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


RECORD OF REVISIONS Page 0-1

May 11, 2006

RECORD OF REVISIONS

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

REVI-SION No.

BULLETIN No.

ISSUE DATE OF NEW PAGES

NUMBERS OF AFFECTED PAGES

DATE OF INSERTION AND SIGNATURE

1 Nov 15, 2004 Pages 0-1, 0-3, 0-4, 1-1, 2-1 to 2-8, 3-2, 3-10 to 3-38, 5-1, 5-7, 5-8, 8-3, 9-1, 9-2, 12-1, 12-2, 12-3, 12-6, 12-8, 13-1

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



RECORD OF REVISIONS Page 0-2 Jun 26, 2001

REVI-SION No.

BULLETIN No.

ISSUE DATE OF NEW PAGES

NUMBERS OF AFFECTED PAGES

DATE OF INSERTION AND SIGNATURE



REVIEW OF EFFECTIVE PAGES - SHEETS Page 0-3

May 11, 2006

REVIEW OF EFFECTIVE PAGES - SHEETS

Section Page - Sheet Date Title Sheet - in English Jun 26, 2001- in Czech 26.6.2001 Record of 0-1 May 11, 2006revisions 0-2 Jun 26, 2001 Review of effective 0-3 May 11, 2006sheets - pages 0-4 May 11, 2006 Contents 0-5 Jun 26, 2001 0-6 blank Jun 26, 2001 Introduction 0-7 Jun 26, 2001 0-8 Jun 26, 2001 Section 1 1-1 Nov 15,2004Engine 1-2 Jun 26, 2001Description 1-3 Jun 26, 2001 1-4 Jun 26, 2001 Section 2 2-1 Nov 15,2004Engine 2-2 Nov 15,2004Operation Limits 2-3 Nov 15,2004 2-4 Nov 15,2004 2-5 Nov 15,2004 2-6 Nov 15,2004 2-7 Nov 15,2004 2-8 blank Nov 15, 2004 Section 3 3-1 Jun 26, 2001Engine 3-2 Nov 15,2004Performance 3-3 Jun 26, 2001 3-4 Jun 26, 2001 3-5 Jun 26, 2001 3-6 Jun 26, 2001 3-7 Jun 26, 2001 3-8 Jun 26, 2001 3-9 Jun 26, 2001 3-10 Nov 15,2004 3-11 Nov 15,2004 3-12 Nov 15,2004 3-13 Nov 15,2004 3-14 Nov 15,2004

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



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

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

Section Page - Sheet Date Section 12 Engine 12-1 Nov 15,2004Accessories 12-2 Nov 15,2004 12-3 Nov 15,2004 12-4 Jun 26, 2001 12-5 Jun 26, 2001 12-6 May 11, 2006 12-7 Jun 26, 2001 12-8 Nov 15,2004 Section 13 Coolant 13-1 Nov 15,2004Injection 13-2 blank Jun 26, 2001 Section 14 Propeller Unit 14-1 Jun 26, 2001 14-2 Jun 26, 2001 14-3 Jun 26, 2001 14-4 blank Jun 26, 2001 Section 15 Engine Controls 15-1 Jun 26, 2001 15-2 Jun 26, 2001 15-3 Jun 26, 2001 15-4 blank Jun 26, 2001 Appendix 1 Jun 26, 2001 2 Jun 26, 2001 3 Jun 26, 2001 4 Jun 26, 2001 5 Jun 26, 2001 6 Jun 26, 2001 7 May 11, 2006 8 Jun 26, 2001



CONTENTS Page 0-5

Jun 26, 2001

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 0-6 Jun 26, 2001



INTRODUCTION Page 0-7

Jun 26, 2001

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 0-8 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.



Section 1 Page 1-1

May 11, 2006

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 1-2 Jun 26, 2001

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 1-3

Jun 26, 2001

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



Section 2 Page 2-1

Nov 15,2004

SECTION 2

WALTER M601E1/E-21 ENGINE OPERATION LIMITS

Operation Limits:

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

Starting at low ambient temperature:

- without preheating: at oil temperature higher than -20 °C

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

The preheating can be stopped when oil temperature reaches +5 °C. Max. gas generator speed: 100 % (100 % = 36,660 rpm) Max. propeller speed: 2,080 rpm Max. interturbine temperature: 735 °C Max. torque of 2,725 Nm (2,010 lb.ft): 106 % (100 % = 2,570 Nm = 1,896 lb.ft) Max. oil temperature: 85 °C Acceptable gravitational load factors:

longitudinal nx = ± 1.5 g vertical ny = + 4.15 g (+ 5 g short-period at landing, short-period at gust load)

- 2.15 g lateral nz = ± 1.5 g ωy = ± 0.45 rad/sec ωz = ± 0.60 rad/sec

Permitted loads at flight manoeuvres are shown in Fig 4-3, Page 4-9 (for coordinate system refer to Page 4-3). For further operation limits see the „Operation Manual“. The engine is approved to operate in severe ice-forming conditions. If interturbine temperature (ITT), torque or propeller speed exceed their maximum permitted values, it is necessary to evaluate them and proceed in accordance with the applicable Diagrams 2-1, 2-2, 2-3 and Table 2-1 when taking further measures for engine operation.



