drainage tube of the fuel pump and FCU (Installation Drawing, Sheet 3, Sector C4) and in addition to it fuel quantity of 0.2 litre (0.05 US gal) bleeds after each shutdown. This fuel is not redelivered to the engine and must be removed by airframe drainage system.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 7 ELECTRICAL SYSTEM AND MONITORING INSTRUMENTS The power unit electrical system is supplied with nominal voltage of 27 V with the minus pole connected to the frame. As a power source a starter/generator complimented by a storage battery located in the airplane is used. The electric system including the monitoring instruments provides for: −

engine starting;

−

storage battery charging;

−

function of the engine monitoring instruments;

−

limiting at start and at the reverse;

−

propeller de-icing (if the LUN 2102.01 additional alternator, additional gearbox and the LUN 7850-7 brush block assembly are used). This AC source can be also utilized for airframe needs.

The electric system of the engine and that of the aircraft are interconnected by four connectors illustrated in the engine Installation Drawing, Sheet 1, Sector 2-5, F-G and Sheet 3, Sector 4-5, G-H and Sector 7E. Wiring of individual connectors is shown in the enclosed wiring diagrams - Figures 7-1 to 7-3. Block diagrams are presented in Figures 7-4 to 7-6. Arrangement of engine instruments is shown in Fig. 12-1. Connector counterparts of the wiring harness and of instruments which are necessary for electric installation of the aircraft are supplied on a special order only. Connector counterparts of the instruments included in the basic engine equipment and installed in the airframe are delivered together with the instruments as an accessory.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 DESCRIPTION OF THE INSTRUMENTS LUN 1333.12-8 Integrated Gas Generator and Propeller Speed Transmitters The integrated gas generator and propeller speed transmitters provide for remote transmitting of the engine speeds. The integrated gas generator speed transmitter is situated in the engine rear part on the front face of the accessory gearbox and transmits gas generator speed. The integrated propeller speed transmitter is situated in the engine front part on the reduction gearbox and transmits propeller speed. The integrated speed transmitter consists of two parts: three-phase alternator and pulse generator. The three-phase alternator is the main part of the instrument, pulse generator is located in the rear part of the alternator. The three-phase alternator supplies alternating current for the speed indicator, pulse generator transmits signal for the limiter system.

LUN 1377-8 Interturbine Temperature Transmitter A set of the LUN 1377-8 transmitters together with the LUN 1370.02-8 ITT indicator provide for the remote measurement of the interturbine temperature. Temperature transmitters are attached by two bolts to the exhaust casing. Nine transmitters are used as a set for one engine. They are parallely interconnected by busbars terminated in soldering eyes. A compensation line transmitting signals for the indicator and for the LUN 5260.04 integrated electronic limiter unit is connected to these soldering eyes. Resistance of the compensating line connecting the indicator should be 8 ± 0.1 ohm at ambient temperature of 20 oC (68 oF). Resistance of the compensating line connecting the limiter system should be 6.2 ± 1 ohm at ambient temperature of 20 oC (68 oF). As far as the principle of operation is concerned the transmitters operate as chromelalumel thermocouples. A 4 mm (0.16 in) dia. hole is provided on the soldering eye for chromel terminal and a 5 mm (0.20 in) dia. hole for alumel terminal in order to prevent incorrect interconnection of terminals. A stream of gas passes by the thermocouple shield and warms the thermocouple measuring tip. Due to this a thermoelectric power is originated which is proportional to the gas temperature. Originated voltage is transferred to the magnetoelectric system of the indicator and to the limiter system. Operational range of transmitter's temperature is from 0 to 900 oC (32 to 1,652 oF); for short-time operation of up to 1,200 oC (2,192 oF).


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Torquemeter Indicator Set A torque indicator set consists of the LUN 1540.02-8 transmitter and the LUN 1539.028 indicator. It provides for engine torque measurements within the range of torque from 0 to 120 %. The set is supplied, together with the engine, and is already calibrated. Pressure transmitter of the torquemeter is located on a bracket in the engine's air intake compartment close to the engine's upper mount. Pressure oil from the reduction gearbox is fed (by a pipe) to the space behind the front bulkhead to the torquemeter pressure transmitter. In the inlet screw union of the torquemeter pressure transmitter a choking coil is situated for smoothing the pulses of the pressure. The transmitter operates as an inductive converter of torquemeter oil pressure into the electric signal. The electric signal of the torquemeter pressure transmitter is evaluated in the indicator. The torquemeter pressure indicator is situated on the control board in the cockpit. It operates as a servomechanical autocompensating bridge. It is designed as an induction sensing unit identical to that one in the transmitter and both sensing units are bridge connected. The voltage in the coils is induced by the current passing through the exciter of the sensing unit winding. The difference between the electric signals from the induction sensing units is amplified by an amplifier and led to the servomotor which turns the armature of the induction sensing unit indicator up to the difference voltage compensation. The induction sensing unit of the indicator is coupled by gears with the instrument pointer which indicates the measured value on the scale calibrated in percentage of torque. The bridge wiring of sensing units is diagonally by-passed by a pair of shunts the value of which can be changed by means of the switch for rough adjustment of the indicator range for 100 % of torque. Fine adjustment is to be carried out mechanically. The set of the torque indicator can be adjusted within the range of torquemeter oil pressure value from 0.93 MPa (135 psi) to 1.03 MPa (150 psi). Each instrument of the set, i.e. transmitter or indicator can be replaced in case of failure. When replacing one, a new adjustment is to be carried out for both instruments according to the instructions in its certificate. A power source of 36 V, A.C. and 400 Hz is necessary for torque indicator set supply. The aircraft should be equipped with this source, as this is not included in the engine installation. Resistance of the electric line between the transmitter and indicator can be max. 0.62 ohm. Alternativelly (on special order) the engine can be fitted with another approved set. The set can be adjusted together with the engine on the engine manufacturer′s test bed or after engine installation in the airframe at customer′s facility acc. to approved documentation.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 LUN 1358-8 Oil Temperature Resistance Transmitter Electric resistance transmitter of oil temperature together with the LUN 1538.01-8 three pointer indicator form a set intended for remote measurement of oil temperature in the engine. The oil temperature resistance transmitter is situated in the engine's rear part on the accessory gearbox case. Its sensor is immersed in the oil tank compartment i.e. it measures the oil temperature in the tank. The temperature transmitter's function is based on the physical property of the thermometric sensor which changes its resistance in accordance with the temperature change. The changes in electric resistance are transferred by a connecting wire to the LUN 1538.01-8 three pointer indicator, which indicates the temperature in centigrades. The three pointer indicator is located on the panel in the cockpit and is included in the airframe installation. Variation of Transmitter Resistance With Temperature from -70 oC to 150 oC (-94 oF to 302 oF) Temperature

Transmitte r resistance

(°C)

(°F)

(Ω)

-70

(-94)

-60

Temperature


Temperature


(°C)

(°F)

(Ω)

(°C)

(°F)

(Ω)

68.36

10

(50)

93.76

90

(194)

125.56

(-76)

71.06

20

(68)

97.36

100

(212)

129.96

-50

(-58)

73.86

30

(86)

101.06

110

(230)

136.41

-40

(-40)

76.86

40

(104)

104.86

120

(248)

138.96

-30

(-22)

79.96

50

(122)

108.81

130

(266)

143.56

-20

( -4)

83.16

60

(140)

112.78

140

(284)

148.36

-10

(+14)

86.56

70

(158)

116.96

150

(302)

153.26

0

(+32)

90.26

80

(176)

121.22

Resistance of the connection wiring between the transmitter and the indicator of 0.16 ohm has been included. Current load must not exceed the value of 10 mA.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 1.25 K LUN 1469.32-8 Min. Oil Pressure Switch The pressure switch is situated on the upper part of the accessory gearbox. This switch provides for signalling the engine's oil pressure drop. A signaller situated on the panel in the cockpit is connected in the pressure switch circuit and becomes lit when the oil pressure drops below the value of 0.125 MPa (18 psi). Following the pressure increase the signaller is switched off again.

