DGLR-MTR-2000.PDF

THE SUCCESSFUL DEVELOPMENT HISTORY OF THE MTR390 TURBOSHAFT 

ENGINE FROM THE DEFINITION OF THE DESIGN TO THE SERIES PRODUCTION 

P. Schinzl 1) , F. Malzacher 1) , P. Moncoutie 2) , B. Brereton 3) , 

G. Schuberth 4) , M. Lauvaux 4) 

1) DaimlerChrysler Aerospace, MTU München 

2) Turbomeca 

3) Rolls Royce 

4) MTR GmbH 

1. ABSTRACT 

This paper presents the MTR390 turboshaft engine 

used to power the Tiger attack helicopter by an emphasis 

of the development history. The MTR organisation 

who produce the engine, and the workshare of the 

three partner companies, MTU, TM and RR is described. 

Modularity, maintenance and the growth 

potential of the engine are covered, as well as engine 

design and test experience. A summary of the programme 

status is presented with an overview of other 

potential applications for the engine. 

2. INTRODUCTION 

The MTR390 (Fig. 1) is a new turboshaft engine in the 

1000kW range being jointly developed by three of the 

largest European engine companies: MTU in Germany, 

TURBOMECA in France, and ROLLS-ROYCE 

in Great Britain. 

Fig 1: General Arrangement 

The three companies have been working together for 

more than 30 years on several programmes like 

Adour, Larzac, RB199, EJ200, RTM322, .... The declared 

goal of the MTR partner companies is to create 

an engine that will not only suit the TIGER/GERFAUT 

programme but will also compete in the military and 

civil market for the next 20 to 30 years. 

The development and qualification programmes have 

been completed together with the AMT programme 

completion in 1997 and led to the qualification of the 

production engine in 1999. As a result of the successful 

completion of all certification requirements the 

MTR390 has obtained its Military and Civil Type Certificates. 

The production contract in has been signed 

preparing the way for production engine delivery in 

2001.

3. ENGINE PARTNERSHIP 

MTU TURBOMECA ROLLS-ROYCE GMBH (MTR) 

GmbH is a joint company established under German 

law in Munich in 1989 by MTU Motoren- und Turbinen- 

Union München GmbH, Turbomeca S.A. and Rolls- 

Royce plc. 

The purpose of the company is 

• to act as contractual party vis-à-vis the authorised 

governmental authorities and other potential customers 

for the MTR390 engine 

• to responsibly manage and coordinate among the 

partner companies the development, airworthiness, 

production preparation, production, marketing, 

sales, licensing and support of the MTR390 

engine. 

MTU: 

MTU Munich is part of DaimlerChrysler AG and responsible 

for all aero engine business as the leading 

German engine company. 

MTU is in charge of the design and manufacture of the 

Combustion Chamber, the High Pressure Turbine, the 

interduct and rear bearing chamber and the oil cooler 

Turbomeca (TM): 

Turbomeca, as part of the French SNECMA Aero 

Engines Group, is the world leader in turboshaft engines 

for helicopters. Its engines cover the largest 

power range from 450 to 3.000 shp and equip helicopters 

made by the world’s leading manufacturers. 

TM is in charge of the design and manufacture of the 

compressor (LP - HP), the gearbox (reduction and 

accessory drive), most of the accessories and the 

FADEC 

Rolls-Royce (RR): 

Rolls-Royce is a world-leading power systems business 

meeting the needs of customers in propulsion, 

electrical power and materials handling around the 

world. 

RR is in charge of the design and manufacture of the 

power turbine, the stub shaft and the power turbine 

shaft 

4. MAIN ENGINE FEATURES 

The MTR architecture is the result of an optimisation 

study, commonly carried out by the three partner 

companies. 

High efficiency 

centrifugal compressor High efficiency 

power turbine 

Main reduction and 

accessory 

gearbox 

Integral oil system 

Fig. 2: Engine architecture 

Annular reverse-flow 

machined ring combustor 

for increased durability 

5. DEVELOPMENT HISTORY 

Advanced air-cooled 

gas generator turbine 

for lower cost and 

increased reliability 

The development of the MTR390 engine started officially 

on 1 st January 1988. Within two years, the 

technical requirements were specified and the main 

development contract was approved by the end 1989 

followed by the first engine run in December 1989. 

