AD-A264 177 (.C/

A GA RDa ADVISORY GROUP FOR AEROSPACE RESEARCH & DEVELOPMENT

7 RUE ANCELLE 92200 NEUILLY SUR SEINE FRANCE

AGARD CONFERENCE PROCEEDINGS 527

Heat Transfer and Coolingin Gas Turbines(Le Transfert Thermique et le Refroidissement L l I •dans les Turbines A Gaz) . T, E

APR.7

Papers presented at the Propulsion and Energetics Panel 80th Symposium heldin Antalya, Turkey, 12th--6th October 1992.

93-08810

NORTH ATLANTIC TREATY ORGANIZATION

1V•"Approved for public releanel '0orSDistributii-n Unlimited •"

Published February 1993

ODstibution and AvaiabWty on Back Cover .

I)

AGARD-CP-527

ADVISORY GROUP FOR AEROSPACE RESEARCH & DEVELOPMENT7 RUE ANCELLE 92200 NEUILLY SUR SEINE FRANCE

AGARD CONFERENCE PROCEEDINGS 527 Accesion ForNTIS CRA&IHeat Transfer and Cooling DTIC TAB

Unannounced

in Gas Turbines(Le Transfert Thermique et le Refroidissement Bydans les Turbines b, Gaz) Bdans les Turb nes A Daz)-":' " "-.1 .... ....................................

Aaliiiiiy CocesI AvaCi a I.or

Dist

Papers presented at the Propulsion and Energetics Panel 80th Symposium heldin Antalya, Turkey 12th--16th October 1992.

- _ North Atlantic Treaty OrganizationOrganisation du TraitW de I'Atlantique Nord

93-0788892n 14 5

The Mission of AGARD

According to its Charter, the mission of AGARD is to bring together the leading personalities of the NATO nations in the fieldsof science and technology relating to aerospace for the following purposes:

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

- Providing scientific and technical advice and assistance to the Military Committee in the field of aerospace research anddevelopment (with particular regard to its military application);

- 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-

- Exchange of scientific and technical information;

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

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

The highest authority within AGARD is the National Delegates Board consisting of officially appointed senior representativesfrom each member nation. The mission of AGARD is carried out through the Panels which are composed of experts appointedby the National Delegates. the Consultant and Exchange Progranmme and the Aerospace Applications Studies Programme. Theresults of AGARD work are reported to the member nations and the NATO Authorities through the AGARD series ofpublications 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 reproduceddiectly from material supplied by AGARD or the authors.

Published February 1993

Copyright C AGARD 1993All Rights Reserved

ISBN 92-835-0701-0

HPnnted by Specialised Printing Services Limited40 Chigwell Lane, Loughton, Essex IGIO 377

Recent Publications ofthe Propulsion and Energetics Panel

CONFERENCE PROCEEDINGS (CP)

Heat Transfer and Cooling in Gas TurbinesAGARD CP 390, September 1985

Smokeless PropellantsAGARD CP 391, January 1986

Interior Ballistics of GunsAGARD CP 392, January 1986

Advanced Instrumentation for Aero Engine ComponentsAGARD CP 399, November 1986

Engine Response to Distorted Inflow ConditionsAGARD CP 400, March 1987

Transonic and Supersonic Phenomena in TurbomachinesAGARD CP 401 March 1987

Advanced Technology for Aero Engine ComponentsAGARD CP 421, September 1987

Combustion and Fuels in Gas Turbine EnginesAGARD CP 422, June 1988

Engine Condition Monitoring - Technology and ExperienceAGARD CP 448, October 1988

Application of Advanced Material for Turbomachinery and Rocket PropulsionAGARD CP 449, March 1989

Combustion Instabilities in Liquid-Fuelled Propulsion SystemsAGARD CP 450, April 1989

Aircraft Fire SafetyAGARD CP 467, October 1989

Unsteady Aerodynamic Phenomena in TurbomachinesAGARD CP 468. February 1990

Secondary Flows in TurbomachinesAGARD CP 469. February 1990

Hypersonic Combined Cycle PropulsionAGARD CP 479, December 1990

Low Temperature Environment Operations of Turboengines (Design and User's Problems)AGARD CP 480, May 1991

CFD Techniques for Propulsion ApplicationsAGARD CP 510, February 1992

Insensitive MunitionsAGARD CP 511. July 1992

Combat Aircraft NoiseAGARD CP 512, April 1992

Airbreathing Propulsion for Missiles and ProjectilesAGARD CP 526, September 1992

Heat Transfer and Cooling in Gas TurbinesAGARD CP 527. February 1993

- iii

ADVISORY REPORTS (AR)

Suitable Averaging Techniques in Non-Uniform Internal Flows (Results of Working Group 14)AGARD AR 182 (in English and French), June/August 1983

Producibility and Cost Studies of Aviation Kerosines (Results of Working Group 16)AGARD AR 227, June 1985

Performance of Rocket Motors with Metallized Propellants (Results of Working Group 17)AGARD AR 230, September 1986

Recommended Practices for Measurement of Gas Path Pressures and Temperatures for Performance Assessment ofAircraft Turbine Engines and Components (Results of Working Group 19)AGARD AR 245, June 1990

The Uniform Engine Test Programme (Results of Working Group 15)AGARD AR 248, February 1990