LUN 1559-8 Fuel Pressure Transmitter NOTE: Measurement/indication of the fuel pressure upstream of the fuel distributor is not required by FAR. If the fuel pressure is not measured it is necessary to blind screw union on the tube delivering fuel to the engine with aid of approved cap. The engine is delivered with shipping cap only. The fuel pressure transmitter shall be situated in the engine nacelle and senses the fuel pressure upstream of the fuel distributor. It operates in a set with the LUN 1538.018 three pointer indicator which is situated on the panel in the cockpit. Fuel pressure for the transmitter is fed from the pipe union which is located on the pipe delivering fuel from the FCU to the engine in the L.H. lower part (Installation Drawing, Sheet 2, Sector D11). A hose connecting the transmitter is an airframe installation part. The fuel pressure transmitter operation results from the diaphragm deflection due to acting pressure. Thus the electrical characteristics of the exciting coils are changed. Due to diaphragm deflection air gap magnitude changes and thus permeance of magnetic circuits changes. Change in the permeance causes change in magnetic flux, which is closed through the air gaps by common armature in the individual circuits. Of course, this influences the value of inductance and impedance of the exciting coils. Currents passing through the coils turn by the transmitter armature. The position of the pointer in the three pointer indicator is found so that the transmitter/receiver armatures angular displacements will correspond with each other. The measuring range of the transmitter is from 0 to 1.6 MPa (232 psi). The operational feeding voltage is within the range of 32.5 to 38 V A.C., 400 Hz ± 2 % and the distortion is up to 20 %. The maximum current for feeding by operational voltage is 80 mA. Each branch resistance of the line between the transmitter and the indicator must not exceed 0.6 ohm.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Equivalent Resistance of the Pressure Transmitter Including the Inductive Transducer Fuel Pressure Value MPa (psi) 0

Resistance Between Transmitter Pins No. 2 and 1 (ohm)


(0)

110.0

150.0

0.2

(29)

115.8

144.2

0.4

(58)

121.0

139.0

0.6

(87)

125.6

134.4

0.8

(116)

130.0

130.0

1.0

(145)

134.4

125.6

1.2

(174)

138.8

121.2

1.4

(203)

143.8

116.2

1.6

(232)

150.0

110.0

LUN 1558-8 Oil Pressure Transmitter The oil pressure transmitter shall be located in the engine nacelle and provides for oil pressure sensing in the engine. It operates as a part of a set with the LUN 1538.01-8 three pointer indicator which is located on the panel in the cockpit. Oil pressure is sensed by this transmitter in the space of the flow adapter on the oil supply pipe at the pressure switch intake of the accessory gearbox's rear wall (Installation Drawing, Sheet 1, Sector F5). Pressure is fed to the transmitter by a hose which is an airframe installation part. The oil pressure transmitter operates on the same principle as the LUN 1559-8 fuel pressure transmitter. The measuring range is from 0 to 0.4 MPa (58 psi), the operational range is from 0.06 to 0.34 MPa (8.7 to 49.3 psi). The operational alternating feeding voltage is within the range of 32.5 to 38 V, 400 Hz ± 2 %; distortion is up to 20 %. The maximum feeding current is 80 mA at operational voltage. Each branch resistance between the transmitter and the indicator must not exceed 0.6 ohm.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Equivalent Resistances of the Pressure Transmitter Including the Inductive Transducer Oil Pressure MPa

(psi)



0

(0)

110.0

150.0

0.04

(5.8)

114.7

145.3

0.06

(8.7)

117.0

143.0

0.10

(14.5)

121.1

138.9

0.16

(23.2)

123.6

133.4

0.20

(29)

130.0

130.0

0.26

(37.7)

135.2

124.8

0.30

(43.5)

138.9

121.1

0.34

(49.3)

142.8

117.2

0.36

(52.2)

145.0

115.0

0.40

(58)

150.0

110.0

Function of Instruments During Engine Starting By depressing the „START“ push-button for a short time, the LUN 2601.01-8 time relay is switched for the period of 20 sec and provides for: − the power switch of the LUN 2132.02-8 starter/generator is switched on; − the current is fed to the ignition system; − the current is fed to the LUN 3191-8 interrupter, which provides for pulse operation of the fuel inlet valve and so regular periodic breaks of two to three sec. in the fuel supply to the torch igniters are ensured. By depressing the „MOTORING RUN“ push-button for a short time, the LUN 2601.01-8 time relay is put in operation, this relay provides for switching on of the LUN 2132.02-8 starter/generator power switch only. Engine starting is further influenced by the starting fuel control unit, which provides for continuos growth of the fuel suuply in the combustion chamber in accordance with the compressor delivery pressure.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 An integrated electronic limiter unit IELU - LUN 5260.04 (which is described separately) is used as fuel supply limiter at interturbine overtemperature during engine starting. Course of starting is fully automated and the pilot′s activity during engine starting is reduced to depressing the push-button „START“ only. The switch for spark plug check provides for separate check of individual ignition units. If the switch is in one of the extreme positions, the voltage is fed to the respective ignition unit. No voltage is fed (conducted) to the ignition units when the switch is in the middle position. Block diagram is in the Fig. 7-5.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 LIMITER SYSTEM - DESCRIPTION 1.

General The limiter system is be fitted with LUN 5260.04 Integrated Electronic Limiter Unit (IELU). The LUN 5260.04 system operates as a two-level limiter of - gas generator speed - torque - interturbine temperature - propeller speed The mentioned parameters are limited above the level of permanently permitted values but in the way still preventing engine deterioration. The limitation of parameters is a matter of importance for the multi-engined aircraft. There is a possibility of coupling with propeller automatic feathering. In case of emergency with one engine out it is necessary to cut-off the limiter of the operating engine and select the maximum contingency rating. The operation of the engine at this power rating is permitted for 2.5 minutes - at parametres corresponding to gas generator speed of 102 %. The manufacturer provides on request the set of necessary data for the whole limitation system design. The limiter system provides for: A. Signalling of monitored parameter limit value exceeding. The following parameters are monitored: gas generator speed

nG

propeller speed

nV

interturbine temperature

ITT

torque

Mk (TQ)

ITT rate of growth at engine start

dITT/dt

The limiter provides for visual signalling if the control current exceeds value of 3 mA which corresponds to exceeding monitored parameter over the adjusted value of nG, nV and ITT by app. 0.15 %.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 B. Protection of the engine against overloading. Limiter system decreases engine fuel delivery when limit values of the monitored parametres have been exceeded. In addition to signals which identify specific engine operation (e.g. engine starting, engine operation in BETA range (reverse)) or inform about undercarriage retracting, ambient temperature and flight altitude signals from the following transmitters are transmitted to the limiter: - gas generator speed transmitter

nG

- propeller speed transmitter

nV

- interturbine temperature transmitter

ITT

and from the torque limiter pressure switch

Mk (TQ)

The limiter evaluates signals of the nG, nV and the ITT and compares them with the values adjusted in the limiter. In case of higher value of the input signal than that adjusted in the limiter, the limiter emits control current. Its value depends on exceeding magnitude and number of exceeded signals. Control current provides for required function of the limiter system. The limiter evaluates also (in addition to the true speed) corrected gas generator speed. Necessary information on the ambient temperature for calculation of corrected speed is provided by the P-5(7) ambient temperature transmitter. Signal from pressure switch of torque limiter which results from exceeding adjusted torque value in the switch, initiates transmitting max value of control current from the limiter. The limiter evaluates also the ITT rate of growth dITT/dt at starting and in case of higher rate than that adjusted in the limiter, the limiter transmits max. control current.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 2.

