Applications of Circulation Control Technologies, Volume 214 (Progress in Astronautics and Aeronautics)

Applications of Circulation Control Technologies

Edited by Ronald D. Joslin Office of Naval Research Arlington, Virginia Gregory S. Jones NASA Langley Research Center Hampton, Virginia

Volume 214 PROGRESS IN ASTRONAUTICS AND AERONAUTICS Frank K. Lu, Editor-in-Chief University of Texas at Arlington Arlington, Texas

Published by the American Institute of Aeronautics and Astronautics, Inc. 1801 Alexander Bell Drive, Reston, Virginia 20191-4344

American Institute of Aeronautics and Astronautics, Inc., Reston, Virginia 1 2 3 4 5 Copyright 0 2006 by the American Institute of Aeronautics and Astronautics, Inc. Printed in the United States of America. All rights reserved. Reproduction or translation of any part of this work beyond that permitted by Sections 107 and 108 of the U.S. Copyright Law without the permission of the copyright owner is unlawful. The code following this statement indicates the copyright owner’s consent that copies of articles in this volume may be made for personal or internal use, on condition that the copier pay the per-copy fee ($2.50) plus the per-page fee ($0.50) through the Copyright Clearance Center. Inc., 222 Rosewood Drive, Danvers, Massachusetts 01923. This consent does not extend to other kinds of copying, for which permission requests should be addressed to the publisher. Users should employ the following code when reporting copying from the volume to the Copyright Clearance Center: 1-56347-789-0/06 $2.50

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Data and information appearing in this book are for informational purposes only. AIAA is not responsible for any injury or damage resulting from use or reliance, nor does AIAA warrant that use or reliance will be free from privately owned rights.

ISBN 1-56347-789-0

Progress in Astronautics and Aeronautics Editor-in-Chief Frank K. Lu University of Texas at Arlington

Editorial Board David A. Bearden The Aerospace Corporation

Abdollah Khodadoust The Boeing Company

John D. Binder viaSolutions

Richard C. Lind University of Florida

Steven A. Brandt U.S. Air Force Academy

Richard M. Lloyd Raytheon Electronics Company

Fred R. DeJamette North Carolina State University

Frank Pai University of Missouri-Columbia

Gail Klein Jet Propulsion Laboratory

Ning Qin University of Shefield

George Eitalbery German-Dutch Wind Tunnels

US.Naval Postgraduate School

Sanjay Garg NASA Glenn Research Center

Ben T. Zinn Georgia Institute of Technology

Eswar Josyula

Peter H. Zipfel U.S. Air Force Research Laboratory

US.Air Force Research Laboratory

Oleg Yakimenko

Foreword

T

HIS collection of papers represents a compilation of the state-of-the-art in circulation-control technologies by two of the foremost experts in the field. The volume is conveniently organized to enable experts and beginners alike to quickly obtain a thorough historical overview and then be brought up to speed on the latest research. The final chapter delves into new areas and draws attention to exciting new ideas in circulation control. A wide range of advanced experimental and numerical methods are discussed by a panel of international experts. The text will prove to be of great value to workers in this field.

Frank K. Lu Editor-in-Chief Progress in Astronautics and Aeronautics

Preface

T

HE GENESIS of this volume originated during the planning of the NASA/ ONR Circulation Control Workshop, which was held March 2004 in Hampton, Virginia. Over two full days, 30 papers and 4 posters were presented, with 110 scientists, engineers, and program managers in attendance. This book was conceived to distribute this rich body of technical information on circulation control to a broader audience and to provide historical documentation to support future circulation control applications. Since that workshop, the papers have been updated and peer-reviewed to arrive at a compilation of the state of the art in circulation-control technologies. The goals of this book are 1) to summarize the history and the state of the art in circulation control technology, 2) to provide a single up-to-date knowledgebase for circulation control design, analysis, and experimental testing, and 3) to highlight prediction tools for circulation control. Goals 1 and 2 are clearly achieved in the chapters by the diverse applications and significant breadth of insights offered by the experts in this field. Goal 3 is most notably achieved by the use and discussion of the diverse range of computational fluid dynamics (CFD) tools for circulation control. Results showing the successful prediction of performance and inadequacies of some predictions are presented for completeness. The book is divided into four sections. The first major section presents a historical overview of circulation control. Because the overview papers are very thorough, many of the remaining chapters present brief introductions. The second major section covers experiments and applications. Section I1 is divided into A. fundamental flow physics, B. aerospace applications, and C. nonaerospace applications. The third major section covers CFD-based prediction tools and some validation with experiments (most of which are detailed in Section 11). Section I11 is subdivided by the different predictive applications. Finally, the last section consists of a single chapter, which introduces a vision for the use of circulation control in a broad spectrum of nonvehicle applications. Although less rigorous than most chapters, this final chapter exposes the reader to some new insights into applications of circulation-control technologies.

Ronald D. Joslin Gregory S. Jones December 2005

xix

Table of Contents Preface

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xix

Overview

Chapter 1 Advantages of Combining BLC Suction with Circulation Control High-Lift Generation

............................

John L . Loth. West Virginia University. Morgantown. West Virginia

3

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Designing a CC Technology Demonstrator STOL Aircraft . . . . . . . . . . . . . . . . . . 5 1974 Flight Testing of the WVU CC Technology Demonstrator . . . . . . . . . . . . . . 12 1979 CC Flight Tests with a Grumman Aerospace A-6A . . . . . . . . . . . . . . . . . . 16 Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18 References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

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Chapter 2 Overview of Circulation Control Pneumatic Aerodynamics: Blown Force and Moment Augmentation and Modification as Applied Primarily to Fixed-Wing Aircraft Robert J. Englar. Georgia Institute of Technology. Atlanta. Georgia

.......

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Coanda Effect . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Applications of Circulation Control. Past and Present . . . . . . . . . . . . . . . . . . . . Powered Lift and Engine Thrust Deflection . . . . . . . . . . . . . . . . . . . . . . . . . . . Other Aircraft Applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nonflying Applications of Circulation Control . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 3 Exploratory Investigations of Circulation Control Technology: Overview for Period 1987-2003 at NSWCCD

.......

23 23 24 25 28 48 53 57 63 64

69

Robin Imber. Naval Air Systems Command. Patuxent River. Maryland; Ernest Rogers and Jane Abramson. Naval Surface Warfare Center-Carderock Division. West Bethesda. Maryland Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ix

69 70

X

Dual-Slotted Cambered Airfoil (LSB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Self-Driven Rotary Thruster (TIPJET) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annular Wing (CC-Duct) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Circular Wing (CC-Disc) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Miniature Oscillatory Valve (CC-Valve) for Unsteady Wing Load Reduction . . . . . Dual-Slotted Low Aspect Ratio Wing (CC Hydrofoil) . . . . . . . . . . . . . . . . . . . . Status of Design Capability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

70 73 79 85 91 93 99 100 101

1I.A. Experiments and Applications: Fundamental Flow Physics

.

