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`GE v. UTC
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`GE-1032.001
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`GAS TURBINE
`HEAT TRANSFER
`AND COOLING
`TECHNOLOGY
`
`GE v. UTC
`IPR2016-01289
`GE-1032.003
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`

`

`ABOUT THE AUTHORS
`
`Je-Chin Han received his B.S. degree from National
`Taiwan University in 1970 and Sc.D. degree from M.I.T.
`in 1976, both in Mechanical Engineering. He joined Texas
`A&M University in 1980 and has been doing research on
`gas turbine heat transfer and cooling technology for air-
`craft propulsion and in land-based power generation or
`industrial applications for more than 25 years.
`
`Sandip Dutta received his B.S. degree from Indian
`Institute of Technology and Ph.D. from Texas A&M
`University in 1995, both in Mechanical Engineering. His
`research areas include internal cooling of gas turbine
`blades.
`
`¯ Srinath Ekkad received his B.S. degree from Jawahaflal
`Nehru Technological University and Ph.D. degree from
`Texas A&M University in 1995, both in Mechanical
`Engineering. He was a member of the Hot Gas Path
`Component Design Group at Rolls-Royce Allison. His
`research focuses on gas turbine heat transfer.
`
`GE v. UTC
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`GAS TURBINE
`HEAT TRAN SFER
`AND COOLING
`TECHNOLOGY
`
`Je-Chin Han, Sc.D.
`HTRI Professor, Texas A&M University,
`Department of Mechanical Engineering,
`College station, Texas
`
`Sandip Dutta, Ph.D.
`Assistant Professor, University of South Carolina,
`Department of Mechanical Engineering,
`Columbia, South Carolina
`
`Srinath Ekkad, Ph.D.
`Assistant Professor, Louisiana State University,
`Depa~ta-~ent of Mechanical Engineering,
`Batori Rouge, Louisiana
`
`GE v. UTC
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`Published in 2000 by
`
`Taylor & Francis
`29 West 35th Street
`New York, NY 10001
`
`Published in Great Britain by
`
`TayIor & Francis
`11 New Fetter Lane
`London EC4P 4EE
`
`Copyright © 2000 by Taylor & Francis
`
`Printed in the United States of America on acid-free paper.
`
`All rights reserved. No part of this book may be reprinted or repro-
`duced or utilized in any form or by any electronic, mechanical, or
`
`other means, now known or hereafter invented, including.photocopy-
`ing and recording, or in any information storage or retrieval system,
`without permission in writing from the publisher.
`
`Library of Congress Cataloging-in-Publication Data
`Han, Je-Chin, 1946-
`Gas turbine heat transfer and cooling technology / je-Chin Han,
`Sandip Dutta, and Srinath Ekkad.
`p. cm.
`Includes bibliographical references (p.).
`ISBN 1-56032-841-X (alk. paper)
`1. Gas-turbines. 2. Heat--Transmission. 3. Gas-turbinesJ
`I. Durra, Sandip, 1963- . II. Ekkad, Srinath, 1968- .
`Cooling.
`IIl. Title.
`TJ778.H24 1999
`621.43’3--dc21 99-26694
`CIP
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`Contents
`
`Preface
`
`CHAPTER 1
`FUNDAMENTALS
`1.1 NEED FOR TURBINE BLADE COOLING
`
`1.I.1 Recent Development in Aircraft Engines
`
`1.1.2 Recent Development in Land-Based Gas Turbines
`
`1,2 TURBINE-COOLING TECHNOLOGY
`
`1.2.1 Concept of Turbine Blade Cooling
`
`1.2.2 Typical Turbine Cooling System
`1.3 TURBINE HEAT-TRANSFER AND COOLING ISSUES
`
`1.3.1 Turbine Blade Heat Transfer
`
`1.3.2 Turbine Blade Internal Cooling
`
`1.3.3 Turbine Blade Film Cooling
`
`1.3.4 Thermal Barrier Coating and Heat Transfer
`
`1.4 STRUCTURE OF THE BOOK
`1.5 REVIEW ARTICLES AND BOOK CHAPTERS ON TURBINE
`COOLING AND HEAT TRANSFER
`
`CHAPTER 2
`TURBINE HEAT TRANSFER
`2.1 INTRODUCTION
`
`2.1.1 Combustor Outlet Velocity and Temperature Profiles
`
`2.2 TURBINE-STAGE HEAT TRANSFER
`2.2.1
`Introduction
`
`2.2.2
`
`2.2.3
`
`2.2.4
`
`Real Engine Turbine Stage
`
`Simulated Turbine Stage
`
`Time-Resolved Heat-Transfer Measurements
`on a Rotor Blade
`
`XV
`
`1
`1
`
`1
`
`4
`
`5
`
`5
`
`7
`
`13
`
`13
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`17
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`21
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`21
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`22
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`23
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`27
`27
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`27
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`31
`31
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`32
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`38