Description Limiter system fitted with the LUN 5260.04 integrated electronic limiter unit consists of the below listed equipment and parts. For block diagram of the limiter system see Fig. 7-6. WALTER a.s. delivers wiring diagram of the system upon request. A. LUN 1333.12-8 integrated speed transmitter (gas generator speed) B. LUN 1333.12-8 integrated speed transmitter (propeller speed) C. LUN 1377-8 interturbine temperature transmitter D. LUN 1476-8 torque limiter pressure switch E. A-037 Radioaltimeter F. Aircraft undercarriage switch G. LUN 5223 Generator speed derivative element H. LUN 5260.04 integrated electronic limiter unit (IELU) I. Electrohydraulic Transducer on the Fuel Control Unit J. LUN 3280-8 Pressure switch for propeller automatic feathering K. LUN 7816-8 Propeller speed governor L. P-5(7) Ambient Temperature Transmitter M. Signallers, interconnecting lines, controls N. LUN 2601.01-8 time relay of the starting panel


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Description of the individual system components A., B. LUN 1333.12-8 Integrated Speed Transmitter Description and function are presented separately, refer to Page 7-2.

C. LUN 1377-8 ITT Transmitter Description and function are presented separately, refer to Page 7-2.

D. LUN 1476-8 Torque Limiter Pressure Switch The switch closes at adjusted level of torquemeter oil pressure, which is proportional to torque. Adjustment is carried out on the test rig by means of a control element which is situated on the switch under cover. A damper which absorbs oil pressure oscillations inside the switch is located in the inlet flow adapter of the switch. The switch is attached to a bracket on the engine below mount ring in 1 o′clock position.

E. A-037 Radioaltimeter Only that signal is used from the radioaltimeter which is obtained as a sum of a signal proportional to the altitude above ground and a signal of the radioaltimeter pilot lamp. The signal is derived so that the signal for the radioaltimeter indicator will be not affected and the permitted loading of the radioaltimeter as shown in its technical specifications will be not exceeded.

F.

Aircraft undercarriage switch This is a limit switch which provides for grounding the pin (17) on LUN 5260.04 if the undercarriage is retracted. It is mounted on the aircraft. Its wiring is shown in the diagram B 071175 in the documentation of L410UVP-E airplane.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 G. The LUN 5223 generator speed derivative element The LUN 5223 derivative element is an electronic fully transistorised device. It serves for stabilization of engine running (for reduction of generator speed fluctuations) in the case of engine power limitation by torque limiter or in the case of torque maintaining at 65 to 75 % with the first limiting level on. The derivative element consists of the converter of frequency derived from the generator speed to voltage which is applied to the derivative circuit input. The derivative circuit processes the voltage change so that, starting with a certain rate of voltage drop, voltage on the derivative block output will drop from 10 V to 0 V (supply voltage for LUN 1476-8). The derivative block includes also a time circuit consisting of a mono-stable circuit which cancels the function of the derivative circuit if more than 6 ± 1 seconds passed following the occurrence of control current in the system of limiters (system intervention or failure state). The derivative block includes a light-emitting diode which serves for checking proper functioning at periodic inspections. The derivative block is mounted in the airframe close to LUN 5260.04.

H. LUN 5260.04 Integrated Electronic Limiter Unit (IELU) The integrated electronic limiter unit is an electronic instrument which is inserted in the spring mounted frame in the airframe. Location of this frame depends on aircraft type and is presented in the aircraft documentation. Wiring except compensation line of the thermocouples is joint to the spring mounted frame and through socket/plug connectors is connected to the limiter after its insertion to the frame. Compensation line of the thermocouples is connected by means of eyes to the limiter.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 LUN 5260.04 is an electronic fully transistorised device. It processes signals from individual transmitters of parameters subjected to limitation (gas generator speed, propeller speed, inter-turbine temperature ITT, torque). At starting and at activating the push-button „CHECK“, ITT temperature limit is reduced and additionally the limited parameter dITT/dt is to be respected. When any of the preset values of these parameters is exceeded, the unit transmits an electric signal to the actuator of the system of limiters on the FCU which provides for the reduction of the supply of fuel to the engine, and thus for the reduction of the parameter subjected to limitation. The limited value of generator speed depends on the ambient temperature; at temperatures below 0 oC, generator speed is limited according to the formula: nGred = (corrected speed) = const. Depending on the radioaltimeter signal, on the position of the undercarriage switch, on the position of the switch on LUN 7816-8, on the state of switch LUN 3280-8 and push-button „CHECK“, engine rating can be limited by the IELU as follows: 1. at the first limiting level, to a rating corresponding to the torque of 65 to 75 %, 2. at the second limiting level, to a rating corresponding to the supply of fuel for generator speed of approximately 60 %. The device has no actuating elements. It enables successive checking of regulation channels of the generator speed limiter, propeller speed limiter and torque limiter by means of a switch and push-button „TEST“ located on the front panel of the IELU. The channel of the ITT limiter can be checked by depressing push-button „CHECK“. The checking of individual channels is performed without removing the device from the aircraft, with the engine running. The device provides for optical signalling of: a) switching-on the device, b) switching-on the first limiting level, c) intervention of any limiter.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Limiting level II is switched on automatically: a) the altitude of flight is in range from H = 700 m SOL to 3,700 m ISA with the undercarriage retracted - signal is provided by the switch on the aircraft undercarriage and A-037 altimeter; b) in the BETA control range and at reverse - signal is passed through the switch on the LUN 7816-8 propeller speed governor; c) the engine is starting; ITT (subject to limitation) is limited to a lower value and further parameter dITT/dt subjected to limitation is being switched on. The function of other limiters remains unchanged. Lower limit for ITT and dITT/dt are engaged only during the operation of the starting panel and if the corresponding pair of contacts in the LUN 3280-8 switch is closed; d) with depressed „CHECK“ push-button when ITT subjected to limitation is switched to a lower value and another parameter dITT/dt subjected to limitation is switched on. There is no change in the function of the other limiters.

I.

Electrohydraulic Transducer on the Fuel Control Unit The electrohydraulic transducer modifies the pressure in the servosystem of the main metering needle in accordance to the control current from the limiter system. Pressure change in the servosystem results in position change of the main metering needle and thus change in fuel supply. The transducer is neither adjusted in operation nor separately checked. Check is carried out within limiter system check.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 J.

LUN 3280-8 Pressure Switch for Propeller Automatic Feathering The pressure switch for automatic propeller feathering is located on a bracket in the space of engine intake in the same plane as the torque indication set transmitter and torque limiter pressure switch. The pressure switch has two levels of contact closing. At the higher pressure level it closes the circuits for automatic trimming of airplane rolling in the electric installation of the airplane and blocks the function of automatic feathering of the second engine. At the lower pressure level it closes circuit of automatic feathering, by an independent pair of contacts switches the limiter over to the lower level of ITT limit when starting-up the engine and simultaneously switches on a circuit limiting the rate of temperature growth.