Chapter 4 Measurement and Analysis of Circulation Control Airfoils

.....................................

105

F. Kevin Owen. Complere Inc., Paczjic Grove. California; Andrew K. Owen. University of Oxford. Oxford. England. United Kingdom

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Experimental Details . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sample Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

Chapter 5

Some Circulation and Separation Control Experiments

105 106 107 107 112 112

..

113

Dino Cerchie. Eran Halfon. Andreas Hammerich. Gengxin Han. Lutz Taubert. Lucie.Trouve. Priyank Varghese. and Israel Wygnanski. University of Arizona. Tucson.Arizona Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discussion of Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

Chapter 6

Noise Reduction Through Circulation Control

Scott E. Munro. Krishan K . Ahuja. and Robert J. Englar. Georgia Institute of Technology. Atlanta. Georgia

113 114 118 162 164 164

........ 167

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Facilities and Instrumentation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Technical Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

167 168 169 171 173 174

xi Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

184 186 186

1I.B. Experiments and Applications: Aerospace

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Chapter 7 Pneumatic Flap Performance for a Two-Dimensional Circulation Control Airfoil

..............................

191

Gregory S. Jones. NASA Langley Research Center. Hampton. Virginia Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NASA CC Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Theoretical Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GACC Airfoil Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Experimental Setup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airfoil Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Appendix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 8 Trailing Edge Circulation Control of an Airfoil at Transonic Mach Numbers

..............................

191 192 193 195 202 207 216 236 237 241

245

Michael G. Alexander. Scott G . Anders. and Stuart K . Johnson. NASA Lungley Research Center. Hampton. Virginia Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Model Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Instrumentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Facli ty . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test Procedures and Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discussion of Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

245 246 247 251 252 253 254 254 263 275 275

Chapter 9 Experimental and Computational Investigation into the Use of the Coanda Effect on the Bell A821201 Airfoil

........... 277

Gerald Angle 11. Brian O’Hara. Wade Huebsch. and James Smith. West Virginia University. Morgantown. West Virginia Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Experimental Apparatus and Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . .

277 278 279

xii Computational Model and Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Experimental Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computational Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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282 285 286 290 291

Chapter 10 Novel Flow Control Method for Airfoil Performance Enhancement Using Co-Flow Jet

................ 293

Ge-Cheng Zha and Craig D. Paxton. University of Miami. Coral Gables. Florida Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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293 294 296 311 312 312

Chapter 11 Experimental Development and Evaluation of Pneumatic Powered-Lift Super-STOL Aircraft

................ 315

Robert J. Englar. Georgia Institute of Technology. Atlanta. Georgia; Bryan A . Campbell. NASA Langley Research Center. Hampton. Virginia

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Experimental Apparatus and Test Techniques. . . . . . . . . . . . . . . . . . . . . . . . . Wind-Tunnel Evaluations and Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Comparison of Measurements and Predictions . . . . . . . . . . . . . . . . . . . . . . . . Potential Applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 12

Use of Circulation Control for Flight Control

315 316 320 321 331 333 333 335 335

........ 337

Steven l? Frith and Norman J. Wood, University of Manchester. Manchester. England. United Kingdom

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Half-Span Cropped-Delta Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Full-Span UAV Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

337 338 339 345 352 353 35 3

xiii

1I.C. Experiments and Applications: Nonaerospace

.

Chapter 13 Pneumatic Aerodynamic Technology to Improve Performance and Control of Automotive Vehicles

.............. 357

Robert J. Englar. Georgia Institute of Technology. Atlanta. Georgia

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Basics of Pneumatic Circulation Control Aerodynamics . . . . . . . . . . . . . . . . . . DOE Pneumatic Heavy Vehicle Model Test Results . . . . . . . . . . . . . . . . . . . . Pneumatic HV Fuel Economy Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Updated Wind Tunnel Evaluations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Sport Utility Vehicles (PSUVs) . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Recommendations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

357 357 358 360 367 371 374 379 380 381 381

Chapter 14 Aerodynamic Heat Exchanger: A Novel Approach to Radiator Design Using Circulation Control Richard J. Gaeta. Robert J. Englar. and Graham Blaylock. Georgia Institute of Technology. Atlanta. Georgia

................ 383

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Technical Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

383 383 386 389 395 397 397

1II.A. Tools for Predicting Circulation Control Performance: NCCR 1510 Airfoil Test Case

.

Chapter 15 Investigation of Turbulent Coanda Wall Jets Using DNSandRANS

.....................................

401

Hermann F. Fasel. Andreas Gross. and Stefan W e n . University of Arizona. Tucson. Arizona

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Investigated Configurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Numerical Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbulent Wall Jet on a Circular Cylinder . . . . . . . . . . . . . . . . . . . . . . . . . . . Circulation Control Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

401 402 403 404 405 415

xiv Conclusions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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418 419 419

Chapter 16 RANS and Detached-Eddy Simulation of the NCCR Airfoil

421

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Geometry. Conditions. and Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computational Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Grid Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial and Boundary Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

421 422 424 425 427 429 430 441 442 442

.......................................

Eric G . Paterson and Warren J. Baker. Pennsylvania State University. University Park. Pennsylvania

.

Chapter 17 Full Reynolds-Stress Modeling of Circulation Control Airfoils

445

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mathematical Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

445 446 448 453 465 465 465

.....................................

Peter A . Chang 111. Joseph Slomski. Thomas Marino. Michael P. Ebert. and Jane Abramson. Naval Surface Warfare Center-Carderock Division. West Bethesda. Maryland

1II.B. Tools for Predicting Circulation Control Performance: NCCR 103RE Airfoil Test Case

.

Chapter 18 Aspects of Numerical Simulation of Circulation Control Airfoils

469

Nomenclature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GeometryandGrid . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

469 470 472

.....................................

R. Charles Swanson. Christopher L . Rumsey. and Scott G. Anders. NASA Langley Research Center. Hampton. Virginia

xv Numerical Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Boundary and Initial Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbulence Modeling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jet Momentum Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Numerical Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Appendix: Coordinates of 103RE Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

Chapter 19 Role of Turbulence Modeling in Flow Prediction of Circulation Control Airfoils

.............................

475 476 476 478 478 495 497 497 497

499

Gregory McGowan. Ashok Gopalarathnam. Xudong Xiao. and Hassan Hassan. North Carolina State University. Raleigh. North Carolina Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Formulation of the Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

499 500 501 502 510 510 510

1II.C. Tools for Predicting Circulation Control Performance: GACC Airfoil Test Case

.