The engine development phase was divided into three 

major steps: 

- Qualification A: Preliminary flight release 

- Qualification B: Military Type Certificate (Musterzulassung) 

- Qualification C: Release of the production standard 

The first development phase – Qualification A – was 

required for the clearance of the MTR390 engine for 

the FTB (flying test bed) and prototype testing. In 

accordance with an agreed qualification procedure, 

engine and accessories tests i.e. 60 hrs endurance 

run, vibration investigation, engine overspeed tests, 

software integration, power turbine overspeed test (cut 

shaft), oil and fuel system performance tests, oil flow 

interruption test in order to clear negative g flights, 

equipment vibration and fire tests etc., were successfully 

performed. The completion of these tests with the 

so-called engine standard 1A ended in the clearance 

for FTB tests in the ‘Panther’ helicopter and flight 

tests in the ‘Tiger’ prototype helicopter. 

The first FTB flight was successfully performed with 

two MTR390 engines on 14 th February 1991 while the 

first flight on the ‘Tiger’ prototype helicopter PT1 took 

place with prototype engines on the 25th April 1991. 

The main subject of Qualification B (Military Type 

Certificate) and C was to satisfy the Airworthiness

equirements and to release the production engine 

standard. 

With a ‘new’ defined engine standard 2A, additional 

engine and rig qualification tests were carried out and 

successfully completed, i.e. 

• rain ingestion test 

• snow ingestion test (simulated with crushed ice) 

• 150 hrs type endurance run 

• overspeed tests 

• vibration surveys 

• oil/fuel clearance runs within several 150 hrs endurance 

tests 

• 30 seconds OEI demonstration (8 shots of 30’’ 

each) 

• inclination tests 

• accelerated mission testing (2 engines, each 

1.200 hours running) representing an ageing 

equivalent to 4.300 Tiger mission hours; 

• sand ingestion tests (with and without sand filter) 

• corrosion test 

• exhaustive aircraft - engine interface qualification 

(control system, …) 

One important step in achieving this qualification 

milestone was the test campaign performed on the 

altitude test facility (ATF) at Centre d’Essais des 

Propulseurs (CEPr) in France. Within several test 

phases the following major qualification justification 

were performed: 

• Demonstration of the starting and restarting envelope 

• Demonstration of the auto-relight function within 

the operating envelope 

• Demonstration of the specified engine performance 

within the operating envelope 

• Clearance of engine fuels and oil brands for operation 

within the specified envelope 

• Icing test with a representative helicopter air intake. 

Based on the engine qualification evidence (engine 

tests, analyses, FMECAs, etc.) the engine standard 

2A was granted the military Type Certificate by the 

German Airworthiness Authorities on the 09.05.1996. 

Subsequent to the issue of the Type Certificate, the 

development of the engine and software concentrated 

on performance improvement, optimisation of acceleration 

and deceleration software laws and consideration 

of maintenance aspects. In order to separate the 

Qualification B and C standards, an interim standard 

(2B) was introduced with improved performance features 

and improved ease of maintenance. This led to 

the introduction of some modifications. Although most 

of the qualification tests were performed with the engine 

standard 2A, all results achieved were transferred 

to the final qualification standard which will be declared 

as engine build standard 2C at the end of the 

engine development phase. The successful comple- 

tion of qualification of the MTR390 engine closed with 

the handing over of the Qualification C Certificate by 

the military customer. 

In parallel to the military certification activities, MTR 

started to gain the civil type certificate. From the very 

beginning of the development programme, a civil application 

of the MTR390 was intended and therefore 

the MTR390 engine specification is based on two 

main airworthiness regulations, the British Civil Airworthiness 

Requirements (BCAR) Section C, issue 

13, which is equivalent to JAR-E change 6 and the 

Military Specification MIL-E-8593A. 

