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ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT~ETC F/t 21/5 HIGH temperature PROBLEMf, IN GAS TURBINE ENGINES. (U)

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ADVISORY GROUP FOR AEROSPACE RESEARCH & DEVELOPMENT

7RUEANCELLE 92200 NEUJLLY SUR SEINE FRANCE

1 AGARD CONFERENCE PROCEEDIN(

1 High Temperature Pro

1 in Gas Turbine Engi

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5S No. 229 1

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APR 18 1978 jj 1

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NORTH ATLANTIC TREATY ORGANIZATION

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DISTRIBUTION AND AVAILABILITY ON BACK COVER

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NORTH ATLANTIC TREATY ORGANIZATION ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT (ORGANIZATION DU TRAITE DE L’ATLANTIQUE NORD)

AGARD Conference Proceedings No.229

HIGH TEMPERATURE PROBLEMS

IN GAS TURBINE ENGINES

Appioved [c.i public lolaasa; Distribution Unlimited

Papers presented at the 50th Meeting of the AGARD Propulsion and Energetics Panel held at the Faculty of Engineering, Middle East Technical University, Ankara, Turkey from

19—23 September 1977

THE MISSION OF AGARD

The mission of AGARD is to bring together the leading personalities of the NATO nations in the Helds of science and technology relating to aerospace for the following purposes;

- Exchanging of scientific and technical information;

- Continuously stimulating advances in the aerospace sciences relevant to strengthening the common defence posture;

- Improving the co-operation among member nations in aerospace research and development;

- Providing scientific and technical advice and assistance to the North Atlantic Military Committee in the field of aerospace research and development;

- Rendering scientific and technical assistance, as requested, to other NATO bodies and to member nations in connection with research and development problems in the aerospace field;

- Providing assistance to member nations for the purpose of increasing their scientific and technical potential;

- Recommending effective ways for the member nations to use their research and development capabilities for the common benefit of the NATO community.

The highest authority within AGARD is the National Delegates Board consisting of officially appointed senior representatives from each member nation. The mission of AGARD is carried out through the Panels which are composed of experts appointed by the National Delegates, the Consultant and Exchange Program and the Aerospace Applications Studies Program. The results of AGARD work are reported to the member nations and the NATO Authorities through the AGARD series of publications of which this is one.

Participation in AGARD activities is by invitation only and is normally limited to citizens of the NATO nations.

The content of this publication has been reproduced directly from material supplied by AGARD or the authors.

Published February 1978

Copyright © AGARD 1978 All Rights Reserved

ISBN 92-835-0209-4

Printed by Technic^ Editing and Reproduction Ltd Harford Houae, 7—9 Charlotte St, London, WIP I HD

U

PROPULSION AND ENERGETICS PANEL

Chairman; Dr Ing. G.Winterfeld, DFVLR, Institut fQr Antriebstechnik, Postfach 90 60 58, D-5000 K6ln 90, Germany Deputy Chairman: Dr J. Dunham, NGTE, Pyestock, Famborough, UK

PROGRAM COMMITTEE

Dr D.K.Hennecke (Chairman), MTU, Munchen, Germany Prof. F.J.Bayley, University of Sussex, Falmer, Brighton, UK Prof. J.Chauvin, VKl, Rhode-St-Gen6se, Belgium 1C J.F.Chevalier, SNECMA, Villaroche, France Prof. D.Dini, Univeisita di Pisa, Italy Mr N.F.Rekos, NASA, Washington, DC, United States

i

HOST NATION COORDINATOR

Col. D.Kaya, Ministry of National Defence, Department of Research and Development (ARGE), Ankara, Turkey

PANEL EXECUTIVE Dipl. Ing. J.H.Krengel, D.I.C.

ACKNOWLEDGEMENT

The Propulsion and Energetics Panel wishes to express its thanks to the Turkish National Delegate to AGARD for the invitation to hold its 50th Meeting at the Middle East Technical University in Ankara, Turkey and for the facilities and personnel made available for this meeting.

CONTENTS

PROPULSION AND ENERGETICS PANEL

TECHNICAL EVALUATION REPORT by R.Eggebrecht and S.Lombardo

SESSION I - MOTIVATION AND SURVEY

PROJECT OPTIMISATION OF MILITARY GAS TURBINES WITH RESPECT TO

TURBINE LIFE

by E.A. White and M.J.HoUand

PROBLEMES DES HAUTES TEMPERATURES DANS LES PETITES TURBOMACHINES par B.Belaygue

PROGRESS IN ADVANCED HIGH TEMPERATURE TURBINE MATERIALS,

COATINGS, AND TECHNOLOGY by J.C.Freche and G.M.Ault

THE STATUS OF SMALL, COOLED, AXIAL-FLOW TURBINES by H.F.Due, A.E.Easterling and J.E.Haas