Chapter 20 Simulation of Steady Circulation Control for the General Aviation Circulation Control (GACC) Wing

........... 513

Warren J. Baker and Eric G. Paterson. Pennsylvania State University. University Park. Pennsylvania

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Geometry. Conditions. and Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computational Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Grid Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial and Boundary Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computational Resources . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

513 514 515 516 518 521 523 523 536 537 537

xvi

.

Chapter 21 Computational Study of a Circulation Control Airfoil Using FLUENT

................................

539

Gregory McGowan and Ashok Gopalarathnam. North Carolina State University. Raleigh. North Carolina Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Configurations and Experiments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Numerical Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

539 540 541 542 545 552 553 553

1II.D. Tools for Predicting Circulation Control Performance: Additional CFD Applications

.

Chapter 22 Computational Evaluation of Steady and Pulsed Jet Effects on a Circulation Control Airfoil

.................. 557

Yi Liu. Lakshmi N . Sankar. Robert J. Englar. Krishan K . Ahuja. and Richard Gaeta. Georgia Institute of Technology. Atlanta. Georgia

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mathematical and Numerical Formulation . . . . . . . . . . . . . . . . . . . . . . . . . . . Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.

557 558 559 561 575 575 575

Chapter 23 Time-Accurate Simulations of Synthetic Jet-Based Flow Control for a Spinning Projectile

.............. 579

Jubaraj Sahu. U S. Army Research Laboratory. Aberdeen Proving Ground. Maryland

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computational Methodology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Projectile Geometry and Computational Grid . . . . . . . . . . . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

579 580 581 584 586 594 595

xvii

. Exploring a Visionary Use of Circulation Control Chapter 24. Coanda Effect and Circulation Control for Nonaeronautical Applications ............................

599

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

599 600 612 612 612

...............................................

615

Author Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

623

Supporting Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

625

IV

Terence R . Day. Vortex Dynamics Pty Ltd. Mount Tamborine. Queensland. Australia

Index

I. Overview

Chapter 1

Advantages of Combining BLC Suction with Circulation Control High-Lift Generation John L. Loth* West Virginia University, Morgantown, West Virginia

Nomenclature CB = circulation control blowing efficiency factor CL = lift coefficient CLopt= optimum lift coefficient where aircraft L / D is maximum C, = blowing coefficient Di = induced drag D,, = parasite drag mcc = circulation control blowing mass flow rate pt = total pressure in the compressor bleed air supply duct q, = dynamic pressure r = circulation control rounded trailing edge radius S, = wing area t,, = non-dimensional circulation control blowing slot height tn = non-dimensional ejector nozzle slot height ts = non-dimensional suction slot height V,, = circulation control blowing velocity V, = equivalent airspeed, corrected for position error Vi = indicated airspeed V , = free stream velocity p = air density Subscripts a = angle of attack c = chord co = free stream conditions

*F’rofessor, Mechanical and Aerospace Engineering. Associate Fellow AIAA. Copyright 0 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

3

4

J. L. LOTH

I. Introduction HE PURPOSE of this paper is to present the advantages of combining boundary layer control (BLC) by suction with circulation control (CC) by blowing for aircraft high lift generation. In the short take-off and landing (STOL) mode, the sharp trailing edge of the wing must be converted into a rounded Coanda surface for CC blowing. Jet engine hot, high-pressure compressor bleed air is the most commonly used source for the blowing air. Ducting this hot, highpressure air to the CC blowing slot involves problems arising from factors such as duct size, weight, pressure loss, required insulation and thermal expansion joints, and jet engine take-off thrust loss. It is shown here how adding an ejector for BLC suction just upstream of the CC blowing slot can diminish the impact of the aforementioned problems. It can reduce the amount of compressor bleed air required, and thus duct size, by more than 50%, provide structural cooling, and improve the CC blowing to free stream velocity ratio, bringing it closer to four, where the theoretical lift augmentation ratio reaches a maximum. Flight test results using such a configuration are provided, together with solutions for in-flight transition from the CC rounded trailing edge to a sharp trailing edge for low-drag cruise. Data were collected in 1974 during flight testing of the first CC Technology Demonstrator Aircraft, at West Virginia University. The use of blowing air to augment airfoil lift had already been proposed',2 in the 1920s. D a ~ i d s o n ,in~ his 1960 British patent application, referred to the concept of blowing over a circular cylinder as circulation control (CC). To improve the lift-to-drag ratio, Kind and Maull? at Cambridge University, experimented with CC on elliptical airfoils. Kind is also credited with developing the first boundary layer theory for CC blowing to correlate his experimental results. At zero angle of attack, elliptic airfoils produce two nearly identical suction peaks at their leading and trailing edges; this results in an aft shift of the center of pressure and thus nose-down pitching moment. Typical streamlines for such an airfoil, computed by S h r e ~ s b u r yare , ~ shown in Fig. 1. A schematic of a CC blowing slot is shown in Fig. 2. Blowing air must be supplied uniformly to the blowing slot. By Coanda turning, the jet generates a high suction force on the rounded surface. The angular position of the lower surface stagnation point, where the Coanda jet separates when meeting the flow from below the airfoil, determines the circulation and lift produced. Even today, most disagreements between computational and experimental results are

T

Fig. 1 Computed streamlines for an elliptic airfoil?

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

5

High -pressure fluid flowing from

Fig. 2 Schematic of a CC trailing edge.

a result of the sensitive relationship between circulation and location of this lower surface stagnation point. In the late 1960s, Robert Williams,6 then at NSRDC (Naval Ship Research and Development Center), started to experiment with CC airfoils developed by Kind. Williams697investigated the feasibility of a heavy-lift helicopter with dual plenum elliptic rotor blades and valves to control CC blowing rate to allow high forward speed. In 1968, the Office of Naval Research (ONR) contracted, with West Virginia University (WVU), research on CC airfoils, including testing at high Reynolds number and away from wind tunnel wall interference. Loth and Fanucci considered the possibility of protruding a CC blown airfoil from the roof of one of the WVU flight test aircraft. However, this would not be safe, because the roll moment produced by such a CC airfoil would exceed the available aircraft aileron control. To satisfy the contract requirement of flight testing CC technology, they decided it would be safer to fly a fixed-wing aircraft with CC blown wings. In the case of blower failure, an elliptic airfoil would not be flyable; therefore, new CC wings were designed at WVU, which were in-flight convertible from a high-speed, low-drag, conventional sharp trailing edge to a rounded trailing edge with CC blowing during slow-speed flight testing. In the following five years, several such CC airfoils were tested at WVU in the wind tunnel there. A comparison of blowing air requirements and lift capability for various high-lift systems was completed in 1973, as shown in Fig. 3.' This indicates that CC high-lift generation is more conservative in blowing air requirement than other methods.