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`vi
`
`CONTENTS
`
`2.3 CASCADE VANE HEAT-TRANSFER EXPERIMENTS
`
`2,3.1
`
`Introduction
`2.3.2 Effect of Exit Mach Number and Reynolds Number
`
`2.3.3 Effect of Free-Stream Turbulence
`
`2.3.4 Effect of Surface Roughness
`
`2.3.5 Annular Cascade Vane Heat Transfer
`2,4 CASCADE BLADE HEAT TRANSFER
`
`2.4.1
`
`Introduction
`2.4.2 Unsteady WaRe-Simulation Experiments
`2.4.3 WaRe-Affected Heat-Transfer Predictions
`2.4.4 Combined Effects of Unsteady-WaRe and
`Free-Stream Turbulence
`2.5 AIRFOIL ENDWALL HEAT TRANSFER
`
`2.5.1
`
`Introduction
`
`2.5.2 Description of the Flow Field
`
`2.5.3 Endwall Heat Transfer
`
`2.5.4 Near-Endwall Heat Transfer
`2.5.5 Engine Condition Experiments
`2.5.6 Effect of Surface Roughness
`2,6 TURBINE ROTOR BLADE TIP HEAT TRANSFER
`
`2.6.1
`
`Introduction
`
`2.6.2 Blade Tip Region Flow Field and Heat Transfer
`2.6.3 Flat-Blade Tip Heat Transfer
`2.6.4 Squealer- or Grooved-Blade Tip Heat Transfer
`2,7 LEADING-EDGE REGION HEAT TRANSFER
`
`2.7,1
`
`Introduction
`2.7.2 Effect of Free-Stream Turbulence
`2.7.3 Effect of Leading-Edge Shape
`2.7.4 Effect of Unsteady WaRe
`2,8 FLAT-SURFACE HEAT TRANSFER
`2.8. I
`
`Introduction
`2.8.2 Effect of Free-Stream Turbulence
`2.8.3 Effect of Pressure Gradient
`
`2.8.4 Effect of Streamwise Curvature
`
`2.8.5 Surface Roughness Effects
`2.9 CLOSURE
`
`47
`
`47
`
`48
`
`52
`
`54
`
`57
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`60
`
`60
`
`62
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`69
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`73
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`77
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`77
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`77
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`79
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`83
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`83
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`85
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`88
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`88
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`88
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`92
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`93
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`100
`
`IO0
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`101
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`106
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`107
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`110
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`110
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`111
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`115
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`116
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`117
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`CONTENTS
`
`CHAPTER 3
`
`TURBINE FILM COOLING
`INTRODUCTION
`3.1
`
`3.1.I Fundamentals of Film Cooling
`3.2 FILM COOLING ON ROTATING TURBINE BLADES
`3.3 FILM COOLING ON CASCADE VANE SIMULATIONS
`Introduction
`3.3.1
`
`3.3.2 Effect of Film Cooling
`
`3.3.3 Effect of Free-Stream Turbulence
`
`3.4 FILM COOLING ON CASCADE BLADE SIMULATIONS
`3.4.1
`Introduction
`
`3.4.2 Effect of Film Cooling
`
`3.4.3 Effect of Free-Stream Turbulence
`
`3.4.4 Effect of Unsteady Wake
`
`3.4.5 Combined Effect of Free-Stream Turbulence
`and Unsteady Wakes
`3.5 FILM COOLING ON AIRFOIL ENDWALLS
`
`3.5.1
`
`Introduction
`
`3.5.2 Low-Speed Simulation Experiments
`
`3.5.3 Engine Condition Experiments
`
`3.5.4 Near-Endwall Film Cooling
`3.6 TURBINE BLADE TIP FILM COOLING
`
`3.6.1
`
`Introduction
`
`3.6.2 Heat-Transfer Coefficient
`
`3.6.3 Film Effectiveness
`3.7 LEADING-EDGE REGION FILM COOLING
`
`3.7.1
`
`Introduction
`
`3.7.2 Effect of Coolant-to-Mainstream Blowing Ratio
`
`3.7.3 Effect of Free-Stream Turbulence
`
`3.7.4 Effect of Unsteady Wake
`
`3.7.5 Effect of Coolant-to-Mainstream Density Ratio
`
`3.7.6 Effect of Film Hole Geometry
`3.7.7 Effect of Leading-Edge Shape
`3.8 FLAT-SURFACE FILM COOLING
`
`3.8.1 Introduction
`
`3.8.2 Film-Cooled, Heat-Transfer Coefficient
`
`vii
`
`129
`129
`
`129
`
`133
`
`14o
`
`140
`
`140
`
`15o
`
`151
`
`151
`
`1.51
`
`154
`
`156
`
`162
`
`162
`
`162
`
`162
`
`169
`
`171
`
`173
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`173
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`175
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`175