SESSION II - TURBINE COOLING TECHNIQUES

ADAPTATION D’UN BANC DE TURBINE AUX RECHERCHES POUR LES HAUTES TEMPERATURES

par J.Fran(ois, Y.Le Bot, I.Michard et P.Deguest

HOT CASCADE TEST RESULTS OF COOLED TURBINE BLADES AND THEIR APPLICATION TO ACTUAL ENGINE CONDITIONS

by H.Kohler, D.K.Hennecke, K.Pfaff and R.Eggebrecht

INVESTIGATIONS ON THE LOCAL HEAT TRANSFER COEFFICIENT OF A CONVECTION COOLED ROTOR BLADE

by W.Kiihl ,

INVESTIGATION ON TEMPERATURE DISTRIBUTION NEAR FILM-COOLED AIRFOILS

by H.Kruae

EROSION PREVENTION AND FILM COOUNG ON VANES by N.Saryal, I.N.Ghantous and A.Cit9i

PERFORMANCE AND DESIGN OF TRANSPIRATION-COOLED TURBINE BLADING by F.J.Bayley

THE INFLUENCE OF TRANSPIRATION COOLING ON TURBINE BLADE BOUNDARY LAYER

by L.S.Han and L.Winget

EXPERIMENTAL EVALUATION OF A TRANSPIRATION COOLED NOZZLE GUIDE VANE

by A.W.H.Morris, J.B.BuIIard and L.D.Wigg

HEAT TRANSFER CHARACTERISTICS OF THE CLOSED THERMOSYPHON SYSTEM by R.W.Stuart-MitcheU and J.Andries

HEAT TRANSFER FROM TURBINE AND COMPRESSOR DISCS by J.M.Owen

w

.v -

Page

iii

vii

Reference

1

2

3

4

5

6

7

8

9

10

11

12

13

14

Reference

SESinON m - COMBUSTORS, AFTERBURNERS, AND NOZZLES

REVUE DES TECHNIQUES DE PROTECTION THERMIQUE DES PAROIS DES FOYERS PRINCIPAUX ET DE RECHAUFFE DES TURBOREACTEURS

par M.Buisson, J.P.Gaillac et B.Deroide 1 5

PRACTICAL SOLUTIONS TO THE COOLING OF COMBUSTORS OPERATING AT HIGH TEMPERATURES

by J. Winter and H.Todd 16

THE INFLUENCE OF COOLANT TURBULENCE INTENSITY ON FILM COOLING EFFECTIVENESS

by R.Best 17

HIGH TEMPERATURE H,-AIR VARIABLE GEOMETRY COMBUSTOR AND TURBD4E:

TEST FACILITY AND MEASUREMENTS (not presented at the Meeting)

by L.Martorano and D.Dini Ig

LOW FREQUENCY COMBUSTION INSTABILITY IN AUGMENTORS

by F.N. Underwood, J.P.Rusnak, R.C.Emst, E.A.Petrino and P.L.RusseU 19

SESSION IV - MATERIALS AND COATINGS

MATERIAUX D’AVENIR POUR TURBINES A HAUTE TEMPERATURE. LES COMPOSITES

ONERA FACE AUX PROBLEMES D’AUBES

par H.Bibring 20

HIGH TEMPERATURE CORROSION OF Ni-BASE FOR TURBINE BLADE ALLOYS IN SULPHATE-CHLORIDE CONTAINING ENVIRONMENTS

by U.Ducati, G.Lecis Coccia, P.Cavalotti and F.Borile 21

PROTECTION D’AUBES REFROIDIES A STRUCTURE INTERNE COMPLEXE

par P.Galmiche 22

COBALT-BASE ALLOYS FOR HOT CORROSION PROTECTIVE COATINGS

by A.Davin, J.M.Drapier, D.Coutsouradis and L.Habraken 23

SESSION V - MECHANICAL PROBLEMS

TRENDS O*' FUTURE TURBINE LIFE PREDICTION; TIME PHASED AUTOMATED ANALYSIS AND TEST VERIFICATION

by J.L.Price and IJ.Cietshon 24

FINITE ELEMENT ANALYSIS OF SOME PROBLEMS ARISING IN COOLED TURBINE BLADES

by P.Beckers, G.Sander and M.Hogge 25

Paper 26 Cancelled

EVALUATION OF A CERAMIC COMBUSTION CHAMBER FOR A SMALL GAS TURBINE ENGINE

by G.Sedgwick 27

SESSION VI - EFFECT OF COOLING ON AERODYNAMIC PERFORMANCE

SYSTEMATIC STUDIES OF HEAT TRANSFER AND niM COOLING EFFECTIVENESS

by J.F.Louis 2S

EFFECTS OF FILM INJECTION ON PERFORMANCE OF A COOLED TURBINE by J.D.McDonel and J.E.Eiawerth

29

Reference

THE INFLUENCE OF JETS OF COOUNG AIR EXHAUSTED FROM THE TRAILING EDGES OF A SUPERCRITICAL TURBINE CASCADE ON THE AERODYNAMIC DATA