11. Designing a CC Technology Demonstrator STOL Aircraft A Bede-4 homebuilt kit was found to provide an economical and suitable frame for test flying a CC wing. The simplest arrangement for in-flight conversion from a sharp trailing edge to a rounded CC blown trailing edge was first investigated. This is a forward folding flap, which exposes its semicircular hinge to provide a rounded trailing edge for CC blowing, as shown in Fig. 4. Dr. Norio Inumaru, visiting WVU from NRL in Tokyo, Japan, designed its drooped leading edge to prevent leading edge stall at high lift. The test model

J. L. LOTH

6

Fig. 3 Performance comparison between powered high-lift systems.'

was fabricated by riveting a sheet metal cuff to the wing leading edge and filling its cavity with foam. In 1970, Model A wing was tested in CC mode in the twodimensional 8 x 10 ft NSRDC wind tunnel, both in the sharp and round trailing edge configurations (Fig. 5 ) . The test data for CL,a, and C , are shown in Fig. 6 . Below stall, in the angle of attack range - 2 < a < 8 deg, they could be

L.E. DROOP DESIGN

Fig. 4 WVU Model A CC wing, wind tunnel tested at NSRDC in 1970.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

7

Fig. 5 WVU Model A wing: left, in cruise; right, in CC mode.

curve-fitted by a linear equation:

In CC mode, the Model A chord length was reduced to 88% relative to cruise mode. All test data shown are referenced to cruise chord length, which effectively lowers CLm,, for the Model A in CC mode. The sharp trailing edge wing had a , = 0.09. Thus, for curve fitting test data in CC two-dimensional value of C mode, ,C , was reduced to 0.09 x 88% = 0.08. Excellent curve fitting was obtained by replacing SCL/SC, with cB/&, where CB is constant and named the blowing efficiency factor. Model A wing test data with CB = 4.3 provided the best curve fit, as shown in Fig. 7 using:

The disappointing performance of the Model A wing prompted a new design called the Model B wing. Instead of folding the flap inward for CC high lift generation, its flap was folded out. The 20% longer chord length, as shown in Figs. 8 and 9, was expected to increase the CC blowing efficiency factor CB from 4.3 to 4.3 x (120%/ 88%) = 5.9. In the Model B wing, great care was taken to achieve blowing

0 -I

0.2

0.4

0.6

0.8

1.0

CP,,

Fig. 6 WVU Model A CC wing, 1970 wind tunnel test results.

a

J. L. LOTH

5 4

1

0

0

0.4

0.2

0.6

0.8

CU

Fig. 7 WVU Model A CC wing empirical curve fit with CB = 4.3.

slot uniformity. This was accomplished by machining and bolting segmented aluminum nozzle blocks to an aluminum 3-in.-diam air supply duct, which also served as the rounded CC trailing edge. This provided a uniform 0.012in.-wide primary blowing slot (Fig. 9). The WVU wind tunnel model tests on a two-dimensional version of the Model B wing are described in Refs. 9 and 10. When applied to the CC Technology Demonstrator aircraft, the source of blowing air had to be selected. Boasson, in his dissertation, proved theoretically that the lift augmentation ratio CBreaches a maximum when Vcc/Vm = 4.'' Such low CC blowing velocity requires a high mass flow rate. Then Ap, the duct friction loss inside the air supplying 3-in.-diameter CC rounded trailing edge,

L.E. DROOP DESIGN

Fig. 8 WVU Model B CC wing, wind tunnel tested at WVU and flown on the CC Technology Demonstrator.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

9

SUCTloN BLC

Fig. 9 WVU Model B CC wing, wind tunnel tested at WVU and flown on the CC Technology Demonstrator.

would far exceed the total pressure required at the CC blowing slot! This prompted the selection of a high-pressure air source, an auxiliary power GTC 85-72 gas turbine. However, its bleed air temperature of about 300"F, created new problems to be solved. The 100-in.-long aluminum CC rounded trailing edges of the wings had to be mounted on sliding bearings to allow for up to 0.5 in. of thermal expansion. Cooling had to be provided to shield the fiberglass wing filled with fuel from the hot CC rounded trailing edge. Incorporating an ejector with suction slot just upstream of the CC blowing slot as shown in Fig. 9 solved this problem. This also reduced the required compressor bleed air mass flow by more than 50% and reduced the blowing velocity ratio V,,/Vm closer to its optimum value. The 3 in. air supply duct contained a bell crank to transfer the torque needed to stow the rounded CC trailing edge within the wing for low-drag, high-speed cruise. The following derivation is included to illustrate the reduction in required compressor bleed air required and increase in blowing momentum obtainable by combining an internal ejector with CC blowing. For simplicity, assume one-dimensional, incompressible flow and constant area ejector with negligible wall shear loss. The dimensionless area ratios used are equal to those in the model B wing. Defining ejector exit CC slot height by tcc= 1, and the dimensionless ejector nozzle slot height by dividing by the CC exit slot height as t, = The remaining area for the suction slot height is divided by exit slot height to give ts = The suction velocity V, is made dimensionless by dividing by the CC blowing velocity V,, = 1. Likewise, the nozzle velocity becomes V,. The subsonic incompressible flow exit boundary condition applies: p n = p s . The Bernoulli equation gives suction gage pressure as ps = -qs = -0.5pV;V:,. We next apply the following equations. Continuity equation:

a.

i.

tsvs

+ t,vn = 1

(3)

J. L. LOTH

10

Inserting values for slot height gives 3 4

-vs

1

+-v, 4

or

=1

v,

=4-3vs

Momentum equation: tsv,2+t,v,2+T= Ps

1

PVCC

(4)