K. LUN 7816-8 Propeller Speed Governor Signal is derived from B 613-3A mechanical microswitch which is mounted under the governor shield. The switch, depending on the position of the propeller feedback ring closes, at the moment when the propeller blade angle starts to be set to value below the minimum flight angle to small positive angles. The closing of the switch activates the system of limiters and engages the second limiting level. This is optically signalled on the panel in the cockpit by lamp „BETA“. The propeller speed governor is mounted on the engine.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 L.

P-5(7) Ambient Temperature Transmitter Ambient temperature transmitter is based on the principle of thermal variation of the electric resistance. Transmitter is installed under engine cowling. The transmitter can be replaced by another approved resistance transmitter with the following properties: Ambient air temperature [°C]

Transmitter resistance [Ohm]

-60

70.9

-50

73.7

-40

76.7

-30

79.8

-20

83.0

-10

86.4

0

90.1

+10

93.6

> +10

> 93.6

The transmitter must be situated in the engine nacelle so to sense true ambient air temperature. The transmitter is not adjusted in operation. NOTE: An independent transmitter shall be installed for each engine (multiengine airplane).


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 M. Signallers, interconnecting lines, controls These items are included in the airframe installation of the limiter system. Wiring diagram should be included in the airframe documentation. The following signallers should be situated within pilot's viewing angle: 1.

„PARAMETER EXCEEDING“ signaller indicates exceeding of some monitored parameter.

2.

„ELECTROHYDRAULIC TRANSDUCER“ („EHT“, „ALERT IELU“, or „IELU OPERATIVE“) signaller indicates that system reduces fuel supply when some monitored parameter exceeds limit.

3.

„BETA“ signaller indicates that propeller blades are adjusted below min. flight angle.

N. LUN 2601.01-8 Time Relay of the Starting Panel The time relay is a part of the starting panel. The relay provides the limiter system with necessary signals for 20 sec, it informs about starting. The relay is installed in the airframe.


27 V

water injection

feathering system

MANUAL AND AUTOMATIC SYSTEMS OF PROPELLER FEATHERING AND EMERGENCY CIRCUIT SWITCH

LUN 3191-8 Interrupter

8

11 12 19 13

4

8

2

3

5

5

5

7

9

V8 2RM 33 KPN 20 Š4V1 - 2 RM 33 BPN 20 G4V1 CONNECTOR WIRING Fig. 7-1

2

3

V 110-2

2 RM 14 KPN 4G

VŠ 18 KPN 5G 1U2

Section 7 Page 7-19/7-20 Jun 26, 2001

1 2 3 4

A B V G D

SWITCH ON THE ECL ON THE ENGINE

TORCH IGNITER FUEL VALVE ON THE LUN 6290.04-8

WK 180 36

1 2 3 4 5 6 7

CHIP SIGNALLER ACCESSORY GEARBOX

5

V 112-2

WK 180 36

1 2 3 4 5 6 7

CHIP SIGNALLER REDUCTION GEARBOX

V 110-2

VŠ 17 KPN 3G 1U2

A B V

N25F-3 (CHAMPION) SPARK PLUG

V 110-2

V 110-1

VŠ 17 KPN 2G1

VŠ 17 KPN 2G1

A B

A B

UNISON

EMERGENCY CIRCUIT VALVE ON THE LUN 6590.05-8

EXCITER

M 304

B4

E 96

E 94

M 28

B8

B2

N25F-3 (CHAMPION) SPARK PLUG

WALTER a.s.

4

STARTIN G PANEL


11 12 19 13

STARTIN G PANEL

27 V LUN 1348-8

LUN 1347-8

Propeller Speed Indicator

Gas Generator Speed Indicator

1 2 3 4 5

27 V 27 V

LUN 1539.02-8 Torque Indicator

7 2 1 4 5 6 3

1 2 3 4 5

lighting

1 2 3 4 5 6 7

red

5V

VŠ 17 KPN 3G1 - 2x

WK 180 36

35 V 400

„BETA“

WK 180 36

lighting

manual and autofeathering system

5V limiter system

limiter system

autofeathering system

limiter system

limiter system limiter system recorder

8

9

10 19 18 32

6 7 11 12 13

4 3 5

8

9

10 19 18 32

27 31 26 28 23

21 20

22

2

17

15 16

24 25

30

14

30

14

2 RM 30 KPN 32 Š1V1 - 2 RM 30 BPN 32 G1V1

V 16

27 31 26 28 23

21 20

22

2

17

15

24 25

16

V8 15

V112-2

1

V 110-1

VŠ 17 KPN 3G 1U2 VŠ 18 KPN 5G 1U2 - 2x

A B V G D

Section 7 Page 7-21/7-22 Jun 26, 2001

LUN 1333.12 Integrated Speed Transmitter (Propeller)

M 36

A B V G D

LUN 1333.12 Integrated Speed Transmitter (Gas Generator)

M 34

WK 180 36

WK 180 36

WK 180 36

1 2 3 4 5 6 7

1 2 3 4 5 6 7

1 2 3 4 5 6 7

LUN 3280-8 Pressure Switch of the Autofeathering System

M 310 Optional

LUN 1540.02-8 Transmitter for Torquemeter Indicator

M 10

LUN 1476-8 Torque Limiter Pressure Switch

M 172

VŠT 18 KPN 5G1U2 VŠ 17 KPN 2G1

A B V

V 112-2

A B

V 110-3

774-S40D4

2RM14 KPN 4G1V1

A B V G D

1 2 3 4

LUN 7880.01-8 Electrohydraulic Actuator

Electrohydraulic Transducer on the LUN 6590.05-8

1.25K LUN 1469.32-8 Min Oil Pressure Switch

Switch on the LUN 7816-8

M 312

M 174

M 14

M 176

A B

LUN 1581-8 Min Oil Quantity Signaller

M 18

WALTER a.s.

4 3 5


CONNECTOR WIRING Fig. 7-2

6 7 11 12 13


Mains 27 V

F = 35 mm2

Starting Push Button

Switch for Spark Plug Check

Motoring Run Push Button LUN 2167.03-8 Regulator

Torch Ignitor Fuel Valve on the LUN 6290.04-8

to the Regulator of the 2nd Engine

F = 1.5 mm2


to Storage Battery

+

F = 35 mm2 LUN 2201.03-8 (2 pcs) or UNISON 9049765-1 Ignition Set

Spark Plug

LUN 2601.01-8 Time Relay of the Starting Panel

Starter/Generato r

Spark Plug

BLOCK DIAGRAM - STARTING SYSTEM Fig. 7-5



lighting

5V

1

ITT Indicator

2 774.24-8

CH

+

-

A

LUN 5260.04-8

+ -

CH

-

A

+

IELU

CH A

CH

-

A

+

LUN 1377-8 Interturbine Temperature Transmitter (9 pcs)

WIRING DIAGRAM - ITT MEASUREMENT Fig. 7-4



Mains 27 V

F = 35 mm

Starting Push Button

Switch for Spark Plug Check

2

Motoring Run Push Button LUN 2167.03-8 Regulator

Torch Ignitor Fuel Valve on the LUN 6290.04-8

to the Regulator nd of the 2 Engine

F = 1.5 mm


2

to Storage Battery

+

F = 35 mm LUN 2201.03-8 (2 pcs) or UNISON 9049765-1 Ignition Set

N 25F-3 or CHAMPION Spark Plug

LUN 2601.01-8 Time Relay of the Starting Panel

2

LUN 2132.02-8 Starter/Generato r

N 25F-3 or CHAMPION Spark Plug

BLOCK DIAGRAM - STARTING SYSTEM Fig. 7-5



LUN 1333.12-8 Integrated Speed Transmitter (Gas Generator) LUN 1333.12-8 Integrated Speed Transmitter (Propeller) LUN 1377-8 ITT Transmitter