by O.Lawaczeck 30

SESSION VII - MEASURING TECHNIQUES

A NEW TRANSIENT CASCADE FACILITY FOR THE MEASUREMENT OF HEAT TRANSFER RATES

by D.L.SchuItz, T.V.Jones, M.L.G.Oldfleld and L.C.Daniels 3 1

HEAT TRANSFER TO A PVD ROTOR BLADE AT HIGH SUBSONIC PASSAGE THROAT MACH NUMBERS

by B.W.Martin, A.Brown and S.E.Garrett 32

TECHNIQUES DE MESURE DANS LES TURBINES A HAUTES TEMPERATURES

par Y.Le Bot, M.Charpenel et P-J.Michard 33

THE MEASUREMENT OF FILM COOLING EFFECTIVENESS ON TURBINE COMPONENTS IN SHORT DURATION WIND TUNNELS

by J.P.Ville and B.E. Richards 34

LOCAL FLAME TEMPERATURE MEASUREMENTS BY RADIATIVE METHODS

by U.Ghezzi, G.Zizak, A.Coghe, F.CignoIi and S.Benecchi 35

SESSION Vm - PREDICTION METHODS

METHODE NOUVELLE DE CALCUL DE L’EFFICAQTE DE REFROIDISSEMENT DES AUBES DE TURBINE PAR FILM D’AIR

par E.Le Grivds and J.J.NicoIas 36

THE EFFECT OF FREE-STREAM TURBULENCE UPON HEAT TRANSFER TO TURBINE BLADING

by F./.Bayley and R.W.Milligan 37

FLOW AND HEAT TRANSFER IN ROTATING COOLANT CHANNELS

by W.D.Morris 38

CALCULATION OF TEMPERATURE DISTRIBUTION IN DISCS AND COOUNG F^jOW IN A TRANSIENT STATE (not presented at the Meeting)

by M.Caprili and R.Lazzeretti 39

A COMPARISON BETWEEN PREDICTED AND MEASURED SPECIES CONCENTRATIONS AND VELOQTIES IN A RESEARCH COMBUSTOR

by W.P.Jones, W.C.tTifford, C.H.Priddin and R.de Chair 40

TECHNICAL EVALUATION REPORT ON 50TH PROPULSION AND ENERGETICS PANEL MEETING ON HIGH TEMPERATURE PROBLEMS IN GAS TURBINE ENGINES

by

R. Eggebrecht MTU Miinchen GmbH

8000 Miinchen 50, Germany

and

S. Lombardo Curtiss-Wright Corp.

Wood-Ridge, N.J., USA

1. INTRODUCTION

The 50th Meeting of the Propulsion and Energetics Panel of the NATO Advisory Group for Aerospace Research and Development was held at the Middle East Technical University in AnKara, Turkey, from 19 to 23 September, 1977. The purpose of the meeting was to review the status of the technology associated with the design and operation of gas turbines at high turbine inlet temperatures. The conference program was arranged by a committee under the chairmanship of Dr. D.K.Hennecke.

The timing of the meeting was most appropriate, since the last meeting devoted to high temperature turbines was the 36th PEP Meeting in Florence, Italy, in 1970. Since that time, much progress has been made in the under- standing and the application of increased entry temperatures to both military and civil aircraft gas turbines.

The maximum cycle temperature at which today’s aircraft gas turbines are designed to operate is increasing as rapidly as the technology of high temperature materials and cooling methods will allow. Increases in cycle operating temperatures result in higher specific output and increased cycle efficiency. From an aircraft systems point of view, the higher specific output raises the thrust-to-weight ratio of the engine, with significant reduction in engine frontal area and nacelle drag. Thus, the benefits of operating an aircraft gas turbine at increased cycle temperatures can be translated into additional payload or range or a combination of both. All modem aircraft, as well as industrial gas turbines, operate at cycle temperatures which require turbine vane and blade cooling, as well as special cooling configurations for other hot section components such as combustors, shrouds, discs, afterburners, etc. In conjunction with cooling, special considerations in materials, and coating selections are required to insure the integrity of the design and reliable operation of these advanced engines. In this meeting, major emphasis was placed on the state-of-the-art of high cycle temperature gas turbines, with regard to heat transfer, performance and materials technology and their intenelationships. In addition, the meeting covered new developments under investigation which offer significant improvements in the performance, cost, efficiency and reliability aspects of advanced gas turbine engines.

The conference consisted of 39 papers adequately balanced in subject matter and between representatives from industry and government-sponsored organizations. The number of countries providing papers was nine, though specialists from other NATO countries were present and took an active part in the discussions.

The conference was divided into eight sessions dealing with the general theme of high temperature problems in gas turbine engines. It is apparent from the attached listing of these session titles and reference papers that the meeting covered a broad scope of activities relating to the high temperature aspects of gas turbines.

The following review ''' evaluation uses slightly different headings for comments on the papers presented and for emphasis of some fei..,.ies which are thought to have an important impact on future developments of high temperature turbines.

2. PROGRESS OF NEW RESEARCH AND DEVELOPMENT TEST FACILITIES 2.1 Static Cascade Rigs

Many of the data presented at the meeting dealt with test results from different cascade rigs, and the progress

vii

i

made to convert these data into a form which would be useful to designers for predicting the performance and durability aspects of high temperature hot section components.

The present status of highly developed cooling technologies was only reached with the help of sophisticated laboratory cascade testing which allows detailed analysis of basic phenomena for comparison with, and further develop- ment of, theoretical methods. However, there is still a fundamental need to increase cooling effectiveness with a minimum of engine performance losses, while the engine manufacturer is resorting to the use of improved manufacturing methods, materials and coatings, which will offer further temperature potential. One can easily foresee that future research work will be faced with new and sometimes more complex cooling configurations for all the hot-end components of advanced gas turbines.