Section 1 Page 2-2 Nov 15, 2004

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

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

2. Check the proper function of the limiter system.

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

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

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

OVERTEMPERATURE LIMITS - STARTING CONDITIONS ONLY

Fig. 2-1

770

780

750

760

730

720

740

30 5 0 10 15 20 25

ITT [°C] INTERTURBINE TEMPERATURE

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

TIME [sec]

NO ACTIONS REQUIRED

AREA „A“

AREA „B“



Section 3 Page 2-3

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.

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

Diagram 2-2a

TIME [sec]

780

800

740

760

700

720



NO ACTIONS REQUIRED

0 20 40 60 80 100 120

AREA „A“

AREA „B“




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

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

OVERTEMPERATURE LIMITS – ALL CONDITIONS EXCEPT STARTING

(Valid for power ratings defined for the event of OEI flight)

Fig. 2-2b


0 1 2 3 4 5 6 7 8 9 10 11 12

TIME [minutes]

780

790

760

770

740

750


NO ACTIONS REQUIRED

AREA „A“

AREA „B“

730



Section 3 Page 2-5

Nov 15, 2004

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

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

OVERTORQUE LIMITS

Diagram 2-3

103

104

109

108

107

106

105

101

102

99

98

100

61 0 2 3 4 5

PROPELLER TORQUE [%]


TIME [minutes]

111

110 AREA „A“

AREA „A“

NO ACTIONS REQUIRED




Propeller speed [rpm]

Measures

up to 2,220 No action required

2,220 to 2,300 Overspeed not longer than 20 sec: Record the rpm in the Engine Log Book. If occurrence of these overspeeds exceeds number 10, the engine must be returned to an overhaul facility for inspection/repair acc. to Overhaul Manual.

Overspeed longer than 20 sec: Ref. to the Propeller Operation Manual

2,300 to 2,400 1) Record the rpm in the Engine Log Book. If occurrence of these overspeeds exceeds number 2, the engine must be returned to an overhaul facility for inspection/repair acc. to Overhaul Manual.

2) Inspect chip detectors and oil filter cartridge for contamination with metal chips. Refer to Section 79.10.00, Maintenance Manual.

3) After engine shut-down turn-by propeller manually. Check for symptoms of power turbine blades seizing (unusual noise). This repeat at 10 min and at 20 min after engine shut-down.

4) Record the results of the check (Item 2) in the Engine Log Book.

5) If the propeller can be manually turned in all three checks without any symptoms of seizing, the engine can continue in operation for remaining T.B.O. without any limitation.

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

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

PROPELLER OVERSPEED LIMITS

Table 2



Section 2 Page 2-7

Nov 15, 2004

ENGINE OPERATING ENVELOPE

Diagram 2-5

H [m] H [ft]

18,000

20,000

14,000

16,000

10,000

12,000

6,000

8,000

6,000

5,000

4,000

3,000

2,000

1,000

2,000

000 0.1 0.2 0.3 0.4 0.5 0.6

4,000

Airbleed Envelope

Flight Mach Number

In-Flight Start






Section 3 Page 3-1

Jun 26, 2001

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.




The following nomenclature is used in this section:

JAR - Joint Aviation Requirements

FAR - Federal Aviation Regulation

F - jet thrust

ISA - International Standard Atmosphere (same as the ICAO Standard Atmosphere Conditions)

ITT - interturbine temperature

nG - gas generator speed

LHV - lower heating value

OAT - outside air temperature

P - absolute pressure

T - absolute temperature

NH - shaft power

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

∆ - difference

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

W - mass flow

Subscripts:

AM - ambient static values of undisturbed air

BL - bleed

F - fuel

G - generator

IN - inlet duct

N - net

TAS - true air speed

C - total

R - corrected

Temperature Conversion Formulae: °F = 1.8 (°C + 40) - 40

°R = °F + 459.67

K = °C + 273.15



Section 3 Page 3-3

Jun 26, 2001

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 %.