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`179
`
`179
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`18o
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`183
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`186
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`186
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`192
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`193
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`194
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`194
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`195
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`viii
`
`CONTENTS
`
`Effect of Blowing Ratio
`
`Effect of Coolant-to-Mainstream
`Density Ratio
`
`Effect of Mainstream Acceleration
`
`Effect of Hole Geometry
`
`3.8.3 Film-Cooling Effectiveness
`Effect of Blowing Ratio
`
`3.8.3.1
`
`3.8.3.2
`
`3.8.3.3
`
`3.8.3.4
`
`3.8.3.5
`
`3.8.3.6
`
`Effect of Coolant-to-Mainstream
`Density Ratio
`
`Film Effectiveness Correlations
`
`Effect of Streamwise Curvature
`and Pressure Gradient
`
`Effect of High Free-Stream Turbulence
`
`Effect of Film Hole Geometry
`
`3.8.3.7
`
`Effect of Coolant Supply Geometry
`
`3.8.3.8
`
`Effect of Surface Roughness
`
`3.8.3.9
`
`Effect of Gap Leakage
`
`3.8.3.10 Effect of Bulk Flow Pulsations
`
`3.8.3.11
`
`Full-Coverage Film Cooling
`
`3.9
`
`DISCHARGE COEFFICIENTS OF TURBINE
`
`COOLING HOLES
`
`3.10
`
`3.11
`
`FILM-COOLING EFFECTS ON AERODYNAMIC LOSSES
`
`CLOSURE
`
`CHAPTER 4
`
`TURBINE INTERNAL COOLING
`
`4.1
`
`JET IMPINGEMENT COOLING
`Introduction
`
`4.1.1
`
`4.1.2 Heat-Transfer Enhancement by a Single Jet
`
`4.1.2.1
`
`4.1.2.2
`
`Effect of Jet-to-Target-Plate Spacing
`
`Correlation for Single Jet
`Impingement Heat Transfer
`
`4.1.2.3
`
`Effectiveness of Impinging Jets
`
`4.1.3
`
`4.1.2.4 Comparison of circular to Slot Jets
`Impingement Heat Transfer in the Midchord
`Region by Jet Array
`
`196
`
`197
`
`198
`
`201
`
`206
`
`208
`
`209
`
`210
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`215
`
`220
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`223
`
`225
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`228
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`230
`
`234
`
`235
`
`235
`
`239
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`243
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`251
`251
`251
`251
`254
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`254
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`256
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`256
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`CONTENTS ix
`
`Jets with Large Jet-to-Jet Spacing
`4.1.3.1
`4.1.3.2 Effect of Wall-to-Jet-Array Spacing
`4.1.3.3 Cross-Flow Effect and
`Heat-Transfer Correlation
`
`4.1.3.4 Effect of Initial Cross-Flow
`
`4.1.3.5 Effect ofCross-Flow Direction on
`Impingement Heat Transfer
`
`4.1.3.6 Effect of Coolant Extraction on
`Impingement Heat Transfer
`
`4.1.3.7 Effect of Inclined Jets on Heat Transfer
`
`4.1.4
`
`Impingement Cooling of the Leading Edge
`
`4.1.4.1 Impingement on a Curved Surface
`Impingement Heat Transfer in the
`Leading Edge
`
`4.1.4.2
`
`4.2
`
`RIB.TURBULATED COOLING
`Introduction
`
`4.2.1
`
`4.2.1.1 Typical Test Facility
`
`4.2.2
`
`Effects of Rib Layouts and Flow Parameters
`
`on Ribbed-Channel Heat Transfer
`
`4.2.2.1 Effect of Rib Spacing on the
`Ribbed and Adjacent Smooth Sidewalls
`
`4.2.2.2 Angled Ribs
`4.2.2.3 Effect of Channel Aspect Ratio
`with Angled Ribs
`
`4.2.2.4 Comparison of Different Angled Ribs
`
`4.2.3 Heat-Transfer Coefficient and Friction
`Factor Correlation
`
`4.2.4 High-Performance Ribs
`4.2.4.1 V-Shaped Rib
`4.2.4.2 V-Shaped Broken Rib
`4.2.4.3 Wedge- and Delta-Shaped Rib
`Effect of Surface-Heating Condition
`
`4.2.5
`4.2.6 Nonrectangular Cross Section Channels
`Effect of High-Blockage-Ratio Ribs
`4.2.7
`
`4.2.8
`
`Effect of Rib Profile
`
`Effect of Number of Ribbed Walls
`
`4.2.9
`4.2.10 Effect of a 180° Sharp Turn
`
`259
`
`260
`
`260
`
`266
`
`267
`
`270
`
`276
`
`278
`
`278
`
`279
`
`287
`
`287
`
`290
`
`291
`
`291
`
`292
`
`295
`
`296
`
`297
`
`301~
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`301
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`304
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`306
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`309
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`313
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`325