Test Cases for Computation of Internal Flows in Aero Engine Components (Results of Working Group 18)AGARD AR 275, July 1990

Test Cases for Engine Life Assessment Technology (Results of Working Group 20)AGARD AR 308, September 1992

Terminology and Assessment Methods of Solid Propellant Rocket Exhaust Signatures Results of Working Group 21)AGARD AR 287. February 1993

LECTURE SERIES (LS)

Ramjet and Ramrocket Propulsion Systems for MissilesAGARD LS 136, September 1984

3-D Computation Techniques Applied to Internal Flows in Propulsion SystemsAGARD LS 140, June 1985

Engine Airframe Integration for RotorcraftAGARD LS 148,June 1986

Design Methods Used in Solid Rocket MotorsAGARD LS 150, April 1987AGARD LS 150 (Revised), April 1988

Blading Design for Axial TurbomachinesAGARD LS 167, June 1989

Comparative Engine Performance MeasurementsAGARD LS 169, May 1990

Combustion of Solid PropellantsAGARD LS 180. July 1991

Steady and Transient Performance Prediction of Gas Turbine EnginesAGARD LS 183, May 1992

AGARDOGRAPHS (AG)

Measurement Uncertainty within the Uniform Engine Test ProgrammeAGARD AG 307, May 1989

Hazard Studies for Solid Propellant Rocket MotorsAGARD AG 316, September 1990

Advanced Methods for Cascade TestingAGARD AG 328 (to be published in 1993)

REPORTS (R)

Application of Modified Loss and Deviation Correlations to Transonic Axial CompressorsAGARD R 745, November 1987

Rotorcraft Drivetrain Life Safety and ReliabilityAGARD R 775, June 1990

iv

Theme

Heat transfer and cooling in gas turbine engines are still key factors to achieve high performance, increased life and improvedreliability. Any progress in this field will lead to a reduction of maintenance cost and fuel consumption.

The purpose of this Symposium was to bring together experts from industry, research establishments and universities to discussfundamental and applied heat transfer problems relevant to gas turbines, to exchange practical experience gained and to reviewthe state of the art.

The Symposium focused on turbine blade cooling (both external and internal heat transfer); heat transfer in combustors, todisks, in labyrinth seals, and in shafts; measurement techniques and prediction methods; as well as interactions.

Theme

Le transfert thermique et le refroidissement continuent a jouer un r6le cli dans l'obtention de meilleures performances.I'augmentation de la duree de vie et l'amdlioration de la fiabilite des turbines i gaz. Tout progres realise dans ce domainepermettra de reduire les cofits de maintenance et de diminuer la consommation de carburant.

L'objet du Symposium ýtait de rassembler des sp~cialistes de l'industrie. des etablissements de recherche et des universitis pourdiscuter des problimes fondamentaux et d'application en transfert thermique dans les turbines a gaz. La reunion a fournil'occasion pour un 6change d'exp~rience pratique et l'examen de l'itat de l'art dans ce domaine.

Le Symposium a traiti du refroidissement des aubes de turbine (le transfert thermique interne et externe), du transfertthermique dans les chambres de combustion, les disques, les presse-garnitures A labyrinthe et les arbres. ainsi que des mithodesde prevision et des techniques de mesure et les interactions qui en resultent.

v

Propulsion and Energetics Panel

Chairman: Prof. Dr Ahmet OCer Deputy Chairman: Mr Robert E. HendersonMiddle East Technical University Acting ChiefODTU Advanced Propulsion DivisionMakina Muh. B6oimii WL/POTAnkara Wright Patterson Air Force BaseTurkey Ohio 45433-6563

United States

PROGRAMME COMMITTEEProfessor Dr Dietmar K. Hennecke M. le Professeur Jacques Chauvin Ing. Claudio Vinci

(Chairman) Laboratoire d'Energetique et de FIAT Aviazione s.p.a.Fachgebiet Flugantriebe Micanique des Fluides ProgettazioneTechnische Hochschule Darmstadt Internes (LEMFI) Corso Ferrucci 112Petersenstrasse 30 Campus Universitaire 10138 Torino, ItalyW-6100 Darmstadt. Germany Bt 502

91405 Orsay Cedex, France Mr William W. WagnerDr Robert Bill Technical Director (Code 07)US Army Propulsion Directorate Mr David P. Kenny Naval Air Propulsion CenterNASA Lewis Research Center Director, Analytical Engineering P.O. Box 7176Mail Stop 77-12 Pratt and Whitney Canada, Inc. Trenton. New Jersey 08628-017621000 Brookpark Road 1000 Marie-Victorin United StatesCleveland, Ohio 44135 Longueuil. Quebec, CanadaUnited States Mr David Way

Professor Jose J. Salva Monfort Superintendent, TurbomachineryProfessor Frans Breugelmans Escuela Tecnica Superior de Propulsion DepartmentHead, Turbomachinery Dept, Ingenieros Aeronauticos Defence Research Agency

Assistant Director Plaza Cardenal Cisneros 3 (Aerospace Division) RAEvon Kirman Institute for 28040 Madrid. Spain Pyestock. Farnborough,

Fluid Dynamics Hants GU14 OLS72 Chaussee de Waterloo United Kingdom1640 Rhode St Gen•se, Belgium