LUN 7816-8 Propeller Speed Governor

LUN 5260.04-8

Integrated electronic limiter unit (IELU)

Switch on the Undercarriage

LUN 2601.01-8 Time Relay

Cockpit Push Button IELU

LUN 6590.05-8 Fuel Control Unit

Switch in the System of Automatic Feathering of the Switch of Feeding Second Engine

A-037 Radioaltimeter

TEST

P-5(7) Temperature Transmitter

SWITCH PUSH BUTTON

Pilot Lamps „IELU“

LUN 1476-8 Pressure Switch of the Torque Limiter LUN 5223 Generator Speed Derivative Element

BLOCK DIAGRAM OF THE SYSTEM OF LIMITERS Fig. 7-6


LUN 3280-8 Pressure Switch for Automatic Propeller Feathering

„PARAMETER EXCEEDING“

„ALERT IELU“

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 8 LUBRICATION SYSTEM The engine is equipped with an independent pressure and scavenge circulation oil system including a single oil tank inside the engine. The lubrication system provides for lubrication and cooling of all bearings and gears in the engine. By means of additional equipment the pressure oil is also used for torque measurement in the reduction gearbox and for propeller speed control. The following appliances are connected to the engine in the aircraft: - Oil cooler - Oil/fuel heat exchanger (optional) - Electric feathering pump (optional) Due to great demands it is necessary to use only the specified grades of oil. Usage of such oils was approved and it is necessary to adhere to the conditions stated in the „Operation Manual“. The oil tank is an integral part of the accessory gearbox case in the engine's rear part. There are 7 litres of oil (1.85 US gallons) in the oil tank but in order to fill the whole oil system including the cooler and the oil/fuel heat exchanger it is necessary to fill a further oil quantity of about 4 litres (1 US gallon), which is necessary to refill following the first engine motoring run. Circulation of oil inside of the engine is effected by a single pressure pump and three scavenge pumps located in the accessory gearbox. In addition to it there is a dual pump in the reduction gearbox; one stage of this provides for oil repumping from the power turbine to the sump in the reduction gearbox at greater pitch angles of the aircraft; the second one is a pressure stage which provides for torque equalisation on the reduction gearbox countershafts and, at the same time for torque measurement. Oil is filtered by the main oil filter (see Installation Drawing, Sheet 5, View R3) with an eye size metal filter cloth of 31.5 µm (1.24 . 10-6 in).


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Protection strainers are arranged at the scavenge pump's intake. The strainers can be checked from the outside of the engine. One strainer is arranged on the reduction gearbox (see Installation Drawing, Sheet 2, Sector CD - 3) and two strainers are situated on the accessory gearbox case. In the oil branch from the generator turbine the strainer is located on the lower part on the accessory gearbox periphery (see Installation Drawing, Sheet 3, Sector C6). In the oil branch from the gears of the accessory gearbox the strainer is located on the lower part of the accessory gearbox's front wall. This strainer is accessible after the chip signaller's removal (see Installation Drawing, Sheet 3, Sector D5). The protecting strainer of the oil enter to the reduction gearbox and the propeller turbine is accessible from the front side of the reduction gearbox (see Installation Drawing, Sheet 3, Sector D11). Filling and checking of oil charge should be enabled by sufficiently large ports in the engine cowling without the necessity of further parts removal. The oil filter, local protecting strainers, and chip signallers must be accessible when carryingout the specified engine maintenance procedures. For oil cooling the lubrication system is fitted with an oil cooler. The cooler is attached to flow adapter installed in the lower part of the accessory gearbox (shown in the Installation Drawing, Sheet 3, Rear View, Sector CD-5) by means of hoses of minimum inner dia. 16 mm (0.62 in).

Leading particulars of the cooler: -

Minimum cooling performance at flow rate of 30 kg/min (66.15 lb/min) at oil inlet temperature of 110 oC (230 oF)

9,600 W (546 Btu/min)

-

Operation pressure

max. 0.25 MPa (36psi)

-

Pressure peak at cold engine starting

max. 0.4 MPa (58 psi)

-

Hydraulic drag of the cooler at flow rate of 30 kg/min (66.15 lb/min) and oil temperature of 90 oC (194 oF)

0.05 ± 0,02 MPa (7.2 ± 2.8 psi)

-

Oil inlet temperature

-

Thermostatic valve ensuring the oil outlet temperature control within the range of 60o to 70o C (140o to 158o F) must be arranged in the cooler; maximum temperature is 85 oC (185 oF).

-

As a part of the cooler a relief valve must be installed which provides for by-pass opening at the pressure of 0.35 MPa (51 psi) within the whole range of the operation temperatures.


max. 130 °C (266 °F)

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 The oil tank filler neck is situated in the common compartment together with the oil dipstick and with the oil filter cap on the upper part of the accessory gearbox case. In order to ensure correct operation of the lubrication system the engine is equipped with the following instruments: - LUN 1358-8 oil temperature resistance transmitter - LUN 1558-8 oil pressure transmitter - 1.25 K LUN 1469.32-8 min. oil pressure switch - LUN 1581-8 signaller of the min. oil quantity in the tank For check of steel parts condition, which can be subjected to the lubrication oil influence, two chip signallers are installed in the engine; one of them is situated in the rear part of the accessory gearbox and the second one is in the oil sump in the reduction gearbox's lower part. The location of these signallers is shown in the Installation Drawing, Sheet 3, Sector D5, and in Sheet 2, Sector C4. A magnetic drainage plug is situated on the bottom of the oil tank. The wiring diagram of electric signallers and transmitters with indicators is shown in Section 7 of this „Installation Manual“. On special order the following instruments can be delivered: - LUN 7840-8 electric feathering pump The following appliances are parts of the L410UVP-E airplane installation: - Fuel/oil heat exchanger - oil cooler These instruments are not included in the engine, but in the airframe installation.

CAUTION:

IF THE OIL COOLER IS NOT ARRANGED IN THE LUBRICATION SYSTEM AND THE FLOW ADAPTERS, WHICH CONNECT THE COOLER TO THE ACCESSORY GEARBOX (SEE INSTALLATION DRAWING, SHEET 3), ARE CLOSED BY A PLUG, IT IS FORBIDDEN TURN THE GAS GENERATOR ROTOR BECAUSE THE OIL PUMPS AND THEIR DRIVES CAN BE DAMAGED.




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 9 COOLING REQUIREMENTS The engine nacelle shall be designed so that the surface of the engine and of its accessories will be cooled sufficiently. The engine and its parts can be cooled by both external and internal flow of cooling air. The conditions inside the nacelle required for external cooling of the engine surface shall be such as to prevent the engine and accessories' surface temperature distribution from exceeding maximum permitted values at all power ratings in both flight and ground operation and at the following engine shutdown. Therefore the nacelle is provided with openings for the necessary cooling air flow, especially between the engine's rear bulkhead and the firewall of the airplane. The efficiency of the cooling system and the inner nacelle layouts effect on the engine surface temperature distribution shall be checked to ascertain that maximum acceptable temperatures are not exceeded at specified points at all power ratings in both flight and ground operation and at any moment following engine shutdown. The specified points are located: - air temperature above hooking eye of reduction gearbox 25 mm (1inch) Maximum temperature: 100 oC (212 oF). - under fixing nut of the flange of the LUN 6590.05-8 fuel control unit intake flow adapter (designated „A“). See the Installation Drawing, Sheet 3, Sector D4. Maximum temperature: 80 oC (176 oF). - on the LUN 6290.04-8 fuel pump intake flow adapter. Point „A“ in its cylindrical part. See the Installation Drawing, Sheet 3, Sector D5. Maximum temperature: 80 oC (176 oF). - under the nut of the ignition unit fixing bolt. See the Installation Drawing, Sheet 1, Sector C5. Maximum temperature: 60 oC (158 oF) for ignition set LUN 2201 - air temperature above hooking eye of the accessory gearbox in distance 25mm (1inch) Maximum temperature: 100 oC (212 oF).


