THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS
92-GT-250
345 E. 47 St., New York, N.Y.10017
ES
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Discussion is printed only If the paper Is published in an ASME Journal. Papers are available
from ASME for fifteen months after the meeting.
Printed n USA.
Copyright © 1992 by ASME
MTR390—Engine for the Future
A. SPIRKL
MTU
Munich, Germany
J. S. DUCOS
Turbomeca
Bordes, France
R. THORN
Rolls-Royce
Bristol, UK
ABSTRACT
The design of the MTR390, a turboshaft engine
in the 1000 kw range, is based on proven technology
and components developed well in advance of the actual
programme start. The engine features a two stage
centrifugal compressor, an annular reverse flow ring
combustor, a single stage air cooled gas generator
turbine, a two stage free power turbine, an integrated
reduction and accessory gearbox with integral oil
system. A full authority digital engine control provides the engine and system monitoring functions.
The development programme, which also includes
a comprehensive integrated logistics support programme,
is running according to schedule, more than 1400
engine hours have been logged including 200 flight
hours with both a Flying Test Bed and the TIGER prototype helicopter.
INTRODUCTION
The MTR390 is a new turboshaft engine in the
1000 kw range being jointly developed by three of
the largest European aero engine companies: MTU,
TURBOMECA and ROLLS-ROYCE for the Eurocopter "TIGER"
anti-tank helicopter (Fig. 1).
M'U, TURBOMECA and ROLLS-ROYCE have set up a
joint company MTU TURBOMECA ROLLS-ROYCE GmbH (MTR)
to coordinate the development, production, marketing,
sale and support of the MTR390 engine, and to act
as a contractor for the German and French Governments
and other customers.
High efficiency
centrifugal compressor
High efficiency
power turbine
rs^f,
Ma in reduction
and accessory gearbox
rs ._
I
I
,(
^"^
/
Integral oil system
for good balllstic
Annular reverse-flow
machined ring combustor
for Increased durability
I
Advanced
air-cooled Vas generator
turbine for lower cost
and Increased reliability
tole arise
Fig. 1: Engine Design
The engine has therefore to meet requirements
based on both civil and military standards as well
as to reach competitive targets in terms of performance, installation, maintainability, durability and
cost of operation.
ENGINE DESCRIPTION
COMPRESSOR (Fig. 2)
This program is far from being the first cooperation between the three companies which have been
working together for more than 25 years on several
programs among which are; Adour, Larzac, RTM322,
EJ200 ...
The compressor is a two stage centrifugal system,
the present design of which results from a development
initiated in 1982. Results of preliminary studies
have shown that compared with the axial-centrifugal
version envisaged at the outset of the programme,
there are significant performance advantages (efficiency (+2%) - flow - pressure ratio) obtained
at a second stage circumferential speed lower than
that of the centrifugal compressor of the axial-centrifugal version.
It is a declared goal of the companies to develop
and produce an engine, which is capable of meeting all challenges of the market in this power class
in the next 20 to 30 years.
The tests carried out on a reduced scale test
rig have confirmed these predictions by giving results
equal or better than target, for example (at nominal
speed):
Presented at the International Gas Turbine and Aeroengine Congress and Exposition
Cologne, Germany June 1-4, 1992
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Target
Obtained
2,5
14
0,8
2,25
14,4
0,814
Flow
Pressure ratio
Efficiency (isentropic)
in this power class to a single stage design (2).
Development of this transonic turbine with cooled
vanes and blades was started in 1982.
In the aerodynamic design of the gas generator
turbine, which included the interturbine duct, attention has been paid special to the reduction of secondary flow losses. The flow path design in the stator
area - optimised by the application of a reverse
flow annular combustion chamber - results in a low
radial pressure gradient at the stator outlet, minimizing boundary layer flows in the radial direction.
Reduction of work extraction close to the rotor blade
tip reduces the sensitivity of turbine efficiency
to blade tip clearance. The potential for a reduction
of tip clearance sensitivity, however, turned out
to be limited, as explained in (3). So an effective
tip clearance control was one of the major development
objectives for this component.
Turbine aerodynamics were optimized on the basis
of an extensive rig programme. Four different aero
design standards were investigated in 20 builds of
a cold flow aero rig, including studies of cooling
air and blade tip clearance effects on turbine performance.