Inserting values for slot height gives 3 1 -V,2+-V,2-0.5V,2= 4 4

1

Substituting V, from above gives the quadratic equation: 5V; - 12Vs

+6 = 0

Solving for Vs < V, gives Vs = 0.7 when inserted in the preceding relation, which gives V, = 1.9. This means that t,V, = x 1.9 = 47.5%; in other words, the nozzle needs to supply only 47.5% of the CC blowing air. The balance of the blowing air tsVs = x 1.9 = 52.5% is supplied by the BLC suction slot and need not be supplied through the CC rounded trailing edge duct. The CC jet exits at near ambient pressure with thrust Tcc = hccVCc.The required ejector nozzle thrust T, is only 0.83Tcc, This demonstrates that incorporating an ejector can 1) provide cooling by boundary layer suction, 2) increase CC blowing momentum by (1-0.83) or 17%, 3) lower the velocity ratio Vcc/VW to increase blowing efficiency factor CB.Furthermore, it reduces compressor bleed air mass flow rate required by 52.5% which lowered duct pressure loss with associated duct size and weight savings. In the WVU wind tunnel model tests, the availability of flap hinge suction also allowed the CC flap to be deflected up to 15 deg without stall for additional lift augmentation. Arrows in Fig. 11 highlight the special features of this aircraft. Arrows have been used to show the CC blowing slot on the top of the 3-in.diameter rounded trailing edge. Suction boundary layer control (BLC) is shown at the flap hinge. The pilot can dump the blowing air overboard by actuating an air dump valve to achieve Direct Lift Control, called (DLC) as indicated. A layout of the WVU CC Technology Demonstrator aircraft is shown in Fig. 10 with a GTC 85-72 gas turbine mounted in the rear passenger seat area. Note that the jet engine exhaust discharges upwards, to prevent igniting the blacktop on the parking area. In Fig. 12 are shown details of the flap retraction and extension mechanism by a two horsepower electric motor. It turns the CC air supply duct inside the

a

9

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

EMPTY WT-

11

1720 I@

W. BROSS 2400 Ib

Fig. 10 WVU CC Technology Demonstrator dimensions and layout.

cabin, which is welded by bell cranks to the two 3-in.-diameter CC rounded trailing edges. For increased roll control at low speed, the ailerons are drooped and blown with compressor bleed air supplied via small ejectors inside the cabin for cooling purposes. To increase aileron effectiveness, they are connected to a flow diverter valve, which alters the left and right wing blowing rate. The bolt shown in the air splitter tee serves as a hinge for the splitter valve inside this tee. Fences and structure used to strengthen the cavity at the bottom of the wing, into which the CC rounded trailing edge retracts.

BLC at flap hinge line

Fig. 11 WVU CC Technology Demonstrator location of CC, BLC, DLC, and space for flap folding.

12

J. L. LOTH

Fig. 12 Compressor bleed air enters into a worm-gear driven pipe, connected by bell cranks to the left and right CC rounded trailing edges.

111. 1974 Flight Testing of the WVU CC Technology Demonstrator Prior to 25 h of flight testing, which started on 10 April 1974, the CC slot was tested for blowing uniformity and its ejector for providing adequate cooling to the fiberglass wings. The flight tests, performed by test pilot Shawn Roberts, started with calibrating airspeed and position error, with the use of a Pitot tube mounted with a boom to the left wing tip (Fig. 13). This boom also contained angle of

Fig. 13 Large position error in cockpit speed indicator against equivalent airspeed based on boom-mounted pitot tube readings.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

13

attack and yaw measuring instrumentation. The aircraft in flight is shown in Fig. 14, with the CC blowing flap deployed. A summary of the flight test data is shown in Fig. 15. Shown are three scales for the lift coefficient, all based on dynamic pressure q, calculated using equivalent airspeed and reduced to sealevel density. The left column indicates the trimmed-out aircraft CL. The middle column has the tail download added to the lift and is termed CL wing On the far right column is the maximum CL value, which occurred at the flap centerline. For example, the average wing lift coefficient increased from 2.0 without blowing, at C , = 0, to 4.3 with blowing at C, = 0.12. Near stall the difference in angle of attack was negligible, thus the blowing efficiency factor C,, from Eq. (2) can be solved from:

G

C, = ACLI C = (4.3 - 2 ) / m = 6.6

(5)

This is more than 10% better than could be expected by extrapolating the Model A test results for the increased chord length, or C, = 4.3 x (1.2/0.88) = 5.86. This improvement can be credited to the utilization of an ejector in the Model B wing. It is interesting to calculate the CC blowing air horsepower required if the blowing air were supplied at standard sea-level conditions. The blowing slot height of 0.050 in. along the two 100-in.-long CC flaps resulted in blowing slot area A,, = 10 in.’. Consider flight with the propeller at idle, with C, = 0.12 and CL = 3.8. From the definition of C, = T,,/(q,S,) and CL= L/(q,S,), calculate blowing momentum

Tcc = (0.1213.8) x (W = 2400 lb) = 76 lbf

(6)

Fig. 14 WVU CC Technology Demonstrator during 25 h of flight testing, starting 10 April 1974.

14

J. L. LOTH

r

w

3

0

f

L

J

J

U

0

t t

.OM

5.6

’

1.7

t.3

-

12

2.1

- LO

20

. !.I

I

1

I

1

q 5 30

Sp

473

-

Vi

KNOTS

Fig. 15 WVU CC Technology Demonstrator flight performance map with CC blowing efficiency factor C, = 6.6.

At sea level density, CC velocity would be

0.00237776 x (10/144)

)’”=678 ft/s

(7)

and mass flow rate would be rit = pAccVcc= 0.002377 x (10/144) x 678 = 0.1 12 slug/s

(8)

Then the blowing power kinetic energy required is equal to V2

l i z s = 0.112 x 0.5 x 6782 = 25742 ft.lbf/s = 46.8 hp 2

(9)

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

15

To minimize blowing power required," the CC blowing velocity V,, should equal 4 times the flight velocity V,. At a flight speed of 39 kn, V,, should then be: 4 x 39 = 156 kn = 260 ft/s. For T,, = 76 lbf, the CC blowing mass flow rate should be 76/260 = 0.288 slug/s or kinetic power required would be as low as 0.5 x 0.288 x 2602 = 9734 ft.lbf/s = 17.7 hp. This reduction in CC blowing power required demonstrates the advantage of optimizing the ratio V C C / V ~ .

The pilot was quite satisfied with the handling qualities, and surprised how well the direct lift control DLC valve worked to make correction on the glide angle on approach to landing without inducing significant attitude changes. The CC flap deployment and stowing process worked well and required less than a 17 lb change in stick force, as shown in Fig. 16. To significantly reduce the stick force during flap deployment or retraction, Loth'* filed U.S. Patent 4,600,172, which allows converting a Fowler flap into a CC rounded trailing edge flap by only folding out a rounded trailing edge, which also supplies the blowing air (Fig. 17). The BLC suction is sufficiently strong to hold the Fowler flap against the CC pipe without the need for mechanical attachments. The ability to stow away the CC rounded trailing edge for high-speed low-drag cruise is an important aspect for operation with circulation control high lift systems. Slow flight was the most challenging aspect of the flight test program. With the propeller at 135 hp, the aircraft could be slowed to 23.5 kn indicated, which corresponds to 33.2 kn calibrated airspeed. This corresponded to a trim lift coefficient of 5.1 and wing average lift coefficient of 5.6 while blowing at 13 psig. Then there is little or no power to spare to prevent the onset of stall, which always started with a rapid roll and up to 1000 ft of altitude loss. Clearly this represents flying on the backside of the power curve, as

FLAP

FULL OUT

FLAP FOLDING ANGLE /3*

w,T:\yE RETRACTED

Fig. 16 WVU CC Technology Demonstrator shows acceptable trim force during flap folding.