In November 1992, MTR contacted the LBA in order to 

discuss possible ways of obtaining the Civil Type 

Certificate for the MTR390 turboshaft engine. As result 

of these discussions MTR decided in November 

1994 to apply for a German Type Certification in the 

frame of a Simplified Type Investigation according to 

§4 LuftGerPO. 

The reason for selecting this way was that the company 

preferred to use the advantage of the German 

Procedures to avoid additional costs potentially resulting 

in a repetition of the majority of certification activities 

which would have resulted when using the Joint 

European Procedures. 

This route was possible because the Military Certification 

Basis was JAR-E, change 6. Therefore MTR 

complies with the German regulations that require the 

same level of safety in the case that the LBA validates 

a German Military Type Certificate. 

At the beginning of June 1997 the final civil documentation 

was submitted and accepted by the German 

Civil Airworthiness Authorities supported by the German 

Military Authorities. Based on the evidence given 

to both Airworthiness Authorities, the Civil Type Certificate 

was granted by LBA on 19 June 1997. 

6. ENGINE DETAILS 

6.1 Compressor 

The compressor is a two stage centrifugal system 

providing a pressure ratio of 14 with high efficiency. 

The important surge margin enabled to delete the 

bleed valve after extensive rig and engine testing. 

The centrifugal compressor is particularly well 

adapted for small turbo-shaft engines because it reduces 

the engine compartment length for installation 

into the aircraft and also vulnerability to battle damage 

for military applications. 

Its simple design with a minimum of parts (2 stages, 

no bleed valve, no IGVs ) is an advantage for a good 

reliability.

It demonstrated particularly good tolerance to FOD 

during the development and qualification tests (sand, 

ice, water, …). 

Both stages can be easily inspected with boroscopes 

in order to monitor the mechanical 

condition. 

6.2 Combustor 

A reverse-flow annular combustor was chosen for its 

advantages regarding the whole engine architecture. 

Best adapted to the centrifugal compressor design, it 

minimises engine length, provides good accessibility 

to fuel nozzles and allows a short and rigid generator 

rotor. 

The combustor [1] can be shortened only provided 

rapid fuel conversion is achieved in its primary combustion 

zone. Primary zone conversion is controlled 

by the fuel mixing and vaporisation subprocesses. 

These processes can be precipitated by achieving a 

preferably small size of fuel droplet, a high relative 

velocity between the drops and the gas, a preferably 

homogeneous fuel distribution in the primary zone, 

and a high turbulence level conducive to efficient mixing. 

Therefore, the fuel conditioning system proved to 

be a key criterion. 

In order to improve the fuel conditioning system MTU 

developed a new type air blast atomiser which is illustrated 

in Fig. 3. 

Fig. 3 Air blast atomiser 

According to this design the fuel is sprayed onto a 

venturi insert by means of a pressure atomiser. The 

two counter-rotating primary zone vortices concentrates 

the burning process in the centre of the primary 

zone. The secondary airstream forms a layer on the 

flame tube wall. This system prevents contact of the 

burning gases with the flame tube walls. Fig. 4 shows 

the flame generated by this type of atomiser.

Fig. 4 Flame structure at atmospheric pressure 

The high homogeneity inside the primary zone leads 

to homogeneous temperature profiles at the combustor 

exit and on the flame tube (s. Fig 62.3). So, the 

achieved OTDF and RTDF values leads also to a 

minimisation of the cooling air for the compressor and 

power turbine: 

OTDF < 30% 

RTDF < 14% 

Fig. 5: Thermo colours pattern after the rig test 

Furthermore, the described design of the fuel conditioning 

system leads the a very stable combustion 

(LBO > 300 kg-air / kg-fuel) which is needed in the 

case of high deceleration of the engine. 

6.3 Compressor Turbine 

The increase of the specific power as well as the reduction 

of the engine weight can be achieved by reducing 

the numbers of stages in the turbomachinery 

components. Therefore, the change from the traditional 

two-stage turbine to a single stage design was 

chosen at the very beginning of the development. The 

technology of a transonic single stage turbine [2] with 

a pressure ratio of 3.6 for small turboshaft engine 

applications (Fig. 6 was developed within a demonstrator 

programme which was partly sponsored by the 

Ministry of Technology of the Federal Republic of 

Germany. 