Fig. 2: Dual Stage Compressor
of:
The structure of the compressor unit consists
POWER TURBINE
- an air intake in cast aluminium light alloy, including an integrated washing system
The power turbine is a two-stage uncooled design,
the aerodynamics of which have been optimised to
give a "flat" efficiency characteristic from cruise
power upwards. Use of a contra-rotating power turbine
relative to the gas-generator takes advantage of
the exit whirl gas angle from the HP turbine stage,
thereby reducing the amount of turning required in
the first stage power turbine nozzle, and saving
an estimated 1% in efficiency.
- an external casing in welded Inconel
- internal casings in cast steel
- stators in steel and titanium, machined from castings and assembled by brazing
- two forged rotors in titanium, joined by a curviccoupling.
Mechanical design features a single crystal
stage 1 blade, the second stage being conventionally
cast. Both blades are shrouded.
The choice of this design structure with thin
walled casings and static parts, together with the
extensive use of titanium, leads to an optimised
balance between dimensions and weight.
The stage 1 nozzle is segmented, each segment
consisting of three vanes. This configuration was
found to present the best tradeoff between life,
cost and performance. The stage 2 nozzle is semisegmented, with a one piece outer ring and an inner
ring cut to form segments of three vanes each.
COMBUSTOR
A reverse flow annular combustion chamber was
chosen for its advantages regarding the whole engine
architecture. Best adapted to the radial compressor
design, it minimizes engine length, provides good
accessibility to fuel nozzles and allows a close
coupling between the compressor and the HP turbine,
thereby reducing the gas generator length.
The power turbine shaft drives forward through
the centre of the gas generator to the reduction
gearbox. The design uses a separate splined stubshaft which allows the power turbine module to be
removed without exposing the rear bearing chamber
to contamination.
A fuel injection and flame stabilisation system
with air blast atomizers and two counterrotating
primary zone vortices concentrates the burning process
in the centre of the primary zone. Thus the system
prevents contact of the burning gases with the flametube walls (which are proportionately larger for
reverse-flow combustor design).
GEARBOX
The complete gear system has been grouped in
a single module located at the front of the engine.
It includes two operational units:
- In the lower part of the module, the reduction
gear, which reduces the rotational speed of the
power turbine (27000 rpm at take off rating) to
the output speed specified by the helicopter manufacturer (8000 rpm). The reduction system comprises
two intermediate trains through which the load
is distributed and balanced by a hydraulic system,
which also provides torque measurement.
GAS GENERATOR TURBINE (1)
Marked progress in turbine aerodynamics, based
on refined analytical design methods and test results,
as well as the availability of improved materials
- powder metal (PM), directionally solidified (DS)
and single crystal (SC) materials - paved the way
for a change from the traditional two-stage turbine
2
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This concept allows a great simplification in
the hydromechanical units limited to the following
elements, most of which are grouped in the accessory
gearbox module:
- In the upper part of the module, the accessory
gearbox which provides the support and drive for
engine accessories:
•
•
•
•
LP and HP fuel pumps
Oil pumps
Oil cooling system
Electrical starter.
- LP and HP fuel pumps
- Filter
- Metering valve
The reduction gear, the accessory gear box and
the engine oil tank are housed in common aluminium
casings, the unit being a separable module of low
size and weight, capable within its existing envelope
of a power increase in the region of 20% without
major redesign.
- Start valve, stop and those for purging of injectors.
serial data link
Helicopter ^
The gearbox casings are designed such a way
that they can accomodate a reduction ratio increase
for a 6000 rpm output shaft speed with minor gear
modification.
Pilot orders
_
-
Helicopter outputs
DC from helicopter
Ergfne
control
system
Engine
monitoring
system
Overapeed
prMectbn
I Helicopter logical
Ilnputsoulputs
DC from helicopter
Sensors B
(Main
control)
Engine logical
outputs
A direct drive layout at 27000 rpm output shaft
speed has also been investigated and only requires
a redesign of the gearbox module; provision has been
made in the engine design so that this arrangement
is feasible (Fig. 3).
Engine
ECOS
Sensors A
!