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`327
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`333
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`X
`
`CONTENTS
`
`4.2.11
`
`4.2.12
`
`Detailed Heat-Transfer Coefficient
`Measurements in a Ribbed Channel
`
`Effect of Film-Cooling Hole on
`Ribbed-Channel Heat Transfer
`
`4.3
`
`PIN -FIN
`
`COOLING
`
`4.3.1
`
`4.3.2
`
`4.3.3
`
`4.3.4
`
`4.3.5
`
`4.3.6
`
`4.3.7
`
`4.3.8
`
`4.3.9
`
`Introduction
`
`Flow and Heat-Transfer Analysis with Single Pin
`
`Pin Array and Correlation
`
`Effect of Pin Shape on Heat Transfer
`
`Effect of Nonuniform Array and Flow Convergence
`Effect of Skewed Pin Array
`
`Partial Pin Arrangements
`
`Effect of Turning Flow
`
`Pin-Fin Cooling with Ejection
`
`4.3.10
`
`Effect of Missing Pin on Heat-Transfer Coefficient
`
`4.4 COMPOUND AND NEW COOLING TECHNIQUES
`Introduction
`
`4.4.1
`
`4.4.2
`
`4.4.3
`
`4.4.4
`
`4.4.5
`
`4.4.6
`
`4.4.7
`
`4.4.8
`
`4.4.9
`
`Impingement on Ribbed Walls
`
`Impingement on Pinned and Dimpled Walls
`
`Combined Effect of Ribbed Wall with Grooves
`
`Combined Effect of Ribbed Wall with Pins
`and Impingement Inlet Conditions
`
`Combined Effect of Swirl Flow and Ribs
`
`Impingement Heat Transfer with Perforated Baffles
`
`Combined Effect of Swirl and Impingement
`
`Concept of Heat Pipe for Turbine Cooling
`
`4.4.10
`
`New Cooling Concepts
`
`CHAPTER 5
`
`TURBINE INTERNAL COOLING WITH ROTATION
`
`5.1 ROTATIONAL EFFECTS ON COOLING
`5.2 SMOOTH-WALL COOLANT PASSAGE
`5.2.1 Effect of Rotation on Flow Field
`Effect of Rotation on Heat Transfer
`5.2.2.1 Effect of Rotation Number
`
`5.2.2
`
`5.2.2.2 Effect of Density Ratio
`5.2.2.3 Combined Effects of Rotation
`Number and Density Ratio
`
`355
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`366
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`371
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`371
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`373
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`377
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`385
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`390
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`391
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`395
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`398
`
`399
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`404
`
`406
`
`406
`
`406
`
`411
`
`414
`
`420
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`421
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`425
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`431
`
`432
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`436
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`439
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`439
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`439
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`439
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`448
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`CONTENTS
`
`5.2.2.4 Effect of Surface-Heating Condition
`5.2.2.5 Effect of Rotation Number and
`Wall-Heating Condition
`
`5.3 HEAT TRANSFER IN A RIB-TURBULATED ROTATING
`COOLANT PASSAGE
`5.3.1 Effect of Rotation on Rib-Turbulated Flow
`5.3.2 Effect of Rotation on Heat Transfer in
`Channels with 90° Ribs
`
`5.3.2.1 Effect of Rotation Number
`
`5.3.2.2 Effect of Wall Heating Condition
`
`5.3.3 Effect of Rotation on Heat Transfer for
`Channels with Angled (Skewed) Ribs
`5.3.3.1 Effect of Angled Ribs and Heating Condition
`5.3.3.2 Comparison of Orthogonal and
`Angled Ribs
`
`5.4 EFFECT OF CHANNEL ORIENTATION WITH RESPECT TO
`THE ROTATION DIRECTION ON BOTH SMOOTH AND
`RIBBED CHANNELS
`
`5.4.1 Effect of Rotation Number
`
`5.4.2 Effect of Model Orientation and Wall-Heating
`Condition
`
`5.5 EFFECT OF CHANNEL CROSS SECTION ON ROTATING
`HEAT TRANSFER
`5.5.1 Triangular Cross Section
`5.5.2 Rectangular Channel
`5.5.3 Circular Cross Section
`5.5.4 Two-Pass Triangular Duct
`
`xi
`
`452
`
`456
`
`458
`
`458
`
`461
`
`461
`
`465
`
`467
`
`469
`
`472
`
`474
`
`474
`
`474
`
`482
`
`482
`
`485
`
`488
`
`489
`
`5.6 DIFFERENT PROPOSED CORRELATION TO RELATE THE
`HEAT TRANSFER WITH ROTATIONAL EFFECTS
`5.7 HEAT.MASSoTRANSFER ANALOGY AND DETAIL
`MEASUREMENTS 503
`
`497
`
`5.8 ROTATION EFFECTS ON SMOOTH-WALL IMPINGEMENT
`COOLING
`5.8.1 Rotatiori Effects on Leading-Edge
`Impingement Cooling
`
`5.8.2 Rotation Effect on Midchord Impingement Cooling
`
`5.8.3 Effect of Film-Cooling Hole
`5.9 ROTATIONAL EFFECTS ON RIB-TURBULATED WALL
`IMPINGEMENT COOLING