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Unlike the external cooling system, this system employs an internal flow of cooling air. This concerns the LUN 2132.03-8 or APC 250 SG 125Q starter/generator and the LUN 2102.01 alternator. It is therefore necessary for the nacelle outer surface to have air intakes (scoops) utilising the ram effect to supply air to these devices. The air intakes, as well as the shape and length of air feeding hoses have to be so designed to ensure air supply with minimum pressure losses. The shape and dimensions of the cooling air intake are shown on the Installation Drawing, Sheet 7.

Cooling of the LUN 2132.02-8 Starter/Generator - permanent nominal load of 200 A; ram air overpressure of min. 1.2 kPa (0.17 psi) at intake vs pressure in the nacelle; operation at decreased overpressure is limited to a period of 15 minutes. Overpressure must not be lower than 0.7 kPa (0.1 psi). - reduced load of 100 A; operation at no ram effect is only possible for a period of up to 30 minutes.

Cooling of the APC 250 SG 125Q Starter/Generator - permanent nominal load of 250 A; ram air overpressure of min. 1.5 kPa (0.21 psi) at intake vs pressure in the nacelle.

Cooling of the LUN 2102.01 Alternator - permanent nominal load of 3/3.7 kW, ram air overpressure of min. 1.2 kPa (0.17 psi) at intake vs pressure in the nacelle; - operation at an overpressure reduced to 1 kPa (0.14 psi) and 0.7 kPa (0.10 psi) is limited to a period of up to 10 and 3 minutes respectively.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 10 EXHAUST SYSTEM The exhaust system directs gas from the power turbine into ambient atmosphere. It consists of an outlet duct, an outlet casing, two inserts, and two exhaust nozzles. The outlet duct and inserts are located in the outlet casing and are parts of the engine layout. The exhaust nozzles and inserts are fastened by bolts to twin opposed exhaust flanges on the outlet casing. The exhaust nozzle flange is shown, incl. all assembling dimensions on Installation Drawing, Sheet 6, View Z3. The exhaust nozzles turn the stream from radial into a roughly axial, slightly downward direction. The shape of the nozzle liner is such as to guide the stream to flow in the required direction. The stream of the exhaust gas produces an additional thrust. CAUTION:

THE L.H. AND R.H. EXHAUST NOZZLES DIFFER FROM EACH OTHER AND ARE NOT INTERCHANGEABLE.

The nozzles pass through openings in the engine nacelle. On the cold engine a clearance of 20 mm (0.8 in) must be maintained along the whole periphery between the engine nacelle and the exhaust nozzle. Exhaust nozzles Nos. M601-418.7 (L.H.) and M601-419.7 (R.H.) are included in the engine standard equipment. Exhaust nozzles differing in shape and design from those described above can also be installed if required, but these must be tested and approved by the engine manufacturer. CAUTION:

1. WHEN ASSEMBLING OR DISASSEMBLING THE EXHAUST NOZZLES, IT IS NECESSARY FOR THE INSERT TO BE LOCATED IN THE OUTLET CASING IN ORDER TO AVOID SIMULTANEOUS REMOVAL OF THE BOTH NOZZLES AND HENCE TO PREVENT THE INNER PART OF THE EXHAUST SYSTEM FROM LOOSENING. 2. THE DIFFERENT SHAPE OF THE R.H. AND L.H. EXHAUST NOZZLES LEADS TO THE NECESSITY OF ASSEMBLING THEM SO THAT THEIR RECESS R 10 (R 0.4 in) ON THE MOUNTING FLANGE WILL BE ALWAYS DIRECTED UPWARDS. PRESCRIBED

SELFLOCKING

WASHERS

MUST

BE

USED

IN

ASSEMBLING.




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 11 AIRBLEED SYSTEM The air bleed for airframe purposes is used for cabin conditioning/pressuring and de-icing the lips of the engine nacelle air inlet and ducts providing internal cooling of both the starter/generator and electric alternator. No further of the engine proper need to be de-iced by air supplied from the compressor. The air supplied to the airframe is bled from the centrifugal compressor outlet via a thermally insulated manifold terminated on the engine's rear bulkhead by a flange and a airbleed pipe union shown on Installation Drawing, Sheet 3, Sector E3. Air pressure and air flow rate which can be bled from the compressor are shown on Diagram 3-3. As follows from the diagram the flow rate of bled air for makes up 62 g/s (0.136 lb/s) at altitude of 4,200 m (13,780 ft), TAS of 400 km/hr (216 kt) and tH ≤ -10 oC (14 oF). This flow rate may not be used at take-off. In this case 25 g/s (0.055 lb/s) of air can be bled for de-icing purposes only. With airbleed on the ITT increases by the value shown on Diagram 3-4 and hence the engine rating shall be set so that the interturbine temperature will not exceed the maximum permitted value.

Bleed Air Purity Harmful substances in the air bled from the engine do not exceed the following values for - carbon monoxide: one part per 50,000 parts - nitrogen oxide: one part per 300,000 parts.




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 12 ENGINE ACCESSORIES This Section contains a list of instruments and accessories which provide for correct engine operation and its check. Non-standard equipment is also included. Speeds, torques and overhang moments of instruments mounted on the engine are presented. Diagram of instrument position on accessory gearbox and reduction gearbox is given in the Fig. 12-1. Basic engine equipment: Description

Designation

Pcs

Fuel control unit

LUN 6590.05-8

1

Fuel pump

LUN 6290.04-8

1

Starter/generator

LUN 2132.02-8 1 or APC 250SG 125Q

Ignition unit

LUN 2201.03-8 or UNISON 90 49 765-1

2

Propeller speed governor

LUN 7816-8

1

Min. oil quantity signaller

LUN 1581-8

1

Min. oil pressure switch

NOTE

on special order

1 on special order

1.25 K LUN 1469.32-8

1

Electrohydraulic actuator

LUN 7880.01-8

1

on special order

Pressure switch of propeller automatic feathering system

LUN 3280-8

1

on special order

Torque limiter pressure switch

LUN 1476-8

1

Integrated speed transmitter (gas generator)

LUN 1333.12-8

1

Integrated speed transmitter (propeller)

LUN 1333.12-8

1

Interturbine temperature transmitter

LUN 1377-8

9

Torquemeter transmitter

LUN 1540.02-8

1

Oil temperature transmitter

LUN 1358-8

1

Engine mount

M601-907.9

3

Additional Alternator Gearbox

M601-55.6

1

on special order

Engine mount ring

B 061 001

1*

on special order

on special order

* Part of the airframe installation


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Accessories: Exhaust nozzles - L.H. model

M601-418.7

1

on special order

- R.H. model

M601-419.7

1

on special order

Instruments installed in the airframe (enclosed in the container): Description

Designation

Pcs

NOTE

Torquemeter indicator

LUN 1539.02-8

1

on special order

Oil pressure transmitter

LUN 1558-8

1

on special order

Fuel pressure transmitter

LUN 1559-8

1

on special order

NOTE: The engine have not to be fitted with the following equipment (on special order): a)