The very encouraging progress is worth noting which has been achieved using the so-called short-duration test facilities at Oxford University, VKl and MIT reported during this meeting by D.L. Schultz (31)*, R.E. Richards (34) and J.F.Louis (28). The tests can be performed at actual engine Reynolds and Mach numbers and actual gas-to-wall temperature ratios. Heat transfer measurements are made by highly developed thin-film techniques, aiid fast data acquisition systems are available for recording and subsequent processing of transient data.

Interesting cascade testing with measurements of heat transfer along a PVD blade profile has been reported by B.W.Martin (32) from Wales University. These tests also employed a transient method by measuring temperature - time responses of blade surface thermocouples when the blade was suddenly introduced into a heated air stream.

One of the key problems in cascade testing, whether in the steady-state or transient mode, is the proper simulation of environmental conditions prevailing under actual engine conditions and influencing, for instance, the cooling performance of vanes, blades, shrouds and combustor liners. F.J.Bayley (37) stated the present position very clearly, when he pointed out that even the relevant characteristics of the engine flow are not yet definable, not to mention simulation under laboratory conditions. H.Kohler (6) in his paper compared surface temperature measurements and associated heat transfer coefficients from static cascade tests with results on a comparable rotor blade operated in an engine at similar Reynolds and Mach numbers and with the same gas-to-wall temperature ratio. The large discrepancies observed, mainly on the leading edge ana pressure surface, highlight the possible effects of engine-related environmental conditions, such as main-stream turbulence originating from unsteady combustion, cooling air admixture and periodic velocity oscillations due to blade wakes.

For investigation of some separate effects of turbulence, F.J. Bayley (37) presented a new experimental set-up at Sussex University consisting of a static cascade with an upstream turbulence generator which was conceived as a rotating squirrel cage. This device allows the investigator to vary the turbulence intensity by using a range of bar diameters and to vary the frequency of velocity fluctuations through the rotational speed. The experiments he reported about were mainly done in the range of up to 6 kHz with measured velocity fluctuations in the range of 24% to 48% turbulence and showed some dramatic effect on local and mean blade profile heat transfer coefficients.

There was general agreement among the speakers that the data obtained from the above described laboratory tests are very useful and necessary for the design of hot-section components.

2.2 Turbine Aerodynamic Rigs and High Temperature Turbine Testing

For investigation of the effect of coolant injection on turbine aerodynamics, cold or warm air turbine rigs are commonly used, as refened to in papers by J.D.McDonel (29) and H.F.Due (4). As reported by A.W.H.Morris (12) at NGTF, a high-temperature single-stage research turbine has been used for recent testing of transpiration-cooled NGV’s at design conditions of 1650 K gas temperature and 4.5 bar inlet pressure.

W.Kuhl (7) described temperature measurements on rotating turbine blades in a single-stage test turbine at the Technische Hochschule, Aachen, which were aimed at analysing blade-profile heat transfer under moderate turbine inlet temperature and pressure conditions up to 1 173 K and 1.5 bar respectively.

A major new test facility presented during this conference in the papers by J. Francois (5) and Y.Le Bot (33) is the so-called French MINOS (Montage inter ONFRA-SNFCMA) operated at CFPr, Saclay.

This test facility is basically a high-temperature test turbine with an upstream engine combustor, in which an attempt is made to simulate engine environmental conditions with respect to the combustor-turbine unit. The max. designed operating temperature is 1800 K and the max. entry pressure delivered from the plant feed system is 4.5 bar.

The authors (5) quote an impressive program of future investigations covering a broad range of high-temperature turbine problems including

turbine aerodynamics

- various heat transfer and film cooling investigations on NGV’s, rotor blades, casings and end walls * Number in parentheses refers to the Paper in the main Conference Proceedings.

vtii

- thermal fatigue tests by means of cyclic variation of cooling air flows

- testing of abradable materials with respect to improvements in running clearances and reduced air leakages.

There is no doubt that each of these problem areas represents a major aspect in the development of advanced turbines. The ambitious targeting for this test facility essentially requires advanced measuring techniques, which are described in the paper by Y.Le Bot (33):

Total pressure and temperature probes designed for fast dynamic response and capable of operating under high temp-'ratures have, for instance, been developed for measuring turbulence at turbine entry, and for analysing rotor downstream wakes. Laser anemometry is seen to be not yet ready for this application. Blade temperatures are being measured by embedded thermocouples and optical pyrometers. For heat transfer analysis on turbine blades, a transient technique involving sudden cooling air flow shut-off is being employed. Heat flux measurements on turbine casing liners can be performed with fluxmeters developed by SNIAS. For tracing cooling air flow paths and evaluating film cooling effectiveness, using the analogy between heat and mass transfer, the rig is designed to allow gas sampling with chromatographic analysis of gas concentrations.

It must be realized that the extent of the instrumentation used causes some changes in comparison with the actual engine situation such as, for instance, wider spacing between blade rows. Another limitation which must be recognized is the rather low operating pressure of the combustor, whose outlet conditions may alter in the actual core engine situation. This comment underlines the author’s opinion that MINOS will at least help to bridge the gap between classical component rig testing and experimental investigations under real engine environmental conditions.