Correction on Installation Losses

Engine installation into the aircraft effects the engine performance in two ways: first,

the power decreases due to pressure losses in the inlet duct; second, the power

decreases due to leakage of exhaust gas in the compressor intake. Entrainment of hot

air/gas from the engine compartment results in some warming of air entering the

compressor. At reverse engine layout when the inlet duct is situated under the engine

the elimination of this effect is very difficult.

a) Pressure Loss Influence:

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

than 90 % pressure loss in the air inlet system ∆PP IN

causes the following

changes in the engine parameters:

Relative power decrease:

∆ ∆NN

1.95P

PH

H IN

= ∗

Relative fuel consumption decrease:

∆ ∆WW

0.9 P

PF

F IN

= ∗

Interturbine temperature increase:

[ ]∆∆

ITT 100 P

P C

IN

= ∗

°

b) Air Warming Influence:

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

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

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

Interturbine temperature increase must be considered if the ITT limit for given

power rating can be exceeded. In this case it is necessary to decrease the gas

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

decreases app. by 17 oC (30.6 oF), power at the same time decreases by 6.4 %

and fuel consumption by 4.4 %.



Section 3 Page 3-5

Jun 26, 2001

The relative pressure loss ∆PP IN

and air temperature increase in the air inlet

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:

T TV

3.6

12009C AM

2

= +

∗ [ K ]

T T TIN C IN= + ∆ [ K ]

P P TT

C AMC

AM

3.5

= ∗

[ kPa ]




P P 1 PPIN C

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 ]

The values of ∆TIN and ∆P P IN

in flight conditions can be determined by means of

Diagram 3-2 for relevant engine model. At zero flight speed at the first step we

calculate the value of nG / Θ with TAM instead of TIN. Then we read in the Diagram 3-

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

At the second step we calculate new value of 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

of TIN and nG/ Θ and then the value of ∆P P IN

can be read in the Diagram 3-1.

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

follows:

T TV

1.9438

11116C AM

2

= +

∗ [ °R ]



Section 3 Page 3-7

Jun 26, 2001

T T TIN C IN= + ∆ [ °R ]

P P TTC AM

C

AM

3.5

= ∗

[ psi ]

P P 1 PPIN C

IN

= ∗ −

∆ [ psi ]

Where:

V = true air speed [ kt ]

TAM = outside air temperature [ °R ]

PAM = ambient pressure [ psi ]

TC = total temperature [ °R ]

PC = total pressure [ psi ]

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

∆P P IN

= relative pressure loss in the air inlet system

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

PIN = total pressure in the compressor inlet [ psi ]

Calculate the corrected gas generator speed nG / Θ [ % ]

In the metric system of units: nG / Θ = nG 288TIN

[ % ]

In the U.S. standard units: nG / Θ = nG 519TIN

[ % ]

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:

T T 273.8 n100

282 BL ING

2

= + ∗

− [ °C ]

Where TIN is in Kelvins, nG is in %.

In the U.S. standard units:

T T 492.8 n100

475.6 BL ING

2

= + ∗

− [ °F ]

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

The engine response to airbleed is as follows:

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

increase in fuel consumption and ITT. This increase can be determined from the curves

in Diagram 3-4, where the difference in fuel consumption with airbleed open/closed

∆WW

F

F BL

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

equivalent diaphragm dia. When the ITT exceeds the limit for the given rating it is

necessary to decrease the gas generator speed. The relation between the 1 %

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

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

Correction on Power Off-Take for Electric Generator.

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

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

100 A of generator loading. Should the ITT increase exceed the limit for the given

rating, it is necessary to decrease the gas generator speed.



Section 3 Page 3-9

Jun 26, 2001

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.




rating

shaft power

equivalent shaft power

ESFC max. gas generator

speed

propeller speed

torque max. interturbine temperature

[kW] [SHP]

[kW] [ESHP]

[g/kW/hr] [lb/ESHP/hr]

GT [%]

[rpm] [N.m] [lb.ft]

[°C]

take-off (5 min)

15 °C (59.0 °F) 560

751

595

798

395

0.6493 98.6 2080

2570

1895 710

sea level static

23 °C (73.4 °F) 560

751

595

798 - 100 2080

2570

1895 735

max. continuous

15 °C (59 °F) 490

657

521

699

410

0.674 96.5 1700 to

2080

2570

1895 680

sea level static

18 °C (64.4 °F) 490

657

521

699 - 97 1700 to

2080

2570

1895 690

take-off with water injection

300 l/hr (79 US gal/hr)

(5 min)

