Table Of ContentAlready published in this series:
FLIGHT TEST INSTRUMENTATION, Volume 1 Edited by M. A. Perry, 1961.
RECENT DEVELOPMENTS IN NETWORK THEORY Edited by S. R. Deards, 1963.
FLIGHT TEST INSTRUMENTATION, Volume 2 Edited by M. A. Perry, 1963.
ADVANCES IN AUTOMOBILE ENGINEERING, Part I "Edited by G. H. Tidbury,
1963.
ADVANCES IN AUTOMOBILE ENGINEERING, Part II Edited by N. A. Carter,
1963.
FLIGHT TEST INSTRUMENTATION, Volume 3 Edited by M. A. Perry, 1965.
ADVANCES IN AUTOMOBILE ENGINEERING, Part III Edited by G. H. Tidbury,
1965.
ADVANCES IN AUTOMOBILE ENGINEERING, Part IV Edited by D. Hodgetts,
1966.
AEROSPACE INSTRUMENTATION, Volume 4 Edited by M. A. Perry, 1967.
COMBUSTION IN ADVANCED GAS TURBINE SYSTEMS Edited by I. E. Smith,
1968.
COMBUSTION AND
HEAT TRANSFER IN
GAS TURBINE SYSTEMS
Proceedings of an International Propulsion Symposium
held at the College of Aeronautics, Cranfield,
April 1969
Edited by
E. R. NORSTER
PERGAMON PRESS
OXFORD NEW YORK TORONTO
SYDNEY BRAUNSCHWEIG
Pergamon Press Ltd., Headington Hill Hall, Oxford
Pergamon Press Inc., Maxwell House, Fair view Park, Elmsford,
New York 10523
Pergamon of Canada Ltd., 207 Queen's Quay West, Toronto 1
Pergamon Press (Aust.) Pty. Ltd., 19a Boundary Street,
Rushcutters Bay, N.S.W. 2011, Australia
Vieweg & Sohn GmbH, Burgplatz 1, Braunschweig
Copyright © 1971 Pergamon Press Ltd.
All Rights Reserved. No part of this publication may
be reproduced, stored in a retrieval system, or trans
mitted, in any form or by any means, electronic,
mechanical, photocopying, recording or otherwise,
without the prior permission of Pergamon Press Ltd.
First edition 1971
Library of Congress Catalog Card No. 72-143801
Printed in Great Britain at The Pitman Press, Bath
08 016524 9
FOREWORD
WITH the accelerating development of the gas turbine engine in size and
complexity and the growth of its technology, there is an obvious need to
supplement this type of Symposium with others covering related topics in
the Gas Turbine field.
However, the success of the 1967 Cranfield International Propulsion
Symposium stimulated so much interest in gas turbine combustion systems
that the Organizing Committee felt the format for the 1969 Symposium should
be virtually the same.
The papers selected for this Symposium gave a good balance between the
work of academic and research establishments and that of industrial organi
zations in the field of Combustion and Heat Transfer in Gas Turbine
Systems. Additional topics of atomization, fuels and high-temperature
materials have been introduced to highlight the increasing spectrum of the
technology associated with the gas turbine engine. Four of the nineteen
papers presented came from outside the United Kingdom and illustrate the
increasing international flavour of Cranfield Symposia.
The discussion of these papers, exchange of ideas and feedback of infor
mation have not only proved to be useful stepping-stones for technical staff
and designers but also stimulus for those engaged in applied combustion
research.
The editor, on behalf of the Organizing Committee, would like to express
thanks to Authors, Chairmen, Delegates and members of the Propulsion
Department of the College of Aeronautics who contributed to the success of
the 1969 Propulsion Symposium and for making the compilation of this
volume possible.
E. R. NORSTER
Vlll
THE COMBUSTION SYSTEM FOR THE
OLYMPUS 593 CONCORDE ENGINE
D. W. HARVEY
Rolls-Royce Limited (Bristol Engine Division)
Introduction
The original concept of an engine for Concorde was an uprated version of
the previous Olympus Mark, the 320, two of which powered the TSR2.
However, the thrust level of 30,0001b was insufficient when in 1964 com
mercial demands on the aircraft necessitated a minimum of 35,000 lb to
cope with increased payload and range—the latter design requirement being
3500, miles typical of a London or Paris to New York run.
The familiar twin-spool straight turbojet design of the Olympus series was
retained together with a cannular configuration of combustion system—on
which considerable experience had been obtained from the 1000's of opera
ting hours of the previous marks including the early Vulcan power units,
marks 104, 201 and 301 and the 320 in the ill-fated TSR2.
Hence the selection of 8 x 10 in. diam. flame tubes each supported up
stream on a sprayer mounted spherical and downstream on the first stage
turbine stators or nozzle guide vanes (Fig. 1).
Principally due to powerplant installation considerations the same major
scantling dimensions as the Olympus 320 were retained—and although the
inlet and outlet duct heights have been altered to meet the demands of
compressor and turbine, the maximum area annulus dimensions remained
unchanged as they have done since the original 201 mark.
Operating Conditions
Figure 2 shows the increases in the major operating conditions of the
combustion system throughout the Olympus series. Maximum inlet pressure
has steadily risen from 10-17 atm.—a relatively low value by present-day
standards but then a high-pressure ratio engine is not necessary for a S.S.T.
The high outlet temperature requirement which was achievable for the first
time with the introduction of cooled turbine blades is shown by the significant
rise from the 1200°K level up to 1510°K—although this figure includes a
built-in margin used for combustion development, full-thrust Concorde
engines will operate some 50°C below this figure.
3
4 D. W. HARVEY
Inlet temperatures at the take-off conditions have also shown a steady rise
over the years culminating with 450°C in the 593. But here we meet the
significant feature unique to this S.S.T. engine where combustion inlet
temperature at cruise is 100°C greater than the take-off value.