From the measurements already made and described in these two papers, it is evident that MINOS has the capability to make a major contribution to solving future high-temperature turbine research and development problems.

3. COOLING TECHNIQUES AND HEAT TRANSFER INVESTIGATIONS

3.1 Recent Work on Convection-Cooled Turbine Blades

Blade cooling is commonly used in present military and civil engines. However, surprisingly little information is available about the local gas-side and cooling-side heat transfer rate for different blade profiles and internal coolant passage configurations under actual engine operating conditions. In order to distinguish the physical phenomena occurring under these conditions, there is still a fundamental need for heat transfer investigations on cascades and research turbines, as reported on by several authors at this conference. The following table summarizes experimental and/or theoretical work devoted to external turbine blade heat transfer in the absence of boundary layer coolant injection.

Author/Ref.

Paper

Investigation Performed

Varied Parameters

Comments

J.F. Louis

(28)

profile heat transfer distribu- tion (p.h.t.d.) for four tran- sonic blade profiles

outlet Mach number; incidence angle

shock tunnel cascade rig

D.L. Schultz

(31)

p.h.t.d. for a high pressure turbine blade

outlet Reynolds number; Tu level

short-duration wind tunnel operating with single-stroke light piston compression

B.W. Martin

(32)

p.h.t.d. for PVD* turbine profile

outlet Mach and Reynolds numbers; Tu level

blades are shifted into hot gas duct and undergi transient heating

F.J.Bayley

(37)

p.h.t.d. for high pressure turbine rotor blade

outlet Mach and Reynolds numbers; Tu level and fre- quency

steady state cascade tests with upstream squirrel cage turbu- lence generator

W.Kuhl

(7)

p.h.t.d. for turbine blade and cooling effectiveness

cooling air mass flow

test turbine with slip ring equipment

H.KbhIer

(6)

p.h.t.d. and cooling effective- ness for four cooling con- figurations with unchanged outer blade profile

mainly outlet Reynolds number and cooling air mass flow

steady state cascade tests and engine measurements on rotor blades by means of thermal paints

J. Francois

Y.Le Bot

(5)

(33)

p.h.t.d. for internally cooled NGV behind combustor

no parameter variation reported

high temperature turbine rig “MINOS”

A review paper using these newly provided heat transfer data for comparison with previously published results by other authors appears to be very desirable. From the amount of available data one could expect that some fruitful incentives for improvements in heat transfer prediction by means of boundary layer theory may arise.

* Prescribed velocity distribution.

IX

One paper, presented by W.D.Morris (38), dealt with heat transfer in rotating coolant channels as affected by Coriolis forces and rotational buoyancy. It turns out that the use of forced convection data obtained with stationary tubes for the prediction of heat transfer in rotating tubes can lead to significant errors of either over- or under- estimations. On the basis of the already available test results of this research work, which has just started, it seems advisable for blade cooling design engineers to closely watch the further outcome of these investigations.

Turbine blade cooling by means of a closed thermosyphon system would offer the advantage of high internal heat transfer coefficients. The paper by R.W. Stuart Mitchell (13) presented new experimental investigations for the stationary vertical, the stationary inclined, and the rotating closed thermosyphon, with water and mercury as working fluid, and gave dimensionless correlations of the measurements. The results suggest that, in add, cion to the commonly used Grashof number based on gravity acceleration, a dimensionless centrifugal acceleration term also has a marked influence on the Nusselt number of the cylindrical tube under investigation.

3.2 Film Cooling of Hot-End Components

An analysis of film cooling physics for any practical design of turbine blades, end wall elements, stators or combustor and afterburner liners is not yet feasible.

No wonder that an increasing amount of research work is being done. The following table gives a survey of film cooling work reported on.

Author/Ref.

Paper

Configurations

Range of Thermodyn. Parameters

Comments

J.F.Louis

(28)

angular injections for flat plate streamwise angles 1 , 20°, 30°, crosswise angles

0°, 30°, 50°, 70°, 90°

blade profile with various film-cooling ejections

Mao = 0,5

To = 556 K

To = 122... 278 K

rii = 0,1 . . . 1,6

exit Mach number = 0,6

To = 450 K

T(. = 293 K

shock tunnel with about 1 0 ms steady flow test time

PcWc

m =

PoWc

B.E. Richards

(34)

flat plate with injection through double row of holes with 30° stream-wise injection angle

Mao ~ 0,6

To = 382 K

To = 267 . . . 365 K m = 0,5 ... 1,5

short duration wind tunnel operating with single-stroke light piston compression

J. Francois

(5)

turbine casing; end wall “MINOS” turbine rig conditions

H. Kruse

(8)

turbine blade leading edge film cooling flat plate

To = 400 K

To = 293 K

rii = 0,5 . . . 2,0

single blade model tests

R.Best

(17)

slot configuration inside tube

Mao =0,1... 0,16

To = 453 . . . 500 K

To = 293 K

Wg/Wo =0,5... 1,9

variation of coolant side Tu

E.Le Grives

(36)

single and multiple row of holes with various stream- wise and crosswise angles

wide range of blowing parameter; thermo- dynamic data not all given explicitly

flat plate experiments; comparison of test results with new analytical prediction method presented by the author