16

J. L. LOTH

Fig. 17 U.S. Patent 4,600,172 allows flap stowing with greatly reduced actuator torque.'*

shown in Fig. 18. Note it takes only half as much power to cruise twice as fast at 70 kn.

IV. 1979 CC Flight Tests with a Grumman Aerospace A-6A The WVU successful demonstration of CC flight on a fixed-wing aircraft motivated the Navy to contract with Grumman Aerospace to convert an A-6A bomber to STOL operation with CC blowing. The challenge of converting an existing large aircraft to operate with CC blowing far exceeded that of building the small WVU CC Technology Demonstrator from scratch. Bob Englar, at NSRDC, began, in 1974, a careful CC wind tunnel test program to cover the

Fig. 18 WVU CC Technology Demonstrator performance safety is limited by the effect of high induced drag on power required.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

17

entire range of operational aspects for the A-6A. His well-publicized test results are currently considered to be the most reliable available data and to which computational solutions are being compared. 13- l7 Of particular interest is Englar’s16 “STOL Potential of the Circulation Control Wing for High-Performance Aircraft.” This contains the performance map of a wind tunnel study of the A-6A wing without a tail, as shown in Fig. 19. When linearized using Eq. (l), the best fit constants are CL, = 0.09 (per degree) and CB = 6.3, as shown in Fig. 20. These results are similar to those found for the WVU CC Technology Demonstrator, although the A-6A has a greater percent of wing area equipped with CC blowing. The drawback of modifying an existing large aircraft is that the CC blowing system had to be an add-on-feature with no possibility for rounded trailing edge retraction to maintain its low-drag, high-speed cruise capability. The magnitude of the CC rounded trailing edge is clearly visible in a close-up photo; see Fig. 21 and in-flight Fig. 22. The STOL performance data for the A-6A were close to those predicted from wind tunnel tests, resulting in 1) 140% increase in usable CL;2) 30-35% reduction in take-off and approach speeds; 3) 60-65% reduction in take-off and landing ground roll; and 4) 75% increase in payload.

-1.4

L,-..

-4

o

4

a

IZ

16

20

24

zs

a in degrees Fig. 19 Wind tunnel test data for wing of Grumman A-6A.

J. L. LOTH

18

4.4

'0'

C+= a!*,'

,,#G= Q,2 3.6

'I

,,,'

,' f

)' h

>

2.8

'5 -I

,'

,#'

w v

-#'

,r'

(CI

C

,' ,'

,/'.

,'8'

,/'

2 <, , ,' #'

# ,'

2 ,4'

0

<, #'

,'

1.2

,*'

'8'

,/I

/' ,f

,i' 0'

,''I

,/ IS

, '

,' /'

0.4

,' *'

'I

',

-0.4 -4

4

-r

12 a in degrees

20

2E

Fig. 20 Empirical curve fit to A-6A wind tunnel data shows blowing efficiency factor Ce = 6.3.

V. Conclusions In 1974, the WVU CC Technology Demonstrator STOL was the first aircraft to demonstrate the high-lift capability of CC. Its wings incorporated an in-flightretractable CC rounded trailing edge to enable high-lift generation by CC blowing on an extended wing, and in-flight conversion to a reduced wing area with sharp trailing edge for low drag, high speed cruise. The use of a retractable CC rounded trailing edge required supplying the hot high-pressure CC blowing air through the rounded trailing edge. To minimize air pressure loss by friction, the duct cross-sectional area had to be at least twice that of its choked-flow area A*. To achieve that, an ejector was inserted inside the CC blowing plenum. The area of its choked flow nozzle was five times smaller than the flow area in its rounded trailing edge. Adding such an ejector provided several other advantages: 1) Its entrainment provided boundary layer suction just upstream of the CC blowing slot, which increased the blowing efficiency factor CB.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

19

Fig. 21 Grumman A-6A CC conversion by an add-on fixed CC blowing duct improving STOL performance at the expense of cruise speed.

2) The entrained flow rate about equaled the nozzle flow rate, thereby doubling the CC blowing mass flow rate. This lowered the Vcc/Vm velocity ratio to increase the blowing efficiency factor CB. It also increased the CC blowing slot height by a factor of four. 3) Ejector entrainment provided wing structure cooling and reduced thermal expansion problems.

Fig. 22 Grumman A-6A during CC flight tests in 1979.

20

J. L. LOTH

4) Ejector entrainment reduced the amount of compressor bleed air required with its associated take-off thrust loss.’o,’’9’8 5 ) During take-off in a jet aircraft, thrust loss associated with compressor bleed, is at least two and one-half times the bleed air momentum 6) In 1979 the second CC Technology Demonstrator STOL aircraft, a converted Grumman A-6, demonstrated excellent STOL performance in terms of increased lift-off weight capability and reduction in required runway length. 7) More research is needed to reduce induced drag associated with flying at high-lift coefficients. Not having to fly on the backside of the power curve would greatly increase safety in STOL flight. Since 1974, numerous other applications for CC blowing over a rounded trailing edge have demonstrated the versatility of this technology. For example: 1) Wake drag reduction behind cars, trucks, torpedoes, etc. 2 ) Propeller downwash drag reduction on tilt rotors. 3) Improved performance of low drag horizontally mounted radiators in cars. 4) Lightweight, hot exhaust gas deflectors on helicopter engines in ground effect. 5 ) Providing an alternative to a helicopter tail rotor, to cancel rotor torque. 6) Noise reduction by wake dissipation on helicopter rotors. 7) Improved effectiveness and control with upper surface blowing (USB). 8) Providing pneumatic control on fixed flight control surfaces. These developments indicate that the future for new CC applications is bright.

References ‘“Wings with Nozzle Shaped Slots,” NACA Translation TM 521, July 1929 (from Berichte Der Aerodynamischen Veruchsenstalt in Wien, Vol. 1, No. 1, 1928). *“The Use of Slots for Increasing the Lift of Airplane Wings,” NACA Translation PW 635, Aug. 1931 (Proceedings L’Aeronautique, June 1931). 3Davidson, I. M., “Aerofoil Boundary Layer Control System,” British Patent No. 913,754, 1960. 4Kind, R. J., and Maull, D. J., “An Experimental Investigation of a Low-Speed Circulation Controlled Airfoil,” The Aeronautical Quarterly, Vol. XIX, May 10, 1968, pp. 170-182. ’Shrewsbury, G., “Numerical Evaluation of Circulation Control Airfoil Performance Using Navier Stokes Methods,” AIAA Paper 86-0286, Jan. 1986. 6Williams, R. M., “Some Research on Rotor Circulation Control,” Proceedings of the Third CALIAVLABS Symposium, Vol. 11, June 1969. ’Williams, R. M., and Howe, H. J., “Two-Dimensional Subsonic Wind Tunnel Tests on a 20% Thick, 5% Cambered Circulation Control Airfoil,” NSRDC TN AL-176, 1070, AD 877764. ‘Loth, J. L., “Some Aspects of STOL Aircraft Aerodynamics,” Business Aircraft Meeting, Wichita, KS, 3-6 April 1973, Paper No. 730328.