`
`504
`
`504
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`512
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`518
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`CHAPTER 6
`EXPERIMENTAL METHODS
`INTRODUCTION
`6.1
`6.2 HEAT-TRANSFER MEASUREMENT TECHNIQUES
`
`6.2.1
`
`Introduction
`
`6.2.2 Heat Flux Gauges
`6.2.3 Thin-Foil Heaters with Thermocouples
`6.2.4 Copper Plate Heaters with Thermocouples
`6.2.5 Transient Technique
`6.3 MASS-TRANSFER ANALOGY TECHNIQUES
`Introduction
`6.3.1
`
`6.3.2 Naphthalene Sublimation Technique
`
`6.3.3 Foreign-Ga,s Concentration Sampling Technique
`6.3.4 Swollen-Polymer Technique
`6.3.5 Ammonia-Diazo Technique
`
`6.3.6 Pressure-Sensitive Paint (PSP) Techniques
`
`6.4 OPTICAL TECHNIQUES
`Introduction
`
`6.4.1
`
`6.4.2
`
`Infrared Thermography
`
`6.4.3 Thermographic Phosphors
`
`6.5 LIQUID CRYSTAL THERMOGRAPHY
`6.5.1 Steady-State Yellow-Band Tracking Technique
`
`6.5.2 Steady-State HSI Technique
`
`6.5.3 Transient HSI Technique
`
`6.5.4 Transient Single-Color Capturing Technique
`
`6.6 FLOW AND THERMAL FIELD MEASUREMENT
`TECHNIQUES
`Introduction
`6.6.1
`
`6.6.2 Five-Hole Probe/Thermocouples
`
`6.6.3 Hot-Wire/Cold-Wire Anemometry
`
`6.6.4 Laser Doppler Velocimetry (LDV)
`
`6.6.5 Particle Image Velocimetry
`6.6.6 Laser Holographic Interferome~ry
`
`6.6.7 Surface Visualization
`
`6.7 CLOSURE
`
`CONTENTS
`
`531
`531
`
`531
`
`531
`
`532
`
`535
`
`538
`
`539
`
`54O
`
`540
`
`540
`
`543
`
`545
`
`546
`
`547
`
`548
`
`548
`
`548
`
`550
`
`553
`
`554
`
`555
`
`557
`
`559
`
`565
`
`565
`
`565
`
`567
`
`568
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`570
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`572
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`572
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`CONTENTS xiii
`
`CHAPTER 7
`
`NUMERICAL MODELING
`7,1 GOVERNING EQUATIONS AND TURBULENCE
`MODELS
`Introduction
`
`7.1.1
`7.1.2 Governing Equations
`7.1.3 Turbulence Models
`Standard k-e Model
`7.1.3.1
`Low-Re k-g Model
`
`7.1.3.2
`
`7.1.3.3
`
`7.1.3.4
`
`7.1.3.5
`
`7.1.3.6
`
`Two-Layer k-g Model
`k-~o Model
`Baldwin-Lomax Model
`Second-Moment Closure Model
`Algebraic Closure Model
`
`7.1.3.7
`7.2 NUMERICAL PREDICTION OF TURBINE
`HEAT TRANSFER
`Introduction
`
`7.2.1
`7.2.2 Prediction of Turbine Blade/Vane Heat Transfer
`7.2.3 Prediction of the Endwall Heat Transfer
`7.2.4 Prediction of Blade Tip Heat Transfer
`7.3 NUMERICAL PREDICTION OF TURBINE
`FILM COOLING
`Introduction
`
`7.3.1
`7.3.2 Prediction of Flat-Surface Film Cooling
`7.3.3 Prediction of Leading-Edge Film Cooling
`7.3.4 Prediction of Turbine Blade Film Cooling
`7.4 NUMERICAL PREDICTION OF TURBINE
`INTERNAL COOLING
`
`7.4.1
`
`7.4.2
`
`7.4.3
`
`7.4.4
`
`7.4.5
`
`7.4.6
`
`7.4.7
`
`Introduction
`
`Effect of Rotation
`
`Effect of 180° Turn
`
`Effect of Transverse Ribs
`
`Effect of Angled Ribs
`
`Effect of Rotation on Channel Shapes
`
`Effect of Coolant Extraction
`
`585
`
`585
`
`585
`
`586
`
`586
`
`587
`
`588
`
`588
`
`589
`
`589
`
`590
`
`590
`
`591
`
`591
`
`592
`
`597
`
`599
`
`601
`
`601
`
`601
`
`607
`
`608
`
`610
`
`610
`
`6!0
`
`614
`
`619
`
`621
`
`625
`
`628
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`xiv
`
`CHAPTER 8
`
`FINAL REMARKS
`8.1 TURBINE HEAT TRANSFER AND FILM COOLING
`
`8.2 TURBINE INTERNAL COOLING WITH ROTATION
`
`8.3 TURBINE EDGE HEAT TRANSFER AND COOLING
`
`8.4 CLOSURE
`
`INDEX
`
`CONTENTS
`
`637
`
`637
`
`637
`
`638
`
`638
`
`639
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`

`Preface
`
`Gas turbines are used for aircraft propulsion and in land-based power genera-
`tion or industrial applications. Modern development in turbine-cooling tech-
`nology plays a critical role in increasing the thermal efficiency and power out-
`put of advanced gas turbines. Research activities in turbine heat transfer and
`cooling began in the early 1970s; since then, many research papers, state-
`of-the-art review articles, and book chapters have been published. However,
`the~’e is no book focusing entirely on the range of gas turbine heat-transfer
`issues and the associated cooling technologies.
`This book is intended as a reference book for researchers and engineers
`interested in working with gas turbine heat-transfer and cooling technology.
`Specifically, this book is for researchers and engineers who are new to the
`field of turbine heat-transfer analysis and cooling design; but it can also be a
`text or reference book for graduate-level heat-transfer and turbomachinery
`classes.