LUN 3280-8 autofeathering pressure switch, if the airplane is not fitted with an autofeathering system.

b)

LUN 7880.01-8 electrohydraulic actuator, if airplane is fitted with neither autofeathering nor manual feathering system.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Instruments installed in the airframe, which provide for operation and check of the engine. Nonstandard equipment, which is not included in the standard equipment (these instruments remain a part of the aiframe): Description

Designation

Pcs

NOTE

Integrated electronic limiter unit (IELU)

LUN 5260.04

1

*

Engine Limiting Unit

LUN 5224

1

*

Fuel filter

LUN 7691.01-8

1

*

Speed indicator (Gas generator)

LUN 1347.03-8

1

on special order

Speed indicator (Propeller)

LUN 1348.03-8

1

on special order

Interturbine temperature indicator Three pointer indicator of

LUN 1370.02-8 LUN 1538.01-8

1 1

on special order on special order

Feathering pump

LUN 7840-8

1

on special order

Pressure switch 0.7 S

LUN 1492.04-8

1

on special order

Propeller time relay

LUN 2601-8

1

on special order

Starting system time relay

LUN 2601.01-8

1

Starting system interrupter

LUN 3191.91-8

1

Alternator

LUN 2102.01

1

or

- oil temperature - oil pressure - fuel pressure (at fuel nozzle inlet)

on special order

* The following instruments shall be always ordered for „first“ new engine installation in the airframe: LUN 5260.04 IELU and LUN 7691.01-8 fuel filter. These instruments remain part of the airframe installation and they are not removed during engine changes.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Cockpit signallers Indication

Description

1.

Min. oil pressure

2.

Chips in oil - reduction gearbox + accessory gearbox

- „MIN OIL PRESSURE“

- „CHIPS“

3.

Min. fuel pressure

- „MIN FUEL“

4.

BETA control

- „BETA“

5.

Electrohydraulic transducer on

- „ELECTROHYDRAULIC TRANSDUCER“ („EHT“, „ALERT IELU“, or „IELU OPERATIVE“)

6.

Parameter exceeding

- „PARAMETER EXCEEDING

7.

Emergency circuit switched on

- „ISOL. VALVE“

The following signallers are used in the case of autofeathering and manual feathering system application: 8. 9.

Switching on of microswitch on the ECL when autofeathering system is on

- „AUTOFEATHER“

Switching on of feathering pump

- „FEATHER PUMP“

Optional (fuel filter): 10.

Fuel filter by-pass (open)

- „FUEL BYPASS OPEN“

Optional (oil tank): 11.

Min. oil quantity

- „MIN OIL“

Optional changes of transmitters, indicators and signallers must be discussed and approved by the engine manufacturer.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Non standard equipment: The engine can be fitted by the following instruments in addition to the standard ones: De-icing system of the propeller unit. This device through pilot's intervention provides for de-icing of the propeller blades in icing conditions. A.C. supplied by the alternator is utilized for propeller blade deicing. Brush bracket is a part of the system; it transfers voltage from the stator part to the propeller rotor. The following instruments are used for deicing of the double acting VJ8.510 propeller unit: Propeller fitted with de-ice boots fed by voltage of 115/200 V AC: LUN 7850-7 brush block assembly

is situated on the reduction gearbox.

LUN 2102.01 alternator

of power 3 to 3.7 kVA, U = 115/200 V, n = 6,300 to 10,150 r.p.m. The alternator is situated on the accessory gearbox and is driven via an additional gearbox, which is installed together with the generator. Integral part with the alternator is its ØJG 3/3.7 kVA regulator, which is installed in the airframe. Cyclic switch for propeller blade heating is also installed in the airframe. The alternator can be utilised for cockpit windscreen deicing and further airframe needs.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Instruments mounted on the accessory gearbox 100 % nG = 36,660 r.p.m.

Item

Sense of drive rotation

Instrument

Speed ratio

Starter 1

Starter/generator

2

Spare drive (hydraulic pump drive)

3

D.C. generator

Fuel Pump

Fuel Control Unit

5

Integrated Gas Generator Speed Transmitter

6a 6b

Overhang moment Nm (lb.in)

11.2 (100) CW

0.2898

9.2

(82)

21 (186)

CCW

0.1974

5.8

(51)

40 (353)

CW

4

Max. Torque Nm (lb.in)

0.1196

4.5

(40)

CCW

0.1223

1.1

(10)

CW

0.1145

0.5

(4,5)

Manual turning by

CCW

0.1145

11.5 (102)

Alternator (manual turning by)

CCW

0.2763

5

This instrume nt must be used This instrume nt must be used 4 (35.4) 0

(44)

40 (353)

Instruments mounted on the reduction gearbox nV = 2,080 r.p.m. 7

Propeller Speed Governor LUN 7816-8

CCW

2.0285

5.7

(50)

7 (62)

8

Integrated Propeller Speed Transmitter LUN 1333.12-8

CCW

2.0285

0.5

(4.5)

4 (35.4)



Facing the Reduction Gearbox

Facing the Accessory Gearbox

For figure description refer to Page 12-6. ENGINE MOUNTED ACCESSORIES Fig. 12-1



Required overall accuracies of measurement and ranges for instruments procured by organization performing the engine installation into airframe. Range Gas Generator Speed (nG)

Propeller Speed (nV)

Interturbine

Accuracy

0 to 110 %

0 to 2500 rpm

Temperature 0 to 900 oC

Oil Temperature (tOIL)

Torque (TQ)


±1.5 %

60 to 110 %

±0.5 %

0 to 1500 rpm

±1.5 %

1500 to 2500 rpm

±0.5 %

600 to 680 oC

o ±15 C

o

(ITT)

Oil Pressure (pOIL)

0 to 60 %

0 to 0.4 MPa

o

-30 to 150 C

0 to 120 %

o

680 to 725 C

±10 C

o 725 to 900 C

±15 C

0.1 to 0.3 MPa

±0.017 MPa

otherwise

±0.024 MPa

o

o

0 to 90 C

o ±4 C

o 90 to 120 C

o ±7 C

80 to 109 %

±2 %

otherwise

±3.5 %

WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 13 COOLANT INJECTION The engine can be fitted for coolant injection at higher atmospheric temperatures as described in the „Operation Manual“. The coolant is injected through nozzles installed on the spray ring surrounding the compressor inlet protecting screen. The spray ring is connected to the coolant tank and the pump via the coolant injection flow adapter on the rear bulkhead; see Installation Drawing, Sheet 3, Sector F4. The spray ring is used also for compressor recovery washing. Three stages of coolant injection can be distinguished in response to ambient conditions as shown in the „Operation Manual“. Pressures required for the various stages at the spray ring inlet must attain the following values: Stage of coolant injection

Coolant flow

Inlet spray ring pressure

l/min

(US gal/min)

MPa

(psi)

I.

1.65

(0.437)

0.075

(10.6)

II.

3.3

(0.873)

0.26

(37.0)

III.

5.0

(1.323)

0.46

(65.0)

The coolant is injected for a period of max. 5 minutes. The coolant supply adapter on the rear bulkhead can also held a device for washing the compressor as described in the „Maintenance Manual“, Section 72.03.00, Page 701. Coolant injection is put in operation by means of cockpit push button. Series connection of terminals No 11 and 12 on the V8 connector (see Fig. 7-1) into push button circuit is required to avoid engine shut down when gas generator rotational speed is decreased without preceding coolant injection switching off.