97.325 kPa (14.12 psi)

33 °C (91.4 °F)

560

751

595

751

-

100

2080

2570

1895

735

intermediate contingency

sea level static

28 °C (82 °F)

560

751

595

798 - 100.5 2080

2570

1895 760

maximum contingency

(10 min)

97.325 kPa (14.12 psi)

28 °C (82.4 °F)

595

798

630

845

-

102

2080

2737

2019

780

NOTE: gas generator speed 100 % = 36,660 rpm 2,080 propeller rpm = 31,023 power turbine rpm

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

WALTER M601E Table 3-1



Section 3 Page 3-11

Nov 15, 2004

rating

shaft power



speed

propeller speed


[kW] [SHP]

[kW] [ESHP]


GT [%]

[rpm] [N.m] [lb.ft]

[°C]

take-off (5 min)

15 °C (59.0 °F) 560

751

595

798

389

0.64 98.1 2080

2570

1895 690

sea level static

28 °C (82.4 °F) 560

751

595

798 - 100 2080

2570

1895 735

max. continuous

15 °C (59 °F) 490

657

521

699

405.9

0.667 96.2 1700 to

2080

2570

1895 660

sea level static

21 °C (69.8 °F) 490

657

521

699 - 97 1700 to

2080

2570

1895 690



(5 min)

sea level static

42 °C (107.6 °F)

560

751

595

751

-

100

2080

2570

1895

735

intermediate contingency

sea level static

32 °C (90 °F)

560

751

595

798 - 100.5 2080

2570

1895 760

maximum contingency

(10 min)

97.325 kPa (14.12 psi)

31.5 °C (88.7 °F)

595

798

630

845

-

102

2080

2737

2019

780


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

WALTER M601E-21 Table 3-2




rating

shaft power



speed

propeller speed


[kW] [SHP]

[kW] [ESHP]


GT [%]

[rpm] [N.m] [lb.ft]

[°C]

take-off (5 min)

15 °C (59.0 °F) 560

751

595

798

395

0.6493 98.6 2080

2570

1895 710

sea level static

23 °C (73.4 °F) 560

751

595

798 - 100 2080

2570

1895 735

climb and max. cruise

15 °C (59 °F) 490

657

521

699

410

0.674 96.5 1700 to

2080

2570

1895 680

sea level static

18 °C (64.4 °F) 490

657

521

699 - 97 1700 to

2080

2570

1895 690



(5 min)

97.325 kPa (14.12 psi)

33 °C (91.4 °F)

560

751

595

751

-

100

2080

2570

1895

735

max. continous

sea level static

28 °C (82 °F)

560

751

595

798 - 100.5 2080

2570

1895 760

Max. take-off

(5 min)

97.325 kPa (14.12 psi)

28 °C (82.4 °F)

595

798

630

845

-

102

2080

2737

2019

780


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

WALTER M601E

Diagram 3-3



Section 3 Page 3-13

Nov 15, 2004

rating

shaft power



speed

propeller speed


[kW] [SHP]

[kW] [ESHP]


GT [%]

[rpm] [N.m] [lb.ft]

[°C]

take-off (5 min)

15 °C (59.0 °F) 560

751

595

798

389

0.64 98.1 2080

2570

1895 690

sea level static

28 °C (82.4 °F) 560

751

595

798 - 100 2080

2570

1895 735

Climb and max. cruise

15 °C (59 °F) 490

657

521

699

405.9

0.667 96.2 1700 to

2080

2570

1895 660

sea level static

21 °C (69.8 °F) 490

657

521

699 - 97 1700 to

2080

2570

1895 690



(5 min)

sea level static

42 °C (107.6 °F)

560

751

595

798

-

100

2080

2570

1895

735

Max. continous

sea level static

32 °C (90 °F)

560

751

595

798 - 100.5 2080

2570

1895 760

Max take-off

(5 min)

97.325 kPa (14.12 psi)

31.5 °C (88.7 °F)

595

798

630

845

-

102

2080

2737

2019

780


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



Section 3 Page 3-15

Nov 15, 2004

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

IN-FLIGHT MEASUREMENT. ENGINE INSTALLED IN THE L410 UVP AIRPLANE.

Diagram 3-2




AIR PRESSURE AT THE AIR BLEED MANIFOLD DELIVERY FLANGE.

VARIATION WITH THE AIR BLEED FLOW RATE AND GAS GENERATOR SPEED.