`In the beginning, we were thinking of covering all aspects of gas turbine-
`related heat-transfer and cooling problems. After careful survey, however, we
`decided to focus on the heat-transfer and cooling issues related to turbine
`airfoils only, because a vast breadth of information on this subject alone is
`available in the published literature. Assembling all the scattered informa-
`tion in a single compilation requires a great deal of effort. The book does not
`include combustor liner cooling and turbine disk-cooling problems although
`they are important to gas turbine hot gas path component designs. The book
`is divided into eight chapters:
`
`Chapter I
`
`Chapter 2
`
`Chapter 3
`
`Chapter 4
`
`Fundamentals. Discusses the need for turbine cooling, gas
`turbine heat-transfer problems, and cooling methodology
`Turbine Heat Transfer. Discusses turbine rotor and stator
`heat-transfer issues including endwall and blade tip region
`under engine conditions as well as under simulated engine
`conditions
`Turbine Film Cooling. Includes turbine rotor and stator blade
`film cooling and a discussion of the unsteady high free-
`stream turbulence effect on simulated cascade airfoils
`Turbine Internal Cooling. Includes impingement cooling, rib-
`turbulated cooling, pin-fin cooling, and compound and new
`cooling techniques
`
`xv
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`xvi
`
`PREFACE
`
`Chapter 5
`
`Chapter 6
`
`Chapter 7
`
`Chapter 8
`
`Turbine Internal Cooling with Rotation. Discusses the effect
`of rotation on rotor coolant passage heat transfer
`Experimental Methods. Includes heatJtransfer and mass-
`transfer techniques, liquid crystal thermography, optical
`techniques,
`flow and thermal field measurement
`techniques
`Numerical Modeling. Discusses governing equations and
`turbulence models, and their applications for predicting
`turbine blade heat transfer and film cooling and turbine
`blade internal cooling
`Final Remarks. Provides suggestions for future research in
`this area
`
`The open literature has many excellent articles available on this subject;
`however, we cannot use all of them in this book. We do not claim any new
`ideas in this book, but we do attempt to present the topic in a systematic and
`logical manner. We hope this book is a unique compilation and is useful for
`the gas turbine community. We are happy to receive constructive comments
`and suggestions on the material in the book.
`
`~[e-Chin Han
`Sandip Dutta
`Srinath Ekkad
`
`May I999
`
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`CHAPTER 1
`
`Fundamentals
`
`1.1 NEED FOR TURBINE BLADE COOLING
`1.1.1 RECENT DEVELOPMENT IN AIRCRAFT ENGINES
`
`Gas turbines are used for aircraft propulsion and in land-based power
`generation or industrial applications. Thermal efficiency and power output
`of ga~ turbines increase with increasing turbine rotor inlet temperatures.
`This is dramatically illustrated in Fig. 1.1, which plots specific core power
`production (which can be related to specific thrust) as a function of turbine
`rotor inlet temperature. The engines tend to track fairly close to the ideal
`performance line, which represents a cycle power output with 100 % effi-
`cient turbines with no leakage or cooling flows. Clearly, increasing rotor inlet
`temperatures is oneof the key technologies in raising gas turbine engine
`performance. Figure 1.2 shows that the rotor inlet temperatures (RIT) in
`advanced gas turbines are far higher than the melting point of the blade
`material: therefore, turbine blades need to be cooled. To double the engine
`power in aircraft gas turbines by the year 2003, the RIT should increase
`from today’s 2500°F to 3500°F using the same amount of cooling air (3-5 % "
`of compressor bleed air). Meanwhile, the compressor pressure ratio should
`increase from today’s 20 times the compression ratio to 40 times the com-
`pression ratio, or even higher, as shown in Fig. 1.3. This means that future
`aircraft gas turbines would have a higher turbine inlet temperature with the
`same amount of hotter cooling air from high-pressure compressor bleed.
`Therefore, high4emperature material development such as thermal barrier
`coating (TBC) or highly sophisticated cooling schemes are two important is-
`sues that need to be addressed to ensure high,performance, high-power gas
`turbines for the next century. To reach this goal, the U.S. Department of De-
`fense (DOD), NASA, and U.S. aircraft gas turbine manufacturers established
`the long-range R&D program known as integrated high-performance turbine
`engine technology (IHPTET). Begun in 1993, it was targeted with doubling
`the engine power by the year 2003 (Daly, 1993). Research and Development
`funds are provided by the U.S. Air Force, Navy, Army, and NASA, and by the
`U.S. gas turbine manufacturers such as GE Aircraft Engines, Pratt & Whitney,
`Allison, and Allied Signal. Research is performed at U.S. government lab-
`oratories, industrial laboratories, and university laboratories. Figure 1.4 is
`a diagram of the Fl17 turbofan engine developed by Pratt & Whitney.
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`2
`
`1000
`
`FUNDAMENTALS
`
`Hydrocarbon
`stoichiometric
`
`900 limit
`
`Specific
`core
`power
`
`800
`
`700
`
`600
`
`5OO
`
`400
`
`300
`
`200
`
`Ideal performance
`
`:1939)
`
`Year 2000+
`
`. Increased flowpath ~1
`¯ Reduced leakage
`. High temperature materials
`= Improved cooling effectiveness
`
`1 O0 -~
`
`w4ooo
`~o
`
`1200 ,1600 2000 2400 2800 3200 3600 4000 4400
`Turbine rotor inlet temperature - °F
`
`~’|~ure 1.1: Increased turbine inlet temperature dramatically improves cycle power
`output (courtesy of Pratt & Whitney; Saumer et al., 1992; AGARD CP 527).