HOST NATION COORDINATOR

Professor Dr Ahmet U3qer

PANEL EXECUTIVE OFFICEMail from Europe: Mail from US and Canada:AGARD-OTAN AGARD-NATOAttn: PEP Executive Attn: PEP Executive7, rue Ancelle Unit 21551F-92200 Neuilly-sur-Seine APO AE 09777France

Tel: 33(1)47 38 57 85Telex: 610176F

Telefax: 33 (1)47 38 57 99

ACKNOWLEDGEMENTThe Propulsion and Energetics Panel wishes to express its thanks to the National Authorities from Turkey for the invitation tohold this meeting in Antalya, Turkey, and for the facilities and personnel which made this meeting possible.

vi

Contents

Page

Recent Publications of PEP iii

Theme/Thime v

Propulsion and Energetics Panel A

Reference

Technical Evaluation Report Tby R.E. Mayle

Keynote Address: Unsteady, Multimode Transition in Gas Turbine Engines Kby R.E. Mayle

SESSION I - TURBINE BLADES: EXTERNAL HEAT TRANSFER

Heat Transfer and Aerodynamics of a 3D Design Nozzle Guide Vane Tested in the 1Pyestock Isentropic Light Piston Facility

by K.S. Chana

Vortex Structure and Mass Transfer Near the Base of a Cylinder and a Turbine Blade 2by M.Y. Jabbari and RJ. Goldstein

Thermal Effects of a Coolant Film along the Suction Side of a 3High Pressure Turbine Nozzle Guide Vane

by T. Arts and I. Lapidus

Etude Experimentale du Transfert de Chaleur pres d'une Paroi Plane Chauffee en 4Presence d'Injections Multiples

par E. Foucault, P. Deniboire, J.-L. Bousgarbies et JJ. Vullierme

The Influence of Non-Uniform Spanwise Inlet Temperature on 5Turbine Rotor Heat Transfer

by G.R. Guenette, G. Pappas and A.H. Epstein

Determination of Surface Heat Transfer and Film Cooling Effectiveness in 6Unsteady Wake Flow Conditions

by M. Sautner, S. Clouser and J.C. Han

Measurement of Turbulent Spots and Intermittency Modelling at Gas Turbine Conditions 7by J.P. Clark, J.E. LaGraff, PJ. Magari and T.V. Jones

Heat Transfer in High Turbulence Flows - a 2D Planar Wall Jet 8by R.B. Rivir, W.T. Troha, W.A. Eckerle and WJ. Schmoll

Heat Transfer with Moderate Free Stream Turbulence 9by S.N.B. Murthy

SESSION II - TURBINE BLADES: INTERNAL HEAT TRANSFER

Echanges Thenniques dana les Canaux de Refroidissement des Aubes de Turbine 10par Y. Servouze et A. Ristori

The Effect of Orthogonal-Mode Rotation on Forced Convection In a I ICircular-Sectioned Tube Fitted with Full Circumferential Transverse Ribs

by W.D. Morris and R. Salemi

vii

Reference

Turbulent Flow and Heat Transfer in Idealized Blade Cooling Passages 12by T. Bo and B.E. Launder

Cooling Geometry Optimization using Liquid Crystal Technique 13by G. Lodigiani, A. Trovati, L. Paci and P. Pirrelli

Fundamental Studies of Impingement Cooling Thermal Boundary Conditions 14by M.G. Lucas, P.T. Ireland, Z. Wang, T.V. Jones and W.J. Pearce

Prediction of Jet Impingement Cooling Scheme Characteristics 15*(Airfoil Leading Edge Application)

by A. Riahi, HJ. Saabas and W. Abdel-Messeh

SESSION III - MEASUREMENT TECHNIQUES

An L2F-Measurement Device with Image Rotator Prism for 16Flow Velocity Analysis in Rotating Coolant Channels

by M. Beversdorff, 0. Hein and R. Schodl

Local Heat Transfer Measurement with Liquid Crystals on 17Rotating Surfaces including Non-axisymmetric Cases

by D.E. Metzger and Y.K. Kim

Paper 18 withdrawn

The Swollen Polymer Technique and its Use for Heat Transfer Investigations on 19Film Cooled Surfaces

by N. Hay, D. Lampard and N. McLeod

The USAF Advanced Turbine Aerothermal Research Rig (ATARR) 20by C.W. Haldeman Jr, M.G. Dunn. C.D. MacArthur and C.G. Murawski

SESSION IV - ROTATING DISKS, LABYRINTH SEALS AND SHAFTS

Transient Thermal Behaviour of a Compressor Rotor with Axial Cooling Air Flow and 21Co-Rotating or Contra-Rotating Shaft

by C. Burkhardt, A. Mayer and E. Reile

Calculs Aerothermiques d'Ecoulements dans des Cavites 22Interdisques de Turbines

par D. Dutoya et Ph. Poncelin de Raucourt

Study of Flow Structure in Rotating Cavities: Numerical Predictions of 23Laminar and Turbulent, Steady and Unsteady Flows

by P. Maubert et al.