Diagram 3-3

p BL =

14

696/

p IN

[psi

]



Section 3 Page 3-17

Nov 15, 2004

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



Section 3 Page 3-19

Nov 15, 2004

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


WALTER M601E

Diagram 3-7



Section 3 Page 3-21

Nov 15, 2004




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



Section 3 Page 3-23

Nov 15, 2004

TAKE-OFF RATING

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

WALTER M601E

Diagram 3-10







WALTER M601E

Diagram 3-11



Section 3 Page 3-25

Nov 15, 2004




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



Section 3 Page 3-27

Nov 15, 2004


NET JET THRUST, INTERTURBINE TEMPERATURE. NO INSTALLATION LOSSES - ISA + 20.2 °C CONDITIONS.

WALTER M601E

Diagram 3-14




TAKE-OFF RATING


WALTER M601E-21

Diagram 3-15



Section 3 Page 3-29

Nov 15, 2004

TAKE-OFF RATING


WALTER M601E-21

Diagram 3-16







WALTER M601E-21

Diagram 3-17



Section 3 Page 3-31

Nov 15, 2004




WALTER M601E-21

Diagram 3-18




TAKE-OFF RATING


WALTER M601E-21 Diagram 3-19



Section 3 Page 3-33

Nov 15, 2004

TAKE-OFF RATING


WALTER M601E-21

Diagram 3-20







WALTER M601E-21

Diagram 3-21



Section 3 Page 3-35

Nov 15, 2004

MAX. CONTINUOUS RATING RATING acc. to JAR



WALTER M601E-21

Diagram 3-22






WALTER M601E-21 Diagram 3-23



Section 3 Page 3-37

Nov 15, 2004



WALTER M601E-21

Diagram 3-24






Section 4 Page 4-1

Jun 26, 2001

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.



Section 4 Page 4-2 May 11, 2006

Engine Mass and Dynamic Characteristics

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

Engine dry mass 207 kg 456 lb

Mass of oil charge 7 kg 15.4 lb

Table 4-1

NOTE:

The following standard equipment delivered with the engine is included in the mass:

Fuel pump, fuel control unit, starter/generator, ignition unit, integrated speed transmitters, oil temperature transmitter, interturbine transmitter, oil pressure transmitter, fuel pressure transmitter, three torquemeter transmitters, min. oil pressure switch, min. oil quantity signaller, engine mounts (three pieces).

Additional equipment mass:

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

Alternator gearbox 2.5 kg (5.5 lb)

LUN 2102 Alternator 9.8 kg (21.6 lb)

Engine mounting ring 5.16 kg (13.5 lb)

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

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



Section 4 Page 4-3

Jun 26, 2001

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

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

y

x

z

+ ωy; εy

+ ωz; εz+ ωx; εx

Axis of the engineDirection of flight




Vibration

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

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

plant. Vibration is measured at steady-state condition after 2 to 3 minutes of engine

operation. Vibration sensors are located in a radial direction as follows (Refer to Figure

4-1):

a) on the reduction gearbox (Designation Sn 1), where the first harmonic component of

power-turbine speed is measured.

b) on the accessory gearbox (Designation Sn 3), where the first harmonic component

of generator speed is measured.

SENSOR

Sn 1 Sn 3

Test bed VEF

(mm/sec)

VRMS

(in/sec)

VEF

(mm/sec)

VRMS

(in/sec)

Propeller test bed 10 0.4 5 0.2

VIBRATION LOAD LIMITS

Table 4-3

Mentioned check of the vibration limits after engine installation and during operation is

not necessary.



Section 4 Page 4-5

Jun 26, 2001

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.



Section 4 Page 4-7

Jun 26, 2001

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



Section 4 Page 4-9

Jun 26, 2001

(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

ny ny

2

1

5

4

3

-1

-2

-1-2 nx21

2

1

5

4

3

-1

-2

-1 nx 1

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

Flight 1 (idling to take-off power) nz = ± 1.5 ωy = ± 0.3 ωz = ± 0.6 εx = ± 4 εy = ± 6 εz = ± 6

Flight 2 (idling to take-off power) nz = ± 1.5 ωy = ± 0.45 ωz = ± 0.6 εx = ± 4 εy = ± 6 εz = ± 6






Section 5 Page 5-1

Nov 15, 2004

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.




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.