`
`COOLING
`-- CONCEPT
`
`AND OTHERS
`
`ME NT
`
`INTRODU CT ION
`OF BLADE COOL1N(
`
`ATED
`COOLING SYSTEMS
`
`260O
`
`2400
`
`2200
`
`2000
`
`1800
`
`1600
`
`1400
`
`1200
`
`UNCOOL~D
`
`TURBINESI
`
`I O001o~O
`
`1960
`
`1970
`
`I
`1980
`
`I
`1990
`
`~
`
`2010
`
`YEAR
`
`l~lgute 1.2: Variation of turbine entry temperature over recent years
`(Clifford, 1985; AGARD CP 390; collected in Lakshminarayana, 1996).
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`NEED FOR TURBINE BLADE COOLING 3
`
`ADV. TURBOFAN-..,.
`
`60--
`
`-
`
`lO~
`
`0
`
`TURBOJET
`TURBOFAN
`TURBOPROP
`
`930 19~
`
`1950 1960 1970 1980 1990 2000 2010
`YFAR OF FIRST FLIGHT
`
`Figure 1.3: Progress in compressor pressure ratio (Rohlik,
`1983; NASA TM 83414; collected in Lakshminarayana, 1996),
`
`FI17 TURBOFAN ENGINE
`(Military Version of PW2040)
`
`-CONTROU.*ED
`DIFFUSION
`AIRFOILS
`
`COMBUSTOR COOLING
`
`CRYSTAL BLdDES
`
`PdETAL DISKS
`
`ELECTRONIC
`ENGINE
`CONTROL
`
`Figure 1.4:Fl17 turbofan engine developed by Pratt & Whitney (courtesy of Pratt
`& Whitney).
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`4
`
`FUN DAMENTALS
`
`All R&D activities are aimed at doubling the capability of turbine engines
`through (1) improved cooling effectiveness, (2) high,temperature materials
`with TBC, and (3) increased flow path efficiency with reducing leakage.
`
`1.1.2 RECENT DEVELOPMENT IN LAND~BASED
`GAS TURBINES
`
`For land-based gas turbines, including power generation (300 MW
`bined cycles), marine propulsion, and industrial applications such as pump-
`ing and cogeneration (less than 30 MW), the RIT should maintain today’s
`level of 2500°F-2600°F due to the constraint of the NO× pollution problem.
`Therefore, the main issue for land-based gas turbines is how to further im-
`prove the thermal efficiency at the current RIT level--for example, how to
`move the efficiency for the stand-alone gas turbine from today’s 35 % up to
`40 %, and for the large gas turbine combined cycle from today’s 55 % up
`to 60 % by the year 2003. To reach this goal, the U.S. Department of En-
`ergy (DOE) and the 0.S. land-based gas turbine manufacturers established a
`long-range R&D program known as advanced turbine systems (ATS). Begun
`in 1992, the ATS program was targeted to increase the combined cycle effi-
`ciency to 60 % by the year 2002 (Davis and Randolph, 1993). Research and
`Development funds are provided by the U.S. DOE and by the gas turbine
`manufacturers such as GE Power Systems, Westinghouse Electric, Allison,
`Solar Turbines, and Allied Signal. Research is performed at government lab-
`oratories, industrial laboratories, and university laboratories. For example,
`Fig. 1.5 is a schematic of the H-Engine Hot Gas Parts developed by GE
`Power Systems under the U.S. DOE ATS program. They used a closed-loop
`
`Gas Turbine. Hot Gas Path Parts
`
`[ ~.~ ~e~r~’~.
`¯ Higher Firing Temperature -~’~’~ ~_/~.~.-~
`Maximizes OMp~t
`~w No~e ~ T Minimizes NO
`Combustion T~mp~ture = Firin~
`
`First-Stage No~e
`
`Firing Te~nperature
`Produces Work
`
`Flgu~ 1.5: Relationship-combustion temperature to fire temperature for H-Engine
`Hot Gas Parts developed by GE Power Systems under U.S. DOE ATS Program (Corman
`and Paul, 1995).
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`1.2 TURBINE-COOLING TECHNOLOGY 5
`
`steam-cooled nozzle with TBC in order to reduce the hot-gas temperature
`drop through the first-stage nozzle. Therefore, the rotor inlet temperature
`(RIT or firing temperature) can be higher to produce more power at the
`same combustion temperature. The GE H-Engine uses the most advanced
`combustor design incorporating fuel/air premixing and lean combustion to
`reduce the environmental pollution problem (dry low NOx). All research
`activities are aimed at (1) combustion and combustion instability issues be-
`cause of the use of lean-premixed dry low NO× combustors, (2) material
`development such as a single crystal blade with TBC and ceramic blades,
`and (3) advanced turbine blade cooling such as a closed-loop steam-cooled
`blade with TBC.