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 14 PROPELLER UNIT Five blade double acting (Avia-Propeller) propellers can be installed on the engine. Necessary control oil pressure in the propeller hub must not exceed 2.9 MPa and max. permitted oil pressure in the propeller hub should be greater than 2.7 MPa in the sea level static conditions. The propeller must not exceed mass limits shown in Fig. 14-1. If mass characterics of the propeller exceed curve (refer Fig. 14-1) WALTER a.s. must be consulted. Distance between propeller center of gravity and engine propeller flange must not exceed 110 mm (4.33 in). The engines are normally fitted for installation of VJ8.510 propeller unit.

Propeller mass

83

Polar mass moment of inertia

10.1 kgm2

34,512 lb in2

5.95 kgm2

20,300 lb in2

Mass moment of inertia about the axis perpendicular to the axis of rotation

kg

183 lb

THE PROPELLER MASS AND DYNAMIC CHARACTERISTICS OF V510 PROPELLER Table 14-1

Max. adjustable reverse power: Double acting five blade Avia propeller, dia 2,300 mm (90.5 in)

app. 325 kW (436 SHP)

There is a pad on the reduction gearbox which is utilized for installation of the brush block assembly when propeller deicing is required (see Installation Drawing, Sheet 7, View T2, Section N4-N4). Avia-Propeller delivers brush block assembly for deiced Avia propellers which can be installed without any modification on WALTER M601 engines; refer to Section 12, Nonstandard equipment.


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 PROPELLER FEATHERING 1)

Emergency feathering Emergency feathering is carried out by means of the propeller control lever. This feathering system does not depend on airframe systems.

2)

Manual propeller feathering (all ratings) effected by means of the push button Propeller is feathered when voltage of 27 volts is fed to the connector of LUN 7880.01-8 electrohydraulic actuator. The system shall ensure simultaneous voltage feeding to emergency control circuit to protect the engine. Propeller manual feathering system of the double acting propellers includes an auxiliary pump which provides for feathering of the stopped engine and even in the case of engine oil loss.

3)

Propeller autofeathering Propeller autofeathering system is an airframe system applied mainly on twin engine airplanes. The system evaluates signals from switch on the engine control lever and from LUN 3280-8 propeller autofeathering pressure switch. The system shall provide for proper operation (with respect to the engine) the following functions: a)

delay of min 4 sec of the signal from the switch closing on the engine control lever.

b)

Switching of the emergency circuit on simutaneously with voltage feeding to LUN 7880.01-8 electrohydraulic actuator and feathering pump if they are included in the feathering system of the double acting propellers.



2

Propeller Mass Polar Moment of Inertia JPVRT [kgm ]

25

20

15

10

5

0 40

50

60

70

80

90

100

Propeller Total Mass (incl. Oil Charge) [kg]

LIMITS OF PROPELLER MASS CHARACTERICS Fig. 14-1




WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 SECTION 15 ENGINE CONTROLS The engine operation is controlled by the following levers: (Schematic diagram is shown in the Fig. 15-1) − engine control lever (ECL) − fuel shut-off valve actuating lever − propeller control lever (PCL)

1

Engine control lever in the cockpit

2

Fuel shut-off valve actuating lever in the cockpit

3

Propeller control lever in the cockpit

4

Engine control lever on the engine

5

Fuel shut-off valve actuating lever on the engine

6

Rope conduit to the PCL on the engine Dwg. No M601-77.6 (optional)

7

Interface for the airplane installation of the PCL (in case of installation of the rope conduit Dwg. No M601-77.6)

8

The PCL on the engine; interface for the airframe installation in case that the rope conduit Dwg. No M601-77.6 is not installed. ENGINE CONTROLS - SCHEMATIC DIAGRAM Fig. 15-1


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Engine Control Lever The engine control lever (ECL) sets the required gas generator speed (engine power) at the forward thrust ratings and the fuel delivery (engine power) at BETA range. At BETA range the engine power corresponds to such a reverse propeller angle that prevents the propeller from overspeeding. At forward thrust ratings the displacement of engine control lever is limited by a stop on the fuel control unit. The stop corresponds to selected gas generator speed of 100 %. At the reverse thrust ratings the displacement of the ECL is limited by a stop that corresponds to max. reverse power.

Fuel Shut-Off Valve Actuating Lever The fuel shut-off valve actuating lever opens and closes engine fuel delivery. With emergency circuit on the lever sets the engine power.

Propeller control lever Propeller control lever sets controlled propeller speed. The lever effects emergency feathering in one extreme position.

Airframe Engine Actuating System Requirements Necessary data concerning airframe interface of individual engine levers are presented in the Installation Drawing, Sheet 6.

Airframe part of actuating system shall provide for: Engine Control Lever: a) Retractable stop on the ECL in cockpit for the idle position. The stop prevents from undesirable displacement of the lever from forward thrust ratings to the reverse thrust and vice versa. Retractable stop from side of forward thrust ratings specifies the ECL position for starting and idle. b) Min. value of actuating moment of 12 N.m (8.8506 lb.ft) on the ECL for max. reverse power setting. This rating is set by an ECL stop. c) Min. value of actuating moment of 1.5 N.m (1.1063 lb.ft) on the ECL on the engine for take-off rating setting. Airframe part of the actuating system is joined to the engine actuating system in the Item 4 (see Fig. 15-1).


WALTER a.s. INSTALLATION MANUAL MANUAL PART No. 0982502 Fuel Shut-Off Valve Actuating Lever a) Retractable stop on the fuel shut-off valve actuating lever in the cockpit for „OPEN“ position („CLOSED“ and „MAX. SPEED“ positions - if power is controlled by the emergency circuit - are ensured by stops on the fuel control unit). b) Min. value of actuating moment of 3 N.m (2.2126 lb.ft) on the fuel shut-off valve actuating lever on the engine to enable reaching of extreme positions of the lever. Airframe part of actuating system is joined to the fuel shut-off valve actuating lever on the engine in the Item 5 (see Fig. 15-1).

Propeller Control Lever a) Retractable stop on the PCL in the cockpit for „MIN. CONTROLLED SPEED“ position. (Lever positions corresponding to „MAX. CONTROLLED SPEED“ and „PROPELLER FEATHERING“ are ensured by stops on the propeller governor.) b) Min. value of actuating moment of 2 N.m (1.4751 lb.ft) on the PCL on the engine to enable reaching of extreme positions of the lever (if the rope conduit Dwg. No M60177.6 is installed the force on the rods of the airframe system is increased by app. 20 %). Airframe part of actuating system is joined to the PCL on the engine in the interface: 7 in case of installation of the rope conduit Dwg. No M601-77.6 (see Fig. 15-1) 8 in case that the rope conduit Dwg. No M601-77.6 is not installed (see Fig. 15-1)





ENGINE INSTALLATION DRAWING Sheet 1

Appendix 1

Jun 26, 2001



Appendix 2

Jun 26, 2001



Appendix 3

Jun 26, 2001


ENGINE INSTALLATION DRAWING SHEET 4

Appendix 4

Jun 26, 2001



Appendix 5

Jun 26, 2001



Appendix 6

Jun 26, 2001


62 (61.97)*

COOLING AIR INLET

88 (88.9)

+1.5 *

)

*

(5.6 )

+1.5 *

145 (142.7)*

(1.3

DIMENSIONS MARKED (…)* ARE VALID FOR STARTER-GENERATOR MODEL 250 SG 125 Q APC USA MANUFACTURER

ENGINE INSTALLATION DRAWING (294)*

Sheet 7

300

Appendix 7 May 11, 2006



Appendix 8

Jun 26, 2001

Installation Manual Walter m601e, m601e-21[1]

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