Fig. 6 Single stage transonic high pressure turbine 

The key aspects which are mainly to be handled are 

on one hand the high aerodynamic loading with su-

personic Exit Mach Numbers for the blading. This 

leads to compression shocks including additional 

shock-boundary layer interactions with the combined 

complexity and sensitivity to geometrical deviations. 

Much attention was given therefore to the aerodynamic 

design. On the other hand the minimisation of 

necessary airfoil cooling flows and the optimisation of 

the tip-clearance are important to get the best result 

with respect to the engine operation line. 

The compressor turbine design including the interturbine 

duct was also affected by the mechanical 

boundary conditions with very high temperature and 

stress levels for life limiting parts. This shows the 

excellent interaction between the necessary technical 

disciplines. 

Fig. 7 shows the loss contributors in this class of 

engine size in comparison with a large engine. It is 

evident that tip clearance and cooling losses are of 

the same order as aerodynamic losses and have to 

be taken careful into account. 

Fig. 7: Thermodynamic losses in the turbine module 

In Fig. 8 published results of several tip clearance 

investigations are plotted together with rig and demonstrator 

engine data from MTU. 

Fig. 8 Influence of tip clearance variations on stage efficiency 

As blade height is only 17mm, a tip clearance of 

0.1mm results in an efficiency change of 1 point. 

Fig. 9 highlights the impact of tip clearance on the 

radial efficiency distribution for 1, 2 and 3 % relative 

radial gap(s/h). The efficiency is affected down to 50% 

blade height. 

Fig. 9: Influence of tip clearance variation on the radial 

efficiency distribution 

Therefore, an efficient system was chosen with liner 

segments hung into a casing of low thermal expansion 

and controlled by forced cooling. The short and 

rigid gas generator rotor contributes to the efficient tipclearance 

control. 

In order to minimise the cooling air both nozzle and 

vane are of single crystal material which permits high 

temperatures. The blade cooling configuration was 

developed to minimise the amount of air and the detrimental 

effects of secondary flows re-entering the gas 

path. 

6.4 Power Turbine 

The power turbine is a two-stage uncooled design. 

The philosophy behind this module was to ensure 

ease of maintainability and repair whilst achieving a 

low cost reduced weight yet reliable design. Several 

features were key in this achievement. 

The power turbine is easily replaced in field through 

the removal of a series of retaining bolts on the power 

turbine case together with the curvic bolts. The module 

then simply slides rearwards and clear from the 

engine. The power turbine bearings and air/oil system 

are retained within the core engine module. The power 

turbine can be replaced without the need to re-test the 

engine. 

The curvic coupling, mentioned previously, transmits 

the torque to the engine reduction gearbox. This curvic 

coupling allows accurate re-assembly of the power 

turbine, reducing the susceptibility of the module high 

speed balance to wear at this interface. 

To ensure Foreign Object Damage (FOD) tolerance a 

single crystal material and equi-axed material was

chosen for the first and second stages respectively. 

The materials chosen offer impact resistance at a 

relatively low density. This is an important consideration 

with the cantilevered power turbine design when 

assessing the highly unlikely event of a blade loss 

and the associated out-of-balance moment that is 

generated. The power turbine foreign object damage 

(FOD) tolerance was optimised and demonstrated 

during the development programme. 

The power turbine rotates in the reverse direction to 

the engine core. This reverse rotation minimises the 

'turning' required by the PT1 NGV to ensure acceptable 

entry conditions to the first stage of blade. This 

contributes to the high overall efficiency of the power 

turbine. 

The efficiency is further improved by the absence of 

an exhaust diffuser and its associated discharge 

losses. Overall module weight is also reduced. 

7. FADEC CONCEPT 

The engine control system is essentially based on a 

FADEC (full authority digital engine control) developed 

from the background acquired over more than 10 

years flight experience on other turboshaft engines 

applications. This concept allows a great simplification 

in the hydromechanical units most of which are 

grouped in the accessory gearbox module. 