J.F.Louis (28) reported about a very comprehensive experimental program performed at MIT on film-cooling configurations mainly of single and double line holes for different streamwise and crosswise angular injection. The importance of these experiments becomes clear when it is considered that most of the practical application in turbines make use of injection holes rather than continuous blowing out of slots. Substantial research work on film cooling with various injection hole configurations was performed about one decade ago at the University of Minnesota, as reported by E.R.G. Eckert at the 1970 AGARD PEP Meeting, and at Arizona State University by D.E. Metzger et al. These investigations were made for very low mainstream Mach numbers and in some cases in the transonic range. The temperature ratio between coolant and mainstream was close to 1 .0. At the same conference C.Liess reported about measurements at VKI downstream of inclined injection holes which were taken at elevated Mach numbers of 0,4 . . . 0,6 but still with only small temperature differences. The present shock tunnel tests at MIT cover nearly the full range of thermodynamic parameters which occur in advanced turbines. The overall correlation used is essentially based on equivalent slot width, the square root of momentum ratio to the 1.35th power and the coolant Reynolds number to the 0.25th power. It must be noted that this correlation describes fairly well the experimental results of individual injection hole arrangements. However, the attempt to describe the isothermal efficiencies of the very different geometric configurations by such an overall correlation parameter leads to a rather wide scatter of data and therefore

is still unsatisfactory. It is interesting to note that Louis can align the effectiveness of hole and slot configurations by a simple geometric “mixing area” correction.

The very high potential of presently available test equipment becomes evident from the presentation by D.L.Schultz (31) and D.E.Richards (34) on tests with the so-called Isentropic Piston Tunnel at Oxford and VKl, respectively. Film cooling experiments using a double row of holes with a 30° injection angle were reported on by Richards. The test results show similar trends to the previous investigations by Eriksen and Goldstein, which were performed under incompressible flow conditions. So far, the results indicate that there is obviously no strong influence of increased Mach numbers. It must be borne in mind, however, that primarily, these tests should prove experimentally the linear relationship between “overall heat transfer coefficient” (based on difference of mainstream and wall temperature) and a non-dimensional coolant temperature. Furthermore, of course, Richards demonstrates the capabilities of the transient test technique based on single-stroke isentropic compression rather than performs any systematic study on film cooling configurations.

As far as continuous blowing is concerned some different analytical methods are already available. In his paper J.F.Louis (28) refers to Demiijian, whose mathematical modelling for angular injection predicts quite well the film behavior in the region near the slot injection and up to blowing rates at which boundary layer lift off occurs and consequently the cooling effectiveness is reduced.

For injection through discrete holes, a new analytical technique was presented by E.Le Grives (36). It must be appreciated that this paper already demonstrates impressive progress in theoretical methods for describing interaction between mainstream and single jets and array of jets by means of the dilution theory. The paper provides valuable references to previous publications by other authors and points out that the future work of the authors will be focussed on curvature effects.

Generally speaking, more fundamental experimental data are obviously needed in order to develop computational methods for the highly complex three-dimensional flow situation in the case of film cooling with hole injection configurations which are aimed at improved cooling effectiveness. Furthermore, this future work will have to simulate more closely engine environmental conditions, in order to study the effect of hot gas mainstream flow characteristics prevailing in turbines behind engine combustors.

In his paper (17) R.Best draws attention to the velocity profiles and turbulence distribution of the coolant flow in the plane just before entering the mainstream boundary layer, which he measured for different slot widths.and for a wide range of blowing rates. He shows that there is an adverse effect of the entering coolant turbulence upon film cooling effectiveness. This effect is more pronounced for coolant to mainstream velocity ratios W^/Wq = 1 . His semi-empirical model fairly well describes the observed experimental phenomena of this type of tangential slot film cooling.

Rather little is known about the interaction of film cooling jets with the mainstream boundary layers of turbine blade or vane leading edges, which have to withstand the highest thermal loading. H. Kruse (8) reported about boundary layer measurements using a minature temperature probe in the vicinity of a turbine blade leading edge simulated by a single airfoil mounted in a small tunnel with adjustable flexible walls. From these investigations, it becomes evident that, for differently angled injection hole arrangements, there is a strong influence of the coolant blowing rate upon the local cooling effectiveness. Furthermore, it is shown that any changes in the stagnation point severely affect the cooling performance when there is only one row of holes near the leading edge.

3.3 Transpiration Cooling

L.S.Han (11) presented a paper on the analytical studies being conducted at Ohio State University on the influence of transpiration cooling on turbine blade boundary layers. The authors described a method by which the external boundary layer and heat transfer distribution can be calculated.