Section 5 Page 5-3

Jun 26, 2001

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)



Section 5 Page 5-5

Nov 15, 2004

AGRICULTURAL AIRPLANE ENGINE AIR INLET SYSTEM

Fig. 5-2

1 - Air Filter

2 - Aerosol Separator (Demister)

3 - Oil Cooler (outside of air inlet system)




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

VARIATION WITH AIR INLET SYSTEM PRESSURE LOSS

Fig. 5-3

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

1W

WF

F

+∆

1N

NH

H

+∆

∆PP0 IN

1.00

0.98

0.96

0.94 0.01 0.03 0.020

∆ITT (°C)

1.00

0.98

0.96

4

2

0



Section 5 Page 5-7

Nov 15, 2004

0

1

2

3

4

5

6

7

8

9

10

11

12

60 70 80 90 100nGR (%)

∆p/

p STR

100

(%

)

∆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]

INLET AIR DISTORTION LIMIT Diagram 5-4






Section 6 Page 6-1

Jun 26, 2001

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.




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.



Section 7 Page 7-1

Jun 26, 2001

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.




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 chromel-alumel 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).



Section 7 Page 7-3

Jun 26, 2001

Torquemeter Indicator Set

A torque indicator set consists of the LUN 1540.02-8 transmitter and the LUN 1539.02-8 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.




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 Transmitter

resistance


resistance


resistance

(°C) (°F) (Ω) (°C) (°F) (Ω) (°C) (°F) (Ω)

-70 (-94) 68.36 10 (50) 93.76 90 (194) 125.56

-60 (-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.



Section 7 Page 7-5

Jun 26, 2001

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 1538.01-8 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.




Equivalent Resistance of the Pressure Transmitter Including the Inductive Transducer

Fuel Pressure Value

MPa (psi)

Resistance Between Transmitter Pins

No. 2 and 1 (ohm)


No. 2 and 3 (ohm)

0 (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.



Section 7 Page 7-7

Jun 26, 2001

Equivalent Resistances of the Pressure Transmitter Including the Inductive Transducer

Oil Pressure

MPa (psi)


No. 2 and 1 (ohm)


No. 2 and 3 (ohm)

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.




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.



Section 7 Page 7-9

Jun 26, 2001

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 %.




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.



Section 7 Page 7-11

Jun 26, 2001

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




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.



Section 7 Page 7-13

Jun 26, 2001

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.




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.



Section 7 Page 7-15

Jun 26, 2001

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.




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.



Section 7 Page 7-17

Jun 26, 2001

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).




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

11 12 19 13

A B V G D1 2 3 4

11 12 19 13

2 RM 14 KPN 4G VŠ 18 KPN 5G 1U2 VŠ 17 KPN 2G1VŠ 17 KPN 3G 1U2WK 180 36

V 110-2

WK 180 36

CHIP SIGNALLER REDUCTION GEARBOX

UNISON

EXCITER

EMERGENCY CIRCUIT VALVE ON THE LUN 6590.05-8

CHIP SIGNALLER ACCESSORY GEARBOX

TORCH IGNITER FUEL VALVE ON THE LUN 6290.04-8

SWITCH ON THE ECL ON THE ENGINE

MANUAL AND AUTOMATIC SYSTEMSOF PROPELLER FEATHERING AND EMERGENCY CIRCUIT SWITCH LUN 3191-8

Interrupter feathering system

water injection

4 8

4 8

E 94 B 2 B 8 M 28

V8 2RM 33 KPN 20 Š4V1 - 2 RM 33 BPN 20 G4V1

E 96 B 4 M 304

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

V 112-2

2

2

3

3

A B V

N25F-3 (CHAMPION) SPARK PLUG

N25F-3 (CHAMPION)SPARK PLUGV 110-2

5

5

A B

VŠ 17 KPN 2G1

A B

STARTING PANEL

STARTING PANEL

V 110-2

7

5

V 110-1

9

5

WA

LTER a.s.

INSTA

LLATIO

N M

AN

UA

L M

AN

UA

L PAR

T No. 0982502

Section 7P

age 7-19/7-20Jun 26, 2001

CO

NN

EC

TOR

WIR

ING

Fig. 7-1

LUN 7880.01-8 Electrohydraulic Actuator

27 V

A B V G D 1 2 3 4

WK 180 36WK 180 36VŠ 17 KPN 3G1 - 2x 35 V 400

VŠ 18 KPN 5G 1U2 - 2x 2RM14 KPN 4G1V1VŠ 17 KPN 2G1

VŠ 17 KPN 3G 1U2

WK 180 36

V112-2

WK 180 36

LUN 1540.02-8 Transmitter for Torquemeter Indicator

LUN 1476-8 Torque Limiter Pressure Switch

LUN 3280-8 Pressure Switch of the Autofeathering System

LUN 1333.12 Integrated Speed Transmitter (Gas Generator)

LUN 1333.12 Integrated Speed Transmitter (Propeller)