FIG. 1. Early design of Olympus 593 flame tube.
The final curve in Fig. 2 shows the stator choking flow function. Bearing
in mind that casing sizes have remained constant and that at windmilling
conditions combustion system temperatures are similar in all marks, then the
W\/T
increase from a value of —-— at outlet of 45 to nearly 75 is indicative of
the increased severity of altitude ignition and relight. The mean air velocity
through the system is nearly 70% greater than in the Olympus 201,
COMBUSTION SYSTEM FOR THE OLYMPUS 593 5
It is important to appreciate that not only have there been general in
creases in combustion operating levels, but also a very wide range of condi
tions is experienced during the normal transatlantic flight profile of the
Concorde. Apart from wide ranges of inlet pressure, temperature and fuel
flow, the variations in system air loading and air-fuel ratio as shown in
Fig. 3 are particularly demanding.
FIG. 2. The increasing demand on the combustion system at take-off.
The highest loading phases are those of approach and taxi, descent, and a
typical diversion, and unfortunately these are also the régimes during which
the weakest operation of the system is prevalent. Therefore we were faced in
the initial design of either ensuring a reasonable primary zone air-fuel ratio
at these weaker, highly loaded conditions and accept rich operation during
the take-off phase with the inevitable high level of smoke generation—or to
6 D. W. HARVEY
run weak during these phases with consequent loss of thermal efficiency and
relight capability.
Combustion Performance Criticality
From the detailed breakdown of the conditions encountered during a
typical flight profile, fairly simple calculations enable the effect of combustion
inefficiency on take-off weight to be determined. An operating point repre
senting each of seven phases was selected and by considering the total amount
FIG. 3. Typical transatlantic flight profile maximum chamber operating loading,
air fuel ratio and fuel temperature.
of fuel used in that phase the efficiency equation shown in Fig. 4 was derived.
The relative importance of maintaining high efficiency levels during each
phase was therefore obtained and is shown graphically in the same figure.
The cruise condition is critical, a 1 % reduction in efficiency necessitating an
extra 1750 lb of fuel to be included to meet the predetermined flight plan.
These figures are based on an overall aircraft weight at take-off of 350,000 lb,
including a payload of 20,000 lb. Next in order of importance are the sub
sonic climb, diversion or hold at alternate and the descent phase, all of which
COMBUSTION SYSTEM FOR THE OLYMPUS 593 7
are conditions of high loading and weak air/fuel operation. If a 1 % loss in
efficiency at all conditions is incurred, then the take-off weight penalty is
over 2500 lb. The efficiency at take-off, however, is relatively unimportant in
this context.
Because of the considerable performance penalty arising from inefficiency
at these weak operating conditions, it was concluded that a comparatively
rich primary zone operation at the take-off condition was essential. A value
of 14/1 was selected, but even this gives nearly 70/1 during the descent phase.
FIG. 4. Effect on fuel penalty of combustion efficiency. Fuel penalty is computed
by the equation:
where a fuel penalty of 10 = 1520 lb.
The criticality of the combustion performance is also experienced when
pressure loss is considered. Engine cycle calculations show that a 1 % increase
in combustion system loss necessitates an extra 1350 lb weight of fuel to be
put aboard. The rate of exchange is similar for loss reduction giving a worth
while design incentive.
8 D. W. HARVEY
A large proportion of the system loss is incurred in diffusing from the
compressor outlet guide vanes to the maximum area annulus. Gradual
controlled diffusion to achieve minimum loss would necessitate increasing
the length of the system. The actual weight increase/inch length local to the
combustion system for the Olympus 593 engine is 8 lb, representing a take-off
weight penalty for the aircraft of 64 lb/in. This comparatively small increase
would therefore be adequately compensated by the return through reducing
the overall system loss. However, at the design stage length priority was
given to both intakes and compressors, and at the present time such a major
redesign would be uneconomic.
Altitude Relight Considerations
Operating conditions are significantly extended when the possibility of a
flight relight is considered. Inlet pressures of less than half an atmosphere and
temperatures of —40° are liable if the guaranteed 30,000 ft clearance at all
subsonic Mach numbers is to be achieved. The air loading of the chamber
during windmilling is nearly doubled compared with the Oly. 201, making
the problem twice as difficult. The simple loading parameter shown earlier is
not only employed in correlating combustion efficiency but can also be
applied to ignition and relight. Figure 5 shows the lines of constant loading
derived from the windmilling characteristics of an Olympus 593 in Concorde.
Included on this plot are ignition and relight results from an early series of
engine tests in an altitude cell at N.G.T.E. The similarity of the loading and
relight loops is clearly shown.
Ignition of the igniter flame tubes must, of course, be followed by cross-
ignition via interconnectors to the adjacent flame tubes together with
sufficient heat release to provide the necessary energy for accelerating the
engine. Within a range of normal fuel flows thermal efficiency at this condi
tion is therefore important—noting that a loading value of approximately
23 is required at 30,000 ft M = 0-6 the demanding conditions can be seen
in Fig. 6 where a typical loading/combustion efficiency plot is shown.
Efficiency levels of less than 40 % may be expected at such air loadings.
Metal Temperatures
The fixing of casing sizes to be the same as the Oly. 320 therefore accen
tuated the problems of relight and the achievement of acceptable efficiency
levels. However, as this constraint led to smaller sized flame tubes than would
have otherwise resulted from a "carte-blanche" design the problem of flame
tube wall temperatures was at least minimized. Fundamental rig tests have
shown that all the flame tube metal temperatures increase at the same rate
as inlet temperatures, whilst the dilution and turbine entry duct zones