The experimental and theoretical work by F.J.Bayley (10) reported on at this meeting confirmed again the very high cooling effectiveness of transpiration cooled turbine vanes and blades compared with other cooled blade configura- tions. He also pointed out the excellent correlation of the heat transfer aspects pertaining to the design of transpiration cooled components. In his paper A.W.H.Morris (12) reported on the experimental evaluation of a transpiration cooled nozzle guide vane. The test program was conducted to evaluate the thermal design of the NGV and to determine the influence of the transpiring flow on stage efficiency. Cascade tests and a single stage high temperature turbine test rig were utilized in this program. The vanes used in this evaluation consist of a POROLOY porous metal airfoil, diffusion bonded to the main structural element. The cascade and engine test results demonstrated again the uniformity in airfoil metal temperatures and the high effectiveness of transpiration cooled blades and vanes. On the mechanical aspects of transpiration air cooled blades, the authors, on the basis of the single stage rig tests, conclude that the uniformity in airfoil metal temperatures possible with transpiration cooled blades and vanes will result in reduced thermal stresses and propensity to thermal cracking. Their test results lead them to believe that pore blockage of the transpiration cooled airfoil structure is not a significant problem. In addition, inadvertent exposure of the turbine to foreign object damage showed extensive maltreatment of the transpiration cooled blades could be tolerated without disastrous consequences. All the above mechanical characteristics of transpiration cooled blades, confirm the test

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results noted earlier by another author and presented at the 36th PEP Meeting in Florence, Italy, in 1970. In noting the quality of the transpiration cooled vanes of the current paper and the quality of the vanes of the earlier reported work, it is obvious that significant strides have been made, over the past seven years, on the fabrication aspects of transpiration cooled structures.

As regards the performance aspects, the authors presented cascade data on profile loss coefficient versus coolant flow ratios on fully transpiration cooled nozzle guide vanes (NGV), as well as for transpiration air cooled vanes with various percentages of the suction surface blocked. Using these data and test results from the testing of the vanes in a single stage turbine test rig, the authors then conducted engine cycle studies comparing the sfc and thrust relation- ships of a transpiration cooled NGV configuration with a conventionally cooled NGV. On the basis of this study, the authors concluded that direct substitution of transpiration cooled NGV offered no significant performance advantages over the higher coolant flow usage of a conventionally cooled NGV turbine stage. In their assessment of turbine efficiency when using transpiration air cooled NGV’s the authors, in applying their cascade results to the cycle studies, have defined turbine efficiency according to the method outlined by L.Y.Goldman of NASA. Consequently, the derived turbine-stage thermodynamic efficiency decreases markedly with increasing coolant flow. The simple application of this efficiency correlation on engine cycle studies, however, appears to be inconsistent with the findings of other investigators. Additional research and engine development testing is indicated to clarify the situation.

3.4 Rotating Disc Heat Transfer

In order to meet the requirements of advanced analytical techniques in the structural design of compressor and turbine discs, improved prediction methods for steady-state and transient temperature distributions are necessary.

The solution of the basic heat-conduction equations for any geometrical configuration appears to be no longer a problem and several mathematical routines are available which tend to use effective finite-element methods. This was also indicated in the paper by M.Caprili (39). It gives, however, details of a different mathematical approach to disc temperature calculation in the case of prescribed surface heat transfer coefficients. Furthermore, the paper demonstrates, in a parametric study, the effect of disc heat transfer coefficient and coolant mass flow on radial temperature distri- butions in a typical turbine disc.

In a general comment, it must be stated that, for several rotating disc arrangements, the analytical methods for calculation of the heat transfer boundary conditions are often based on rough empirical methods. It can be said that past progress in this field of heat transfer research has not kept pace with fast heat conduction solution procedures which are nowadays being widely used. Therefore, it must be appreciated that one paper by J.M.Owen (14) was devoted solely to the very problem of heat transfer from turbine and compressor discs. Whereas several previous publications have already dealt with different rotating disc and cavity arrangements, this paper in particular presents heat transfer measurements for

(a) the situation of central axial throughflow and

(b) the situation of radial outflow of coolant between co-rotating discs.

The experiments reveal strong vortex breakdowns for the situation (a) and identify different heat transfer regimes for the situation (b). Judging from the present results, the author’s concluding view must be shared that much more research work is necessary for establishing theoretical or even empirical prediction methods.

4. EFFECT OF TURBINE COOLING ON AERODYNAMIC PERFORMANCE

From the angle of aerodynamic losses, the most attractive blade profile position for ejection of cooling air is seen to be the trailing edge of the blade. O.Lawaczeck (30) presented cascade wake flow measurements in a wide range of downstream subsonic to supersonic flow conditions. The experimental results provide basic turbine design data in terms of downstream flow angles and loss coefficients for this type of coolant ejection.

The evaluation of the effects of film-cooled vanes and blades on turbine aerodynamic performance and the effect on overall cycle thermodynamic efficiency was the subject of the paper presented by J.D.McDonel (29). The testing was done on ^ sin^e-stage turbine test rig which featured five independent coolant supplies for independent variations of coolant-to-mainstream temperature ratios, pressure ratios, and mass flow ratios. The program included five test configurations, including two different film cooling designs, and three combinations of film-cooled and solid airfoils. Test results were presented showing the effects of the individual and combined vane and blade cooling air flow ratios on overall stage efficiency. These results were then compared with previously reported analytical methods, and the correlation was quite good. Using the results of the turbine rig, McDonel conducted a cycle analysis study program on a typical high temperature high-performance core engine to demonstrate the effects of cooling air utilization on overall engine performance. The base line turbine inlet temperature of the core engine was 1478 K (2200° F).