1.25K LUN 1469.32-8 Min Oil Pressure Switch

LUN 1581-8 Min Oil Quantity Signaller

Switch on theLUN 7816-8

limiter system

LUN 1348-8 Propeller Speed Indicator

6 7 11 12 13

6 7 11 12 13

M 10 M 14 M 174 M 172

V 8

V 16 2 RM 30 KPN 32 Š1V1 - 2 RM 30 BPN 32 G1V1

Optional

M 310 M 36 M 176 M 18

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

1

2

15

A B V

2

A B

774-S40D4VŠT 18 KPN 5G1U2

A B

„BETA“

limiter system

limiter system

recorder

manual and autofeathering system limiter system

1615

15

V 110-1

2524

A B V G D

M 34

1 2 3 4 5

LUN 1347-8 Gas Generator Speed Indicator

LUN 1539.02-8 Torque Indicator

1 2 3 4 5

4 3 5 8 9

autofeathering system

limiter system

10 19 18 32

10 19 18 32

27 31 26 28 23

27 31 26 28 234 3 5 8 9

5 V

lighting

lighting

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

red

5 V

WK 180 36

1 2 3 4 5 6 7

21 20 22

21 20 22 17

M 312

17

Electrohydraulic Transducer on the LUN 6590.05-8

16

A B V G D

24 25

V 112-2 V 110-3

1430

27 V 27 V

1430 WA

LTER a.s.

INSTA

LLATIO

N M

AN

UA

L M

AN

UA

L PAR

T No. 0982502

Section 7P

age 7-21/7-22Jun 26, 2001

CO

NN

EC

TOR

WIR

ING

Fig. 7-2



Section 7 Page 7-23

May 11, 2006

BLOCK DIAGRAM - STARTING SYSTEM

Fig. 7-5

Motoring Run Push Button

Starting Push Button

Switch for Spark Plug Check

LUN 2167.03-8 Regulator

Starter/Generator

LUN 2601.01-8 Time Relay of the Starting Panel

Spark Plug

Spark Plug

LUN 2201.03-8 (2 pcs) or

UNISON 9049765-1 Ignition Set

LUN 3191-8 Interrupter

Torch Ignitor Fuel Valve on the LUN 6290.04-8

Mains 27 V

to the Regulator of the 2nd Engine

to Storage Battery +

F = 35 mm2

F = 1.5 mm2

F = 35 mm2




WIRING DIAGRAM - ITT MEASUREMENT

Fig. 7-4

ITT Indicator 1 2

lighting

5 V

774.24-8

LUN 5260.04-8

IELU

CH

A CH

CH

+ - +-

LUN 1377-8

Interturbine Temperature Transmitter (9 pcs)

+

-

+ -

AA

A

CH



Section 7 Page 7-25

Jun 26, 2001

BLOCK DIAGRAM - STARTING SYSTEM

Fig. 7-5

Motoring Run Push Button

Starting Push Button

Switch for Spark Plug Check

LUN 2167.03-8 Regulator

LUN 2132.02-8 Starter/Generator

LUN 2601.01-8 Time Relay of the Starting Panel

N 25F-3 or

CHAMPION Spark Plug

N 25F-3 or

CHAMPION Spark Plug

LUN 2201.03-8 (2 pcs) or

UNISON 9049765-1 Ignition Set

LUN 3191-8 Interrupter

Torch Ignitor Fuel Valve on the LUN 6290.04-8

Mains 27 V

to the Regulator of the 2nd Engine

to Storage Battery +

F = 35 mm2

F = 1.5 mm2

F = 35 mm2




BLOCK DIAGRAM OF THE SYSTEM OF LIMITERS

Fig. 7-6

LUN 5223 Generator Speed Derivative Element

LUN 5260.04-8Integrated electronic limiter unit

(IELU)

TEST SWITCH

PUSH BUTTON

LUN 1333.12-8 Integrated Speed Transmitter (Propeller)

LUN 1333.12-8 Integrated Speed Transmitter (Gas Generator)

LUN 1476-8 Pressure Switch of the Torque Limiter

LUN 6590.05-8 Fuel Control Unit

A-037 Radioaltimeter

LUN 1377-8 ITT Transmitter

LUN 3280-8 Pressure Switch for Automatic Propeller Feathering

LUN 7816-8 Propeller Speed Governor

„ALERT IELU“

Cockpit

Push Button IELU

Switch in the System of Automatic

Feathering of the Second Engine Switch of Feeding

Pilot Lamps

„IELU“

„PARAMETER EXCEEDING“

P-5(7) Temperature Transmitter

Switch on the Undercarriage

LUN 2601.01-8Time Relay



Section 8 Page 8-1

Jun 26, 2001

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).