Modelling Thermal Behaviour of Turbomachinery Discs and Casings 24by R.D. Monico and J.W. Chew

Flow and Heat Transfer between Gas-Turbine Discs 25by X. Gan, M. Kilic and J.M. Owen

Heat Transfer and Leakage in High-Speed Rotating Stepped Labyrinth Seals 26by W. Waschka, S. Wittig, S. Kim and Th. Scherer

Fluid Flow and Heat Transfer in the Entrance Region of an Annulus between 27Independently Rotating Tubes with a Turbulent Axial Flow

by H. Pfitzer, T. Rothe and H. Beer

'Paper available and distributed but not presented at the Symposium.

viil

Reference

SESSION V - COMBUSTORS

Paper 28 withdrawn

The Effect of Main Stream Flow Angle on Flame Tube Film Cooling 29by H. Klinger and D.K. Hennecke

Impingement/Effusion Cooling 30by G.E. Andrews et al.

Evaluation of Multi-Dimensional Flux Models for Radiative Transfer in 31Cylindrical Combustion Chambers

by N. Selquk

Calculs d'Ecoulements Turbulents Reactifs dans les Foyers Aeronautiques 32par P. Caiilau et F. Dupoirieux

SESSION VI - DESIGN, INTERACTIONS

Aero-thermal Design of a Cooled Transonic NGV and 33Comparison with Experimental Results

by S. Colantuoni, A. Colella, L. Di Nola, D. Carbone and D. Marotta

Paper 34 withdrawn

The Aerodynamic Effect of Coolant Ejection in the Leading Edge Region of a 35Film-cooled Turbine Blade

by A. Beeck, L. Fottner, E. Benz and S. Wittig 36

On the Development of a Film Cooling Layerby F. L6pez Pefia and T. Arts

Paper 37 withdrawn

SESSION VII - PREDICTION METHODS

Paper 38 withdrawn

Modelisation d'un Ecoulement Turbulent en Presence de 39Jets Parietaux Discrets de Refroidissement

par J.M. Maurice, F. Leboeuf et P. Kulisa

Coupling of 3D-Navier-Stokes External Flow Calculations and Internal 3D-Heat 40Conduction Calculations for Cooled Turbine Blades

by A. Heselhaus, D.T. Vogel and H. Krain

A Navier Stokes Solver with Different Turbulent Models Applied to 41Film-Cooled Turbine Cascades

by F. Bassi, S. Rebay, M. Savini. S. Colantuoni and G. Santoriello

Navier-Stokes Analysis of Three-Dimensional Flow and 42Heat Transfer inside Turbine Blade Rows

by C. Hah

Cooling Predictions in Turbofan Engine Components 43by A. Matesanz, R. Rebolo, A. Viedma and M. Rodriguez

ix

43-1

COOLING PREDICTIONS IN TURBOFAN ENGINE COMPONENTS

A. MATESANZ, H. REBOLO, A. VIEDMA & M. RODRIGUEZ

SENER Ingenierfa y Sistemas. S.A.Parque Tecnoldgico de Madrid

28760 Tres Cantos, Madrid, Spain

(1) Also in Universidad Polit4cnica de MadridSchool of Aeronautical Engineering,

Pl. Cardenal Cisneros 3, 28040 Madrid, Spain

SUMMARY T2 = cooling injection temperature.

x = distance from the point of injection.The aim of this work is to show how the Ac= correcting factor due to spectralmetal temperature measured in a convergent- overlapping.divergent nozzle and in a turbine exhaust c = inner face wall emissivity.diffuser of a turbofan engine, can be gpredicted with reasonable approximation Cg= gas emissivity = Cc02Pc02+ cH2 OPH2 o-cusing the data available in the openliterature. It is shown how the simplified n = film cooling effectivenessfluid dynamic equations with the gas dynamic viscosityappropriate experimental correlation allow e = angle from stagnation point in profilethe prediction of these results in other leading edge.flight conditions than those tested. p = density

p"),.= density and viscosity at aLIST OF SYMBOLS temperature T*

S= Stefan-Boltzman constant.C = specific heat at main stream a

temperature. = non dimensional distanceCp2= specific heat at cooling flow -0,25

temperature. , [ R,2 2-Dh= Hydraulic diameter M.S [ p.

d = local curvature diameter Subscriptsh = convective heat flux coefficientK = thermal conductivity coefficient g = gasL = inner blade passage length w = wall

M =blowing factor P2 V2 1,2 = cold gas injection

PW VW i INTRODUCTIONNu= Nusselt number The modern engine solid surfaces exposed toPr= Prandtl number internal flow need improved thermalq = heat flux protection because high temperature cyclesRe 2 = Reynolds number for the cooling are used to increase performance. During

the past decades, different cooling methodsP2 V2 . S have been used to reduce the metal

parameters 2temperature, therefore minimizing theU2 required amount of cooling air.

R xS,d= Reynolds number based on xsd. The simplest way to cool surfaces is byconvective cooling. In this process, heat

s = cooling slot height. flows by conduction from exposed metalsurfaces to un-exposed ones, whic'z are

T,= absolute gas temperature. cooled by air flowing usually parallel toit. Convective cooling is used whenever low

T = recovery temperature = T++O,9rT-T levels of cooling effectiveness arer ele.Owj] required. This limitation exists due to the

TW= actual wall temperature fact that the air supply is somehowTaw = adiabatic wall temperature limited. On the other hand, high

effectiveness levels tend to increaseTOT = static and stagnation temperature thermal stress problems [1-3].