COMBINING BLC SUCTION AND CC HIGH-LIFT GENERATION

21

’Loth, J. L., Fanucci, J. D., and Roberts, S. C., “Flight Performance of a Circulation Control STOL Aircraft,” AIAA Paper 74-994, 6th Aircraft Design, Flight Test and Operations Meeting, Los Angeles, CA, Aug. 1974. “Loth, J. L., Fanucci, J. D., and Roberts, S. C., “Flight Performance of a Circulation Control STOL Aircraft,” Journal of Aircraft, Vol. 13, No. 3, 1976, pp. 169-173. “Loth, J. L., and Boasson, M., “Circulation Control STOL Optimization,” Journal of Aircraft, Vol. 21, No. 2, 1984, pp. 128-134. ‘’Loth, J. L., “Retractable Rounded Trailing Edge for Circulation Control Wing,” U.S. Patent No. 4,600,172, issued 15 July, 1986. 13 Englar, R. J., “Investigation into and Application of the High Velocity Circulation Control Wall Jet for High Lift and Drag Generation on STOL Aircraft,” AIAA Paper 74-502, 17-19 June 1974. 14 Englar, R. J., “Circulation Control for High Lift and Drag Generation on STOL Aircraft,” AIAA Journal of Aircraft, Vol. 12, No. 5, 1975, pp. 457-463. ”Englar, R. J., Trobaugh, L. A., and Hemmerly, R. A., “Development of the Circulation Control Wing to Provide STOL Potential for High Performance Aircraft,” AIAA Paper 77578,6-8 June 1977. 16Englar,R. J., Trobaugh, L. A., and Hemmerly, R. A., “STOL Potential of the Circulation Control Wing for High-Performance Aircraft,” Journal of Aircraft, Vol. 15, No. 3, 1978, pp. 175-181. ”Englar, R. J., Hemmerly, R. A., Moore, H., Seredinsky, V., Valckenaere, W. G.,and Jackson, J. A., “Design of the Circulation Control Wing STOL Demonstrator Aircraft,” AIAA Paper 79-1842, Aug. 1979; also published in Journal of Aircraft, Jan. 1981. 18Loth,J. L., “Circulation Control STOL Aircraft Design Aspects,” NASA Circulation Control Workshop, 19-21 Feb. 1986, NASA Ames Research Center, NASA Pub. CP2432, pp. 569-588. ”Loth, J. L., Funk, M. S., “Thrust Savings Limitations with Blown High Lift Wings,” AIAAIAHSIASEE Aircraft Design, Systems and Operations Meeting, St. Louis, MO, 14-16 Sept. 1987.

Chapter 2

Overview of Circulation Control Pneumatic Aerodynamics: Blown Force and Moment Augmentation and Modification as Applied Primarily to Fixed-Wing Aircraft Robert J. Englar* Georgia Institute of Technology, Atlanta, Georgia

Nomenclature Aj = blowing slot area b = wingspan c = chord length cf = flap chord length c d , cD = 2-D or 3-D drag coefficient CDE= equivalent drag coefficient C,, CL = 2-D or 3-D lift coefficient C,,,, = maximum lift coefficient CN = yawing moment coefficient CT = thrust coefficient C y = side force coefficient C , = 2-D or 3-D jet momentum coefficient (see Eq. 1) h, hj = blowing jet slot height L/D,, = equivalent lift/drag ratio m = measured jet mass flux P d , P D = duct total pressure q = freestream dynamic pressure ($pV2) r = circulation control R = universal gas constant Re = freestream Reynolds number, based on chord length c S = wing area *Principal Research Engineer, Aerospace, Transportation, and Advanced Systems Lab., Georgia Tech Research Institute. Associate Fellow AIAA. Copyright 0 2005 by Robert J. Englar. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

23

R. J. ENGLAR

24

S, = ground roll Td = duct total temperature V = freestream velocity vj = blowing jet velocity, isentropic x/c = nondimensional chord-wise location xTO = takeoff distance a = angle of attack y = ratio of specific heats Sf= flap deflection angle Sj =jet deflection angle p = freestream density pj = blowing jet density Ic, = yaw angle (side wind angle)

I. Introduction HE USE of pneumatic devices in the form of blown jet airfoils has been employed or been under consideration in the field of aerodynamics as far back as the 1930s, and perhaps even earlier.’,* In most of these devices, which generally fall into the categories of jet flaps or blown flaps, a jet sheet exits from the trailing edge of the airfoil at a fixed angle or tangent to a flap with a sharp trailing edge. This augments aerodynamic forces by entraining and deflecting the airfoil flowfield pneumatically, rather than solely by deflecting a mechanical surface. These are “pneumatic flap” lift augmentors, which have been shown to be successful if a sufficient onboard source of compressed air is available. The aerodynamic concept now known as “circulation control” (Fig. 1) is a logical follow-on to these devices, with one very important difference, which has made a significant performance improvement. The tangential jet sheet exits over the curved trailing edge of the surface replacing the flap, and this curvature can turn through a full 180deg or more. The jet remains attached to that curved surface because of a balance between the subambient pressure in the jet sheet and the centrifugal force in the jet going around the curvature. Initially, at very low blowing values, the jet entrains the boundary layer to prevent aft flow separation, and is thus a very effective boundary layer control (BLC; see Fig. 1 lift plot). Eventually, as the jet continues to turn, a rise in the static pressure plus viscous shear stress and centrifugal force combine to separate the jet sheet, and a new stagnation point and stagnation streamline are formed on the lower surface. The large flow entrainment rate of this jet and the large deflection of the stagnation streamline produce a pneumatic camber, and thus pneumatic control of the airfoil’s circulation and lift. Although it is a very effective BLC device, the interest in this concept comes from its ability to further augment the circulation and lift, and thus giving rise to the name “circulation control (CC).” Several additional benefits became obvious from early experimental investigations of the concept as a means of lift augmentation:

T

1) Only very small flap size or even nonmoving control surfaces were required. 2) Lift augmentation could be achieved independent of airfoil angle of attack.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

25

TANQENTIAL BLOWING OVER ROUNDED TRAlLlNQ EDGE SURFACE

Fig. 1 Basics of circulation control aerodynamics.