Serial date
link
EMOS
j
AC
Alternator
01 LP pump
Actuator Order
Y
Metering
valve
Overspeed
stop
order
Shot-oft valve
Back up unit
(auxiliary control)
Adaptable drives and
intake configuration
MTR390
Low speed drive
I
8000 rpm and
6000 rpm Fig. 4: Control and monitoring System
Common core
-
a
The electronic unit is mounted in a separate
box which can be installed either directly on the
engine, or in the helicopter equipment compartment
(as is the case of the HAP, PAH2/HAC applications).
^
n
'"
1i ) - L
h
The essential principles which govern the system
design are:
MTR390T
High speed drive
27000rpm
- Full authority, ensuring the control of all ratings
and all powers, from stop the to maximum authorized
rating.
-^. S. ---^ ` s
- Detection and correction of control system failures,
to minimize their effect on mission reliability
and the work load of the pilot.
Fig. 3: MTR390 Derivatives
(1TT CvOmwM
- Safety of operation resulting from the incorporation
of auxiliary emergency loops, with electronic or
manual control.
The complete integral oil system features (4):
- integral gearbox/oil tank casing providing mass
reduction, simple routing of oil piping and minimum
external pipes for enhanced survivability
- Autonomous electrical supply from the engine (except
for starting).
- Pilot assistance provided by:
- only 2 bearing chambers
the supply of parameters for engine surveillance
• Simulation of failure cases by means of a training function
- no in-field pressure setting
- oil sight glass on left and right side
- Maintenance aid
- gravity oil replenishment
- of the system itself, with the help of the failure self-detection and its modular breakdown
ENGINE CONTROL AND MONITORING SYSTEM (ECMS)
The engine governing system is essentially based
on a digital concept developed from the experience
acquired over some years on other turboshaft engine
applications.
- of the engine by:
• performance control
limit exceedance control
calculation of life utilisation (counting
of cycles)
3
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DESIGN FEATURES (5)
Ratings
The simple, robust design of the 2-stage centrifugal compressor makes the engine insensitive to erosion
and foreign object damage. A short engine is achieved
by the incorporation of an annular combustion chamber,
which with its air-blast fuel-injection system ensures
low-pollutant combustion. The gas generator turbine,
which features cooled single-crystal blades and powder
metal disk for long life, is of single-stage design
to reduce cost and engine weight. The speed of the
2-stage power turbine is matched to the 8000 rpm
requirement of the helicopter manufacturer via a
reduction gear box. The engine accessories are easily
accessible and can be speedly changed without the
need for special tooling.
Normal operation
Take-off (5 min)
Maximum continuous
SRAPT POWER (ISA 0/0)
OPERATION
- See Pig. 6
- Orovth potential
snort cerm0 $
long torn:
50 $
- Comtat mleeion vlth a high
number of accels and de0e15
em bests for 6000 h life end
high reliability
lb/shp/h
958
873
1285
1171
274
277
0.451
0.458
One engine inoperative
Super emergency (20s)
Super emergency (308)
Contingency (2.5 min)
Intermediate (30 min)
1160
1138
1027
958
1556
1526
1378
1285
—
—
—
—
274
0.451
Average new engine
SLS/ISA
Uninstalled
Output shaft speed 8000 rpm
Fig. 6: Engine Performance
Gas temperature rating and design concept are the basis for
considerable power growth potential.
--1. L7
- Acceleretlon 0- 95 $ T0, 1
3
- Manoeuvre loads
2,5 red/s
a 1 g
J
STEP 3
TIT
Further nch.
(tooled PT1)
a lal cantrilugel
replacement
^. I!
.20
m,
mm
^.:✓
cis
- Inetellatton attitude: a 15 0 roll
_J'I (i
0_^^^
E iNONE
^^
STEP 1
l
1
I
STEP2
Increee-E TIT
a lctcenhi110ga
replacement
IncreaveE TIT
("VP- T Inlertlucl)
- Self-contained oil system
-
g/kW IT
4u
- Front drive
0000 to 5320 RPM
shp
kW
U
INSTALLATION
- Output shaft
- speed
1
- dieppleos,ent rel.
to engine C 1125
Specific fuel consumption
Outer shaft power
The main design requirements are highlighted
in Fig. 5.
aPTl noIIb
BASIS
S hon term
tong term
TIT: Turtles Inlet Tempenure
PT1: Power Turbine Siegel
lot, fuel 000eusPtlon
150 xelght
lov production cost
Out, maintenance costs modular design,
good msintalnab:llty, high performance retention
Fig. 7: Growth Potential
The engine is supplied complete with all accessories and systems necessary for independent operation.