A special type of convective cooling is byof the main stream, means of impingement. It is used whenever

TM OT. FT 0,]2 large heat transfer coefficients are neededT 0 av %] on the un-exposed surfaces (4-8]. Another

43-2

method to obtain high heat transfer Xcoefficients is to place fins or ribsnormal to the coolant flow path [9].

When higher levels of cooling effectivenessare required it is necessary to use more "sophisticated alternatives. The most common T.method is to insert a secondary fluid into T0the boundary layer on the surface which isto be protected. There are different meansof injecting this fluid such as ablation, Fig.l. Single slot film coolingtranspiration and film cooling. In ablation configuration.cooling, a heat shield ablates andsecondary fluids enter the boundary layer[10] In transpiration cooling the coolant The convective coefficient h is calculatedenters the boundary layer through a porous independently from the adiabatic wallmaterial [11-13). Both methods are used to temperature Taw.protect the region where coolants areadded. Unfortunately the application ofthese systems is difficult because ablation This magnitude is related to the concept ofhas a limited time span and porous effectiveness of the film cooling, figurematerials are not strong enough to be used 1:in engines. The basic mechanism of film Tr T awcooling is the introduction of a secondary Tr Tfluid at several locations along a surface 2

to protect that surface not only in theinjection area but also in the downstream The correlation found more suitable for theregion a12-14]. flow conditions and geometry is the modelof Kutatelache & Leanter, in reference (14]

Thermal studies on aircraft engines, because the flow can be consideredparticularly modern ones with large flight twodimensional and compressible and theenvelopes, require a great amount of difference of temperature between the walltesting and instrumentation. Designers have and the main stream is important.different tools for obtaining enginecomponent temperature predictions. These The effectiveness is given from thistools can be classified as: 1) theoretical, reference by the expression:2) experimental and 3) numerical methods.Theoretical analysis is usually aimed atfinding parameters to describe the =temperature of an adiabatic wall downstream + [ (40 + *)of the coolant injection (15]. Real and + p2three-dimensional effects are studied on p2test rigs with controlled conditions When the cooling injection is made by two[16-18). Predictive numerical calculations consecutive slots, the total effectivenessare being widely used nowadays for can be treated as the combination of twodifferent cooling applications and cooling layers where the externalconfigurations, because they help to speed temperature for the inner flow is theup the design and to reduce the number of "adiabatic wall temperature" for theexperiments (19-20]. external one [21], figure 2.In this paper a method to predict full T -scale engine temperatures is described. TawlThis method combines the fluid dynamic =T - TIequations with semi-empirical data obtainedfrom the open literature. TawI- Taw2The result is an easy to use procedure. 72 =Tawl - T2

2. NOZZLE COOLING PREDICTION

The prediction of the petal temperature of and the total effectiveness is:

a convergent-divergent nozzle has been Tw - T a 2performed through a computer code that uses = T - T + n2 (1 -correlation models available in the open 2literature with actual data obtained from atest program that includes a reduced scalehot test, a full scale rig test and a DVE(Design Verification Engine) program. T, ...

2

The film cooling produced when the relative T,cold layer is injected at the beginning of _Tathe convergent petal of the nozzle, Twproduces a reduction in the metal To

temperature along the petal.

The heat transfer rate is modelled by: TO

q - A.h [Taw - Tw] Fig.2. Double slot film coolingconfiguration.

43-3

The convective coefficient h is obtained Tw/Treffrom semiempirical correlation for filmcooling with a high velocity compressible 2.20boundary layer. s

As the blowing factor M < 2 the correlationused is the one proposed by Lefebre (22].

K2

x S 0,01 m hr = 2uK 1.80

0,7Nux = 0,057 R ex

R P2 V 2 .X

ex= 2 1.40

x > 0.01 m hy = Ux 2

N =0,0256 es H.S 1.0e

0 0.4 0.8 1.2

R = V RelatiVe X Dositiones U2 Fig.4. Comparison of predicted temperaturesand fixed condi nozzle test results.

The wall temperature of the petal is

obtained for an equilibrium model where theconvective and radiative heat transfer Tw/Trefwith the cavity between the nozzle petalsand the fairing flap, and the conductionalong the petal are modeled in theclassical way. 2.6

It may be of interest to note that theradiative heat transfer with the hot gas ofthe main flow is modeled as (15 & 23).

q 1 ' 1,5(2,5 2,5) 2.2-

A 2 a +C )J Cg Tg (Tg' T

3. NOESLE THERNRL TESTS

The model resulting from this information 1.8gave good results for wall temperature Prctedestimation always with a conservative + Teouplemargin that the test in the scale modelhelped to reduce (figure 3). The full scaleresults with fixed convergent-divergent 1.4.nozzle tested in a hot rig showed very goodagreement between the predictions and the 0 0.4 0.8 1.2actual values from the thermocouples and Relative X positionthermal paints. (see figures 4 and 5). Fig.5. Thermal points, predicted

2.00 - temperature and thermocouples values infixed condi nozzle.