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 carrying-out 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 max. 130 °C (266 °F)

- 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.



Section 8 Page 8-3

Nov 15, 2004

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.






Section 9 Page 9-1

Nov 15, 2004

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.


- 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.


- 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)





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.



Section 10 Page 10-1

Jun 26, 2001

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.







Jun 26, 2001

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.







Nov 15, 2004

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 NOTE 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 on special order

Ignition unit LUN 2201.03-8 2 or UNISON 90 49 765-1 1

Propeller speed governor LUN 7816-8 1 Min. oil quantity signaller LUN 1581-8 1 on special order Min. oil pressure switch 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 on special order 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 * Part of the airframe installation




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.




Nov 15, 2004

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 * or 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 LUN 1370.02-8 1 on special order Three pointer indicator of LUN 1538.01-8 1 on special order

- oil temperature

- oil pressure

- fuel pressure (at fuel nozzle inlet)

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 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.




Cockpit signallers

Indication Description

1. Min. oil pressure - „MIN OIL PRESSURE“

2. Chips in oil - reduction gearbox

+ accessory gearbox - „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. Switching on of microswitch on the ECL when autofeathering system is on - „AUTOFEATHER“

9. 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.




Jun 26, 2001

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.



Section 12 Page 12-6 May 11, 2006

Instruments mounted on the accessory gearbox

100 % nG = 36,660 r.p.m.

Item

Instrument

Sense of drive

rotation

Speed ratio

Max. Torque

Nm (lb.in)

Overhang moment

Nm (lb.in)

Starter 11.2 (100) 21 1 Starter/generator D.C. generator CW 0.2898 9.2 (82) (186)

2 Spare drive (hydraulic pump drive) CCW 0.1974 5.8 (51)

40 (353)

3 Fuel Pump CW

0.1196

4.5 (40)

This instrument must be used

4 Fuel Control Unit CCW

0.1223

1.1 (10)

This instrument must be used

5 Integrated Gas Generator Speed Transmitter CW 0.1145 0.5 (4,5)

4 (35.4)

6a Manual turning by CCW 0.1145 11.5 (102) 0

6b Alternator (manual turning by) CCW 0.2763 5 (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)




Jun 26, 2001

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 Accuracy

Gas Generator Speed (nG) 0 to 110 % 0 to 60 %

60 to 110 % ±1.5 %

±0.5 %

Propeller Speed (nV) 0 to 2500 rpm 0 to 1500 rpm

1500 to 2500 rpm ±1.5 %

±0.5 %

Interturbine Temperature

(ITT)

0 to 900 oC 600 to 680 oC

680 to 725 oC

725 to 900 oC

±15 oC

±10 oC

±15 oC

Oil Pressure (pOIL) 0 to 0.4 MPa 0.1 to 0.3 MPa

otherwise ±0.017 MPa

±0.024 MPa

Oil Temperature (tOIL) -30 to 150 oC 0 to 90 oC

90 to 120 oC ±4 oC

±7 oC

Torque (TQ) 0 to 120 % 80 to 109 %

otherwise ±2 %

±3.5 %




Jun 26, 2001

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

l/min (US gal/min)

Inlet spray ring pressure

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.







Jun 26, 2001

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 kg 183 lb

Polar mass moment of inertia 10.1 kgm2 34,512 lb in2

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

5.95 kgm2 20,300 lb in2

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.




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.




Jun 26, 2001

0

5

10

15

20

25

40 50 60 70 80 90 100

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

Pro

pelle

r Mas

s P

olar

Mom

ent o

f Ine

rtia

J PV

RT

[kgm

2 ]

LIMITS OF PROPELLER MASS CHARACTERICS

Fig. 14-1







Jun 26, 2001

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




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).




Jun 26, 2001

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 M601-

77.6 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)




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

Appendix 1

Jun 26, 2001

ENGINE INSTALLATION DRAWING

Sheet 1


Appendix 2

Jun 26, 2001


Sheet 2


Appendix 3

Jun 26, 2001


Sheet 3


Appendix 4

Jun 26, 2001


SHEET 4


Appendix 5

Jun 26, 2001


Sheet 5


Appendix 6

Jun 26, 2001


Sheet 6


Appendix 7

May 11, 2006

62 (61.97)*

145

(142

.7) *

88 (88.9) *

(294)*

300

COOLING AIR INLET

(5.6 +1.5) *

(1.3 +1.5)*

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


Sheet 7


Appendix 8

Jun 26, 2001


Sheet 8

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Installation Manual Walter m601e, m601e-21[1]

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