The integral oil system has been designed with special
emphasis on reduced vulnerability. Automatic engine
start, gas generator and power turbine speed, torque
equalization between the engines (twin-engined versions) and the observance of limits are all controlled
by a modern FADEC system based on experience gained
with other engines. An engine condition-monitoring
system supplies all information necessary for rapid
fault diagnosis and correction.
Fig. 5: Main Design Requirements
Since protection against the ingress of foreign
objects into the engine intake is not necessary for
all flights altitudes, an integral particle separator
(IPS) is not fitted. When needed, an aircraft air
filter can be installed providing a much more effective filtration than an IPS. A screen is provided
for trapping coarser particles, and if required,
a sand separator can be fitted by the helicopter
manufacturer. As experience in the recent Gulf war
showed, this is necessary for operations in a sandy
atmosphere, even with engines fitted with an IPS.
The helicopter's vulnerability is markedly reduced
by appropriate exhaust control and by the use of
an infrared suppressor (IRS).
MAINTAINABILITY (Fig. 8)
Maintenance actions will be required "on condition" only.
In addition to the maintenance aid system integrated in the control system electronic box, the
engine is equipped with devices for monitoring of
mechanical health:
The various engine ratings (Fig. 6) have been
set such that on the one hand the high emergency
power in the event of failure of one engine of the
helicopter, which operates predominantly at low altitudes, will not result in critical situations, and
on the other hand, a potential power increase of
up to 50% can be achieved (Fig. 7).
- indicators of oil and fuel filter pre-clogging
and blockage
- magnetic plugs for the detection of metal particles
at several points of the oil system
- ports for boroscope inspection
- vibration pick-ups (optional).
4
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From the very beginning of the design, maintainability studies have been carried out to facilitate
maintenance operations on the installed engine, and
in the shop. These studies have been conducted on
mock-ups and have been taken into account during
the engine definition and are being continued on
the helicopter prototype.
Today, Oct. 1991, - 14 bench engines, including
2 engines presently installed in the "Flying Test
Bed" (FTB) and 3 prototype engines (Fig. 10) delivered
to Eurocopter for the start of the "TIGER" development
have run a total of more than 1400 h, including approx.
200 hours in flights up to 4100 m. Parallel to the
engine tests a great variety of mechanical and aero/
thermodynamic component tests have been carried out.
All these actions will ensure a high availability
rate and a reduction in the user's operating costs.
cons
MODULE
- Minimum inspection requirements
Fully modular conception
- On condition monitoring - Minimum epecific tools
- Built-in Ooepreeeer vaah - Adjustment From module replacement
- Health monitoring system - No teat at user 1eve1
Some of the most important tests so far are
the endurance tests. These tests were carried out
at full rated temperatures, including all increments
for deterioration, instrumentation accuracy etc.,
all specified powers have been demonstrated including
20 second super emergency power as well as the specified uninstalled acceleration rates (
3s from 0
to 95% take-off power).
- No core-locking
Layouts after test showed the engine to be in
a perfectly good condition. Together with other specification compliance tests successfully carried out
with the same engines and other engines before, this
shows the quality of the engines to be fully compliant
with the requirements specified for Eurocopter's
flight test programme.
- By-pass iodination of Fool and oil filters
- Non handed engine
LIDS gEFca^&Bls UNIT
Adjustment free replacement
- Revered Sna0013at1on impoeeible
- stendard tools for r.pl.oement
- Change time einimioed
MAINTENANCE REQUIREMENTS MINIMIZED
- DIAGNOSTIC POSSIBILITIES MAXIMIZED
Fig. 8: Maintainability Concept
MODULARITY
The MTR390 is fully modular and comprises three
easily changeable modules:
gearbox, gas generator and power turbine.
The modules are interchangeable between engines
in the field and module fastenings have been kept
to a minimum to facilitate removal and replacement
in minimum time.