1.80When the true variable convergent-divergentnozzle was tested in the Design

1.60 " + Verification Engine (DVE) the differencesbetween the prediction and the measurementin the beginning of the divergent petal

1.40 were important. But if a cooling ingestion+ through the gap in the throat is

Cat~ considered, this disagreement disappears1.20 -and the model appears to be consistent and

to give accurate temperature predictions.(Figure 6).

1.000 0.2 0.4 0.6 0.0:8 4. * ZAUST DIfFUSER COOLING CIRCUIT

Rab XPOS~n The cooling circuit of the exhaust diffuser

Fig.3. Initial prediction and scaled hot has been designed with a method thattest results along nozzle petal. modelise the more relevant fluid factors

43-4

Tw/Tref which receives strong radiative heat fluxfrom the afterburner. The dominant feature

2.4 in this part of the cooling circuit is the2.4! combination of cross-flow and impingement

cooling. The radial geometry simplifies theW/o -Inct iecta mathematical model which follows the scheme

of (26].

* The calculation of the flow through the

0 holes in each station is also modified toI0 allow for the cross flow. Correlation from

(27] is used to take this effect intowits• ?i~ct l o account. The convective heat coefficient

I Ieproduced by the impingement is modelledI following (4 & 28].

To reproduce this correlation here isbeyond our limits, so our advice is toconsult the original papers.

e- r.P'rde s Wall temperature is obtained by a thermal* Therii•ecouoIes balance between external and internal

radiative heat flux and the impingement asalready commented. Upon the reheat luminous1.2• flame emissivity and radiative terms of

0 0.4 0.8 1.2 this equilibrium are modelled after

Relative X position references [3 & 15).

Fig.6. Predicted and measured temperatures The flow behavior inside the lateral doublewith and without throat cold air injection, wall is calculated in a similar way with

the simplification that no impingementeffect is present.

around the circuit and use the availabledata in the open literature. This The mass flow and thermal predictions areprediction method includes theoretical performed jointly in an iterative mode. Itanalysis and semienpirical correlations starts with a hypothetical mass flow at thecoupled to numerical simulation wherever beginning of the circuit. Its' purposefurther refinement is required. A more being to match the pressure, at the exitdetailed description of this design method located downstream of the cascade, obtainedhas been given elsewhere [24]. Therefore by the program with the actual value.here we will only comment at the asoectsrelated to the general objectives of thispaper.

The geometry of the cooling circuit can be Vseen in figure 7, the fixed blade diffuser 1cascade has an inner passage with coolingcapillaries in the trailing edge. Thethermal analysis of the blades requires adetailed numerical process with NASTRANcode for the solid pieces and CFD code forthe convective heat transfer in the bladeboundary layer. But for the purpose of thecircuit flow prediction a simpler model isused estimating the internal frictioncoefficient using (25] and the inner heatflux with the expression from (2].

Nu = " - -36 Re ' P

V

The mean blade temperature is fixed by aheat equilibrium analysis where theexternal heat transfer is modelled in the Fig.7. Exhaust diffuser cooling circuit.leading edge with the correlation

Nu=1.14 Re 11 [ ,41 -!I-J 5. EXHAUST DIFFUSER TEMPERATURE ASSESSMENT

Calibration tests were carried out to checkand in the rest of the blade surface with the mass flow predicted through the coolingthe standard Nusselt value for turbulent circuit. It was performed in a rig thatboundary layer. reproduces the geometry and flow

conditions, and is divided into two phasesThe remaining heated flow which reaches the to obtain separately the mass flow throughplenum is forced to cross the perforated out both capillary trailing edge tubes andwall to impinge onto the rear wall. This is discharge exit downstream blade trailingneeded to enhance the cooling of this wall edge.

43-5

Figures (8 and 9) show the comparison4.4- between experiments and theoretical

predictions for the capillary tubes and the7 1 total cooling mass flow.

4.0 z W

The engine validation tests were performed- to assess the pressure and temperature

3.6-1 predictions. We will focus on the cone,

U9 lateral and rear walls where no furthercalculation was performed. The figures

0 3.2-, (10,11 and 12) show the comparison ofa!_______ predicted and measured temperature in the

¶ range of compressor speed, while figure

2.8 [13] shows the relative error for both.8~ temperature and pressure in several circuitlocations.

2.4

I .1 1.3 1.5 1.7 0.901PRESSUIRE RA 770 0nez

Fig.8. Results of calibration test mass 0.80____f low through capillary cooling tubes in the 0.0-blade trailing edge.

4.8 __________________ _ 0. 702

4.4 0.60-.

4.0S ~0.0 so -

3.6240 60 8010

3. Fig.ll. Comparison between predicted andexperimental values of temperature at the

2.8 --. ICat periphery of the diffuser rear wall.

2.41.1 1.3 1.5 1.7 0.901,

PRESSURERA4 770

Fig.9. Results of calibration test. Total Caclaecooling mass flow in exhaust diffuser 0.80-1cooling circuit.

0.70,

1.00 1 .

-L~nmwikI0. 60¶

0.00)-oj ~ M

/~u~ 0.50

I,40 60 80 1 0070~1 RW LP cai-ipwsa Spee (V

Fig.12. Comparison between predicted and

I I experimental values of temperature at an0.60 intermediate section of the lateral wall.