The results of this study were presented in parametric form, showing how engine output and efficiency varied with cooling air flow usage for various increases in turbine inlet temperature. This paper clearly demonstrates the potential cycle performance gains resulting from increases in engine cycle temperatures. In addition, it also points out

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that, if the cooling is inefficient, the iricreased coolant flow lates and film injection losses can erode the potential cycle performance gains very rapidly. Curves were presented which showed specific limits which must be placed on the coolant flow rate for specific temperature increases. These results should be helpful tc designers in establishing coolant flow limits during engine preliminary design studies. This paper was most timely, since film cooling of vanes and blades is now well established in modern aircraft engines.

Small turbines are generally accepted “to be different” and to have their own problems. During the last 10 to 15 years the small, cooled, axial-flow turbine has been the subject of several research programs. H.F.Due (4) presented a very useful review paper on the special aspect of the aerodynamic performance of the small turbine. The author presented several experimental results of various US-industry and government-sponsored investigations and concluded that considerable efforts are still necessary in order to improve turbine design methods, with special emphasis on prediction of coolant effects.

Besides the author’s statement, it is believed that the expected better understanding of aerodynamic losses will guide new approaches to cooled turbine designs which offer still further potential for reduction of losses owing to adverse interaction between turbine mainflow and discharging cooling air.

5. COMBUSTORS AND AFTERBURNERS

Investigation of different liner cooling configurations of combustors and afterburners of aero engines was the general subject of the paper presented by M.Buisson (15). It presents rig measurements of cooling effectiveness by means of the gas analysis technique and the application of thermal paints and outlines the basic features of a simplified analytical approach which is being used for predicting the wall temperature of combustors and afterburners. Furthermore the advantage of combustor liner sandwich design, which employs effective convection cooling before coolant ejection takes place, is emphasized.

J. Winter (16) discussed various practical solutions for combustor cooling problems typically associated with a reverse-flow annular combustor and with a cylindrical flame tube combustor operating in a regenerative gas turbine engir e. This paper is mostly devoted to the very typical development problems combustion engineers are faced with, when component life has to be increased or more potential for engine uprating is necessary. The subject of this paper is seen to be very suitably placed in this “High Temperature Problems” conference, the intention of which is, on the one hand, to cover the wide scope of present scientific research work and, on the other hand, to deal with application problems which influence the direction of future research work.

A topical area of combustor-related research work is the development of analytical models which describe exhaust species concentrations as well as overall combustor performance. The paper of W.P.Jones (40) et al., presented by C.H.Priddin, dealt with measurements of species concentrations and velocities in a small-scale research combustor, these being compared with predictions of their mathematical model of chemically reacting flow which uses finite-difference equations. The present status of this model describes the profiles of fuel/air ratios and UHC concentrations quite well, but exhibits shortcomings in the prediction of CO concentrations along the combustor axis. The authors discuss possible approaches to overcome the present limitations in future developments in the model. Without doubt, some of these improvements can be easily incorporated into the existing mathematics as, for instance, a modified probability function or an additional reaction mechanism for NOX formation.

The authors regard the introduction of adequately prepared fuel breakdown physics, with model capabilities to describe ignition and extinction limits, as a rather more longterm development. Hesitation may, therefore, be justified in sharing the optimism expressed in the authors’ concluding statements which suggest that only a little further development is necessary.

Alternative aviation fuels under consideration for future aircraft engines will influence especially the combustor system design. The paper by L.Martorano (18) deals with Hj-air combustion in a coaxial-stream cylindrical combustor up stream of a small single-stage research turbine featuring variable nozzle guide vanes.

The primary zone air loading of the combustor can be varied by movable inlet baffles and testing has been done over a wide range of fuel/air ratios, but no specific details of combustor measurements are given. It can be expected that the impact of alternative fuels on engine combustor design, as well as on cooling techniques, will attract increasing interest in future high temperature turbines. In this sense, the subject of this paper is believed to be also of considerable importance for future High Temperature conferences.

The last paper of the combustion session was devoted to the severe problem of low-frequency combustion in mixed-flow afterburners known as rumble or chugging. F.N.Underwood (19) categorised the several possible mechanisms which are normally seen to cause or regulate this special phenomenon. The paper gives a status report on a research project which is aimed at developing a reliable empirical and analytical model to aid afterburner design.

The experimental rig test data presented identify airflow dynamics and fuel distribution as main rumble contributors and the overall mathematical model of the augmentor system is shown to already predict typical rumble conditions.

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Improvements by incorporation of a more adequate combustion model are necessary and were announced by the author.

6. HIGH TEMPERATURE MATERIALS AND COATINGS

Six papers were given, discussing the properties, characteristics, and selection of materials for use in hot section components operating at high turbine inlet temperatures. G.M.Ault (3) presented a very comprehensive survey on the status, progress, and future potential of advanced processes, materials, and coatings currently under development by the gas turbine community for advanced high temperature engines. As noted in previous papers (12, 29), significant payoff in engine performance can be achieved by minimizing the amount of cooling air used. Thus, the development of advanced materials and coatings, together with the development of improved cooling techniques are keys to realizing the full performance benefits of the high temperature gas turbine. In his paper, Ault (3) also predicted that pre-alloyed, powder-metallurgy-processed super alloys will afford increased strength and