PROGRAMME STATUS
Fig. 9 shows the overall programme which was
officially launched in December 1987. The development/
certification programme is planned to run for 6000
hours with an additional 2400 hours accelerated mission
testing before entry into service. The first bench
engine ran 19th Dec. 1989, approx. 2 months ahead
of the target date set by the contracting authority.
90
91
92
93
95
96
97
98
99
The reason for this satisfactory situation after
roughly 2 years of engine development is first of
all that engine design was based on proven technology.
Secondly, in order to gain a maximum amount of experience right at the beginning of the engine development
programme, the first bench engines were equipped
with very comprehensive instrumentation including
more than 600 measurement points.
2000
The instrumentation has allowed the recording
and the analysis of a considerable amount of important
data on the engine behaviour, its overall performance
and the performance of major components and internal
systems. On the basis this information, design changes
could be defined immediately and incorporated in
the first flight engines and the endurance engine.
It should be noted that such changes were minimized
thanks to the excellent overall test behaviour of
the initial engines.
Oueliticetion test
First flight
First run
94
Fig. 10: Flight Engine
Design &
Fdevelopment
Production
In service: HAP PAH-2 HAC
Support
I
90
I
I
I
91
92
93
I I I I 1
94
95
96
97
98
99
More than 25 years of experience in working
together has allowed the three partner companies
to set up an effective and fast reacting organisation
perfectly suited to this kind of programme.
2000
Fig. 9: Programme Status
5
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iIc,II.T..nr, TFU
REFERENCES
The MTR390 engine is being developed together
with a comprehensive Integrated Logistics Support
(ILS) programme.
(1) Trappmann K., Ducos J.S. and Sanderson A.R.
"MTR390 - A New Generation Turboshaft Engine"
15th European Rotorcraft Forum, 1989, Paper
No. 89 - 90
This ILS programme has been structured and
staffed to optimize engine support resources. In
place support system procedures optimize requirements
for support equipment, spare parts, personnel resources and other related logistics resources.
(2) Trappmann K.
"Design Characteristics of a New Generation
Turboshaft Engine"
4th International Symposium on Air Breathing
Engines, 1985, ISABE 85 - 7047
The major task of ILS is to assure the identification of requirements of each support resource and
that exacting and thorough on-going procedures are
established to govern the planning, design, procure- ment, timely deployment and operational evaluation of each resource.
(3) Hourmouziadis G. and Albrecht G.
"An Integrated Aero/Mechanical Approach to High
Technology Turbine Design"
AGARD 69th Symposium of the Props. & Energ.
Panel, 1987
(4) MTR Document
"MTR390 Turboshaft Engine Briefing"
MD 6.002, Issue 1, 1990
The primary features of the ILS programme are:
- to minimize the logistics cost impact of introducing
a new engine:
(5) Spirkl Dr. A., Muggli Dr. W. and Holly L.
"A New Proposal for an Old Problem - The right
engine for the right helicopter"
17th European Rotorcraft Forum, 1991, Paper
No. 91 - 51
to ensure condition monitoring maintenance for
maintenance significant items
to maximise inspection periods
to minimize maintenance operation items
• to favour maximum use of standard hand tools
• to produce Technical Manuals according to ATA
100
to provide Logistics Material Support according
to AECMA 2000M
• to provide comprehensive training suited to
user personnel skill level
- to ensure operational supportability:
a Logistics Support Analysis (LSA) Programme,
and reporting LSAR, according to Mil-Std 13882A, assure all support requirements have been
identified and provide support data compatible
with user data systems
elements of operation and support which are
peculiar to the user are entered into the LSA
process to generate a LSAR and other supportability findings peculiar to each service and
its unique needs
maintainability demonstrations are performed
to validate task times, skill levels, training,
technical manuals, support equipment and spare
parts requirements
CONCLUSIONS
The MTR390 is a state of the art engine, well
supported by an extensive background of research
and demonstrator engines to ensure trouble free introduction of the lates technology.
Thanks to this background the programme is running on schedule and the MTR390 will provide the
modern solution not only for the already committed
TIGER helicopter programe but also to fulfil other
military or civil needs for turboshaft in the 1000
kw range.
6
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