0.50 f 6. CONCLUSIONS40 60 80100

fta"LP a"Pl" S (WThe comparison between predicted and actualtest results of calibration rigs and engine

Fig.10. Comparison between predicted and test beds shows that for design purposesexperimental values of temperature at the the accuracy obtained is enough and nocenter of the diffuser rear wall. other complicated method is needed.

43-6

DEPAM1$SFROMMALEAWD VALUES 1%) 8. ANDREWS, G.E. et al. "Full CoverageImpingement Heat Transfer: The Variation in

3.0 •Pitch to Diameter Ratio at a Constant Gap",

University of Leeds, Leeds, U.K.

2.0 9. WEBB, R.L., ECKERT, E.R.G. andGOLDSTEIN, R.J. "Heat Transfer and Frictionin Tubes with Repeated-Rib Roughness". Int.

1.0 / J eat Mass Transfer, vol. 14, pp. 601-617,Y///I 19 71.

0.0 10. LAUB, B. "Thermochemical Ablation ofT.antalum Carbide Loaded Carbon-Carbons.Paper 80-0290 AIAA 18+h Aerospace Sciences

-1.0- P2 MxMeeting. 1980.

P,42 111. LIBRIZZI, J. and CRESCI, R.J.

M2.2 "Transpiration Cooling of a TurbulentT Boundary Layer in a Axisymmetric Nozzle".

t43 AIAA Journal, vol. 2, April 1964, pp.

-3.01 617-624.

P1 P2 TI T2 11 12 M3 12. MOFFAT, R.J. and KAYS, W.M. "A Reviewof Turbulent-Boundary-Layer Heat Transfer

Fig.13. Differences between prediction and Research at Stanford, 1958-1983". Advancesmeasured values. P, T and M are pressure, in Heat Transfer, vol. 16, Academic Press,air temperature and metal temperature Orlando 1984, p. 241.respectively.

13. HARTNETT, J.P. "Mass Transfer Cooling",Handbook of Heat Transfer Applications, pp.

Relevant engineering design parameters, 11, 1111.

such as metal temperatures, pressures andmass flow were predicted within 3% of 14. GOLDSTEIN, R.J. "Fluid Cooling".accuracy. It is possible to explore a broad Advances in Heat Transfer, Academic Press,range of flight envelope points using this 1971.

simple design tool. Further improvements onthe modeling of reheat radiative properties 15. LEFEBVRE, A.H. "Gas Turbinewould be needed if greater pcurowertes Combustion", McGraw-Hill, New York, pp.woldbenedd fgratraccuracy were 2725required. 287-295.

REFERENCES 16. LUCAS, G.J. and GOLLADAY, R.L."Gaseous-Film Cooling of a Rocket Motor

1. ECKERT, E.R.G. and DRAKE, M.M. Jr, "Heat with Injection Near the Throat". NASA TNand Mass Transfer", New York, USA, D-3836, 1967.McGraw-Hill, 1959, pp. 167-253. 17. STEPHEN DAPELL, S. "Effect on Gaseous

2. NECATI OZISIK, M. "Heat Transfer. A Film Cooling of Coolant Injection throughBasic Approach", New York, USA, Angled slots and Normal Holes". NASA TNMcGraw-Hill, pp. 281-415. D-299, 1960.

3. CHAPMAN, A.J., "Heat Transfer", New 18. SMAO-YEN KO and DENG-YING LIU,

York, USA, McMillan, 1984. "Experimental Investigations onEffectiveness, Heat Transfer Coefficient,

4. KERCHER, D.M. and TABAKOFF, W. "Heat and Turbulence of Film Cooling", AIAATransfer by a Square Array of Round Air Journal, vol. 18, August 1980, pp. 907-913.Jest Impinging Perpendicular to a FlatSurface Including the Effect of Spent Air". 19. DIBELIOS, G.H. et al. "Numerical

ASME I. of Eng. for Power, Jan 1970. Predictions of Film Cooling Effectivenessand the Associated Aerodynamics Losses with

5. OBOT, N.T. and TRABOLD, T.A. a Three-Dimensional Calculation Procedure".

"Impingement Heat Transfer within Arrays of ASME 90-GT-226.Circular Jets: Parts I and II".Transactions of the ASME, vol. 109, 1987. 20. STOLL, J. and STRAUB, J. "Film cooling

and Heat Transfer in Nozzles". Journal of

6. DABAGH, A.M. et al. Turbomachinery, January 1988, vol. 110.

"Impingement/Effusion Cooling: TheInfluence of the Number of Impingement 21. SELLORS, J.P., "Gaseous film cooling

Holes and Pressure Loss on the Heat with multiple injection stations", AIAA J.,

Transfer Coefficient", Gas Turbine and vol. 1, Sept. 1963, pp. 2154-2156.

Aeroengine Congress and Exposition, 1989, , "A89-GT-188. ' 22. LEFEBVRE, , Aproposed method forcalculating film cooled wall temperatures

7. ABDUL HUSAIN, R.A.A. and ANDREWS, G.E. in gas turbine combustion chambers". A.H.,

"Full Coverage Impingement Heat Transfer at ASME 72-WA/HT-24High Temperatures". Gas Turbine andAeroengine Congress and Exposition, 1990, 23. HOLMAN, J.P., "Heat Transfer",

90-GT-285. McGraw-Hill.