It provides carefree engine operation, control and 

monitoring. 

It is essentially a single channel with electric manual 

back up and analogue overspeed protection. 

The control function provides basically automatic start 

sequence, rotor run up, and then maintains the Rotor 

at the nominal value. 

In flight the engine is protected against limit exceedance, 

surge and flame out for a minimum of pilot 

workload. 

Many other functions are also available like training 

mode which simulate an engine failure for pilot training 

and automatic relight in flight in case of flame out. 

A lot of work has been done including several altitude 

tests for optimisation of engine transient accelerations 

and decelerations. 

The result of this optimisation is one of the key elements 

enabling TIGER aircraft impressive manoeuvrability. 

The FADEC also provides maintenance help functions: 

• Built In Test: monitors the defect of the Control 

Unit, all the accessories and aircraft interfaces 

electrically linked to the Control Unit. 

After a defect is detected the pilot is informed of 

the consequences if any on the engine operation 

and complementary localisation information helps 

the maintenance personal to repair after the flight. 

• Engine Performance Check: monitors that the 

minimum specified power is available. 

• Limit Exceedance Monitoring: records potential 

limit exceedance duration and peak values of 

main parameters (torque, turbine temperature, 

gas generator and power turbine speeds) in order 

to define the maintenance actions. 

• Life Usage Monitoring: computes the life consumption 

of the life limited parts (impellers, disks, 

…) 

8. MODULARITY AND MAINTENANCE 

The engine consists of four separate, interchangeable 

modules, including the FADEC can be changed without 

the need to retest the engines, nor are any special 

checks or adjustments required on module 

change. 

Maintenance time and costs are reduced by „On 

Condition monitoring“ carried out by the FADEC, with 

its built in diagnostic aids, pre-clogging and by-pass 

indicators for the fuel and oil filters and chip detectors 

in the oil system to monitor bearing condition. The 

built in test failure detection system also informs the 

maintenance personal of the location of faults, to 

minimise maintenance time. 

A minimised toolkit is required for first line maintenance, 

consisting of just a few hand tools. All line 

replaceable units (LRU) can be removed by one operator, 

in an average time of 15 minutes, and there is 

no wire locking. 

9. TEST EXPERIENCE 

During the programme, a total 17.700 hours of engine 

running have been completed. Development bench 

running of 10200 hours has been accumulated, including 

1500 hours of endurance testing and 1.800 hours 

of AMT. Flight engines have logged up a total 6500 

hours, 5800 hours in the Tiger, and 700 hours in the 

Panther FTB. 

The engines in the flight programme have proved to be 

very reliable. In the 6500 flight hours there has been 1 

unplanned „in flight“ shutdown, caused by a design 

problem which has now been corrected and 1 basic 

unplanned engine removal. Considering this is during 

the development flight programme, it is an impressive

esult. The engine was designed from the outset to be 

a simple and reliable design, with a low parts count to 

lead to a reliable engine. 

10. GROWTH POTENTIAL 

The MTR390 features a power growth capability of 

50%: 

• a first growth step of 10% could be achieved by 

increasing the turbine entry temperature; up to 

level already tested on other programme; 

• a second growth step of 20% could be achieved 

by modifying the compressor 1 st stage such as to 

increase the air mass flow within the same size; 

• additional steps could be envisaged later on by 

means of further temperature increase and / or 

deeper compressor modifications. 

11. SUMMARY 

The MTR390 has now finished a successful development 

test and flight programme. This leads to production 

of the engine for the Tiger attack helicopter due to 

begin in 2001. The engine has also been designed 

and developed with other military and civil applications 

in mind, and to this end, has obtained both Military 

and Civil Type Certificates. 

12. LITERATURE 

[1] F. Joos, B. Simon, B. Glaeser, S. Donnerhack 

Combustor Development for Advanced Helicopter 

Engines 

ISABE 1991 

[2] H.-J. Dietrichs, F. Malzacher, K. Broichhausen 

Aerodynamic Development of an HP-Turbine for Advanced 

Turboshaft Engines

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