`
`1.2 TURBINE-COOLING TECHNOLOGY
`1.2.1 CONCEPT OF TURBINE BLADE COOLING
`
`" Advanced gas turbine engines operate at high temperatures (2500-
`2600°F) to improve thermal efficiency and power output. As the turbine in-
`let temperature increases, the heat transferred to the turbine blades also
`increases. The level and variation in the temperature within the blade mate-
`rial (which causes thermal stresses) must be limited to achieve reasonable
`durability goals.
`The operating temperatures are far above the permissible metal temper-
`atures. Therefore, there is a need to cool the blades for safe operation. The
`blades are cooled by extracted air from the compressor of the engine. Since
`this extraction incurs a penalty to the thermal efficiency, it is necessary to
`understand and optimize the cooling technique, operating conditions, and
`turbine blade geometry. Gas turbine blades are cooled internally and ex-
`ternally. Internal cooling is achieved by passing the coolant through several
`¯ enhanced serpentine passages inside the blades and extracting the heat
`from the outside of the blades. Both jet impingement and pin-fin cooling are
`also used as a method of internal cooling. External cooling is also called film
`cooling. Internal coolant air is ejected out through discrete holes or slots to
`provide a coolant film to protect the outside surface of the blade from hot
`combustion gases.
`The engine cooling system must be designed to ensure that the max-
`imum blade surface temperatures and temperature gradients during oper-
`ation are compatible with the maximum blade thermal stress for the life
`of the design. Too little coolant flow results in hotter blade temperatures
`and reduced component life. Similarly, too much coolant flow results in re-
`duced engine performance. The engine cooling system must be designed to
`minimize the use of compressor bleed air for cooling purposes to achieve
`maximum benefits of the high inlet gas temperature.
`Highly sophisticated cooling techniques in advanced gas turbine engines
`include film cooling, impingement cooling, and augmented convective cool-
`ing. Figures 1.6 and 1.7 show the cutaway view of the General Electric CF6
`turbofan engine and the stage-1 high-pressure nozzle guide vane (NGV) for
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`6
`
`FUN DAMENTALS
`
`Figure 1.6: Cutaway view of GE CF6 turbofan engine (Treager, 1979).
`
`the GE-CF6 turbofan engine, respectively. The cooling air comes from the
`14th-stage compr~essor bleed and impinges on the inner walls of the NGV.
`After impingement cooling, the spent air provides film cooling through the
`leading edge holes, gill holes, midchord holes, and trailing edge slots. Fig-
`ure 1.8 shows the stage-1 internally cooled high-Pressure turbine rotor blade
`for a GE CF6 turbofan engine. The cooling system is based on the use of
`convective cooling in the leading edge region and film cooling through the
`gill holes, augmented convective cooling with rib turbulators in the midchord
`region, and squealer tip-cap cooling and augmented convective cooling with
`pin fins in combination with film cooling at the trailing edge. The optimum
`
`l~igure 1.7: The stage-1 high-pressure turbine nozzle vane for the GE CF6 engine
`(Treager, 1979).
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`

`1.2 TURBINE-COOLING TECHNOLOGY 7
`
`TIP-CAP HOLES
`
`SQUEALER TIP J m’~,P
`
`, -- SQUEALER TiP HOLE
`
`HOLES
`
`HOLES
`
`BLADE ~
`PLATFORM ~
`
`Figure 1.8: The stage-1 high-pressure turbine rotor blade for the GE CF6
`(Treager, 1979).
`
`engine
`
`AIRFOIL AIR-INLET HOLES
`
`combination of these cooling techniques to meet the highly complex design
`requirements is the key to designing air-cooled turbine blades and vanes.
`
`1.2.2 "13q~IC81, TURBINE COOLING SYSTEM
`
`Gas turbine-cooling technology is complex and varies from engine man-
`ufacturer to engine manufacturer. Even the same engine manufacturer uses
`different cooling systems for various engines. Most turbine cooling systems
`are proprietary in nature and also are not available in open literature. How-
`ever, most cooling-system designs are quite similar regardless of engine
`manufacturer and models. The following paragraph will discuss a typical
`turbine cooling system using NASA’s energy efficient engine (E3), developed
`by GE Aircraft Engines (Halila et al., 1982). This is used as an example since
`it is available in the public domain. Note that the cooling systems for today’s
`advanced gas turbine engines have improved beyond the E~ engine.
`Figure 1.9 is an overall view of the rotor, stator, and casing cooling supply
`system. The stage- 1 nozzle is cooled by air extracted from the inner and outer
`combustion liner cavities, and the stage-1 rotor is cooled by air extracted at
`the compressor diffuser midspan. The stage 2 nozzle coolant comes from
`the stage 7 compressor bleed, and the stage 2 rotor coolant comes from the
`stage-1 rotor inducer system. Figures 1.10 through I. 12 show the cooling air
`supply for the nozzle pitch line and the inner-band and outer-band design.
`This design includes two separate impingement inserts and trailing edge
`pressure side bleed slots. The design also uses both impingement cooling
`and film cooling at the nozzle leading edge and midchord region with two
`rows of compound angle holes on the pressure side and two rows of diffusion-
`shaped holes on the suction side. The vane i