43-7

24. MATESANZ, A., VIEDMA, A., VELAZQUEZ, A. Ronald Press. Co., New York, 1953, vol. I,& RODRIGUEZ, M., "Cooling study in the p. 228.exhaust diffuser of a reheat turbofan",International Symposium on Heat Transfer in 27. FLORSCHUETZ, C.W. & ISODA, Y., "FlowTurbomachinery, Athens, 24-28 August 1992. distribution and discharge coefficient

effects for jet away impingement with25. HACLAND, S.E., "Simple and explicit initial cross flow", J. Eng. for Power,formulas for the friction factor in vol. 105, pp. 296-304, 1983.turbulent pipe flow", J. Fluid Eng., vol.105, pp. 89-90, 1983. 28. FLORSCHUETZ, L.W. & S.U, C.C., "Effects

of cross flow temperature on heat transfer26. SHAPIRO, A.H., "The dynamics and within an array of impinging jets", Trans.thermodynamics of compressible fluid flow". ASME, vol. 109, Feb. 1987.

Discussion

QUESTION 1:DISCUSSOR: J. Salva Monfort, Escuela Tecnica Superior de Ingenieros Aeronauticos

Can you explain in more detail how you have calculated the cooling flux through thecapillary tubes in the trailing edges of the blades?

AUTHOR'S REPLY:The flow through the cooling capillary tubes in the trailing edge of the blades is calculatedusing the same turbulent one-dimensional equations as for the main passages of theblade. Some problems of choking due to the high heat transfer can arise. In this case, it isnecessary to detect the Mach number increase and to reduce the mass flow for thatcapillary tube accordingly.

QUESTION 2:DISCUSSOR: D.T. Vogel, DLR

You used many of empirical constants in your calculations. Are these constants relatedto your special problem, or is it possible to calculate other cooling configurations'?

AUTHOR'S REPLY:The correlations used are all chosen from open literature taking into account thegeometry and dimensionless parameters of the problem. Please refer to the originalpaper to find out if it can be applied to other configurations.

REPORT DOCUMENTATION PAGE

I!. Recipient's Reference 2. Originator's Reference 3. Further Reference 4. Security Classificationof Document

AGARD-CP-527 ISBN 92-835-0701_0 UNCLASSIFIED/UNLIMITED

5. Originator Advisory Group for Aerospace Research and DevelopmentNorth Atlantic Treaty Organization7 Rue Ancelle, 92200 Neuilly sur Seine, France

6. TitleHEAT TRANSFER AND COOLING IN GAS TURBINES

7. Presented at the Propulsion and Energetics Panel 80th Symposium held inAntalya, Turkey, 12th-- I6th October 1992.

8. Author(s)/Editor(s) 9. Date

Various February 1993

10. Author's/Editor's Address 11. Pages

Various 510

12. Distribution Statement There are no restrictions on the distribution of this document.Information about the availability of this and other AGARDunclassified publications is given on the back cover.

13. Keywords/DescriptorsCombustors Impingement coolingConvective heat transfer Rotating discsCooling Seals technologyFilm cooling Turbine bladesGas turbines Turbulent flowHeat transfer Wake flow

14. Abstract

The Conference Proceedings contains the 38 papers presented at the Propulsion and EnergeticsPanel 80th Symposium on "Heat Transfer and Cooling in Gas Turbines" which was held from12th-i16th October 1992 in Antalya. Turkey. The Technical Evaluation Report and the KeynoteAddress are included at the beginning, and discussions follow most papers.

The Symposium was arranged in the following Sessions: Turbine Blades: External Heat Transfer(9); Turbine Blades: Internal Heat Transfer (6); Measurement Techniques (4); Rotating Disks,Labyrinth Seals and Shafts (7); Combustors (4); Design, Interactions (3); and Prediction Methods(5).

Heat transfer and cooling in gas turbines are still key factors for achieving high performance,increased life and improved reliability. Any progress in this field will lead to a reduction ofmaintenance cost and fuel consumption. The purpose of the Symposium was to bring togetherexperts from industry, research establishments and universities to discuss fundamental andapplied heat transfer problems relevant to gas turbines, to exchange practical experience gainedand to review the state of the art.

04 0

0< ODE

A . cO

E 8 E tý -r s E2.c C 0 0 0 q 0-

A . *. 2 M C4 0

m(0 .= 0 o'0

Ru z .5 t

~~~j(0 0J Cr-0 o

u 0 - m-

C t~ o ~ D. u A (

< C <

r-.,r-ci 4d

UOU

uC >

~Q .~ CC

0 00 "U Z 0 0

a Eta 00 u to.~ Z

0; -- Q

0 -E

C0 00 CU om :E 02v

C4~~0 - - -0 0 N0 (S o 9 0, - '

U< P< ~0 'r-

QO ~ w~~A

~ ~ ~JDu~ Uv~ U, ~ U Iz

i.79

- C. -

ca~~-!cm~

SE

0.0 b> OI.- C;-0IEWob

ccQ 1

= 0

030

z ~ z

CL' 0

E '=

20 20

c~cc

~0 0

\-0 E~ - a~

E~ .s~

ILI . f

Uý 0- 0 ~rS eq C4

tj z o