3) Jet turning angle was no longer limited by physical jet exit angle or blown flap deflection angle. 4) Very high force augmentation was generated per unit of input blowing momentum. Roughly 70 years have passed since the very earliest revelation of this type of curved pneumatic device, and a very large variety of pneumatic configurations have been proposed and evaluated. The author has been actively involved with many of these since around 1967. To further expose this wide range of actual and potential applications, this paper will discuss a large number of these pneumatic devices with which the author is familiar from both past and current research, as well as provide an indication of where the use of CC aerodynamics may be heading. It is by no means a complete and exhaustive study of all known efforts, but rather contains representative cases from a wide variety of pneumatic force/moment augmenting and modifying devices. This paper concentrates primarily on fixed-wing aircraft, but CC is certainly not limited to that application alone. The following examples will confirm the multiple uses of CC devices as the following: 1) aerodynamic force and moment augmentors (Fig. 1 shows ACL/CIL= 80, or 8000% return on the invested momentum); 2) aerodynamic force and moment reduction if/when needed (drag in climb out and cruise); 3) aerodynamic moment control and stability augmentation; and 4) aerodynamic device simplifier (moving parts elimination, complexity and weight reduction). 11. Coanda Effect The CC concept is actually based on the now well-known Coanda E f f e ~ t , ~ - ~ named after the Romanian inventor Henri Coanda, who claimed to have discovered it in Paris prior to 1935. There is a Romanian postage stamp (and

26

R. J. ENGLAR

associated story) showing that Coanda had originally used the Coanda device for a totally different purpose: as a means to deflect the exhaust of a radial piston engine away from a wooden aircraft fuselage. During its first flight, these shielding plates actually entrained the hot exhaust flow inward, igniting and destroying the aircraft. Figure 2 shows the basic Coanda device as later formulated by him (after the fire exhaust incident) and its application to a fixed-wing aircraft (which in this case appears to be a form of BLC device). Note that in these (and in all other Coanda cases found), Coanda aligns acuteangle “steps” downstream of one side of a jet nozzle to deflect the jet to that side and entrain large masses of fluid from the opposite side. The distinctive steps and angles were intended to generate a separated vortex flow at each corner, and thus enhance mixing there. The concept was applied by Coanda to many other devices, including car engine exhaust scavengers, wind-tunnel turning vanes, thrust augmentors, water propulsion units, injection wind tunnels, deflection surfaces, and rotary pumps. However, efficiency questions arose because of added friction along all the steps and separated flow at each corner. Nevertheless, the concept forms the basis for the present CC aerodynamics; an infinite number of small-angled steps simply becomes a continuous curved surface with even greater entrainment capability and less energy loss due because its lack of discrete corners. The following discussion will present a number of favorable applications of CC aerodynamics, where the governing difference between CC and the jet flap/blown flap will be the continuously curving surface downstream of the tangentially blown jet, with force augmentation/modification being mainly a factor of jet blowing parameters, not the angles of the sharp flap trailing edge or the jet angle relative to the chord line. The main emphasis here will be on fixed-wing devices and applications. Application of CC to rotary-wing aircraft offers many additional benefits, as discussed in Refs. 6 to 9. These are based on

Original Coanda Device, Approx. 1935

Fig. 2 Coanda devices and high-lift, low-drag wing?

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

27

high-lift generating capability independent of angle of attack, as Fig. 3 shows, thus eliminating the previously required cyclic and collective pitch blade mechanisms. Before proceeding, it is important to define the blowing momentum coefficient C, as

where the last definition only holds for two-dimensional incompressible flow (pj = pm). Typically, the jet velocity in ft/s is calculated from isentropic relationships as:

where the subscript d implies total conditions in the blowing plenum duct, subscript 03 is freestream, R = 1716 ft2/(s2 OR), and y =1.4 for air. A jet

X

U

z

THICKNESSICHORD, tlc

Fig. 3 Maximum lift of blown circulation control airfoils.

R. J. ENGLAR

28

expansion to the actual static pressure just outside the jet slot would yield higher calculated values of vj and thus C,, but would vary as the external flow conditions or shape changed, so would be hard to duplicate as a universal design parameter. Mass flow m is almost always measured under test conditions using appropriate flow meters, but can be calculated isentropically as well using compressible flow relationships." Before going any further, please note that there is nothing that prohibits the jet velocity from being supersonic unless the geometry is such that a shockdown back to subsonic flow causes the jet to detach from the curved surface. It will soon be seen that it is often advantageous to have a higher speed jet than a lesser speed one. The momentum term mvj can, of course, also be thought of as a jet thrust. 111. Applications of Circulation Control, Past and Present

A. Circular Cylinder Stopped-Rotor Aircraft An early application of CC was developed by the British National Gas Turbine Establishment (NGTE) in the mid-l960s, when it was desired to produce a stoppable-rotor VTOL aircraft. In this concept, a blown two-bladed rotor could produce very high lift per blade just to get the aircraft to hover, then be stopped and stowed within the helicopter fuselage for forward fixed-wing flight.697A circular-cylinder cross-section slotted-pipe rotor appeared to be an ideal solution, because, as Fig. 3 shows, its thickness/chord ratio of 1.0 presents the possibility of C, = 4.rr if flow can be made to stay attached. As the figure shows, values even greater than 4.rr were generated by blown CC cylinder rotor blades when excess thrust in the vertical direction (the jet flap effect) was included at higher t / c values. However, the high drag of a 100% thick circularcylinder airfoil proved to be a difficult problem and reduced the aerodynamic efficiency of these airfoils to unacceptable values. A similar circular lifting surface' was also pursued by NASA Langley in the 1960s to provide lift on takeoff and landing by blowing on the circular fuselage cross-section of a hypersonic aircraft, as well as for return after launch of missile or rocket boosters having circular cross-sections. Whereas lift coefficient values over 20 were measured at very low Reynolds numbers for an end-plated-cylinder tunnel model with multiple slots, a single-slotted cylinder produced Cl = 18 at C, = 6. This lift augmentation of only three times the input C, implied the need for a large air supply. The associated drag coefficient of over 9 gave a lift/drag ratio of only 2, or even less if the blowing coefficient were added to the drag to yield an equivalent drag coefficient. Clearly, high lift was available, but the lift-associated drag and required blowing coefficient posed serious problems.

B. Elliptic-Airfoil CC Rotor As interest in circular cylinder CC blades for helicopters was lessening in the United Kingdom, it was rising dramatically in the United States in the late 1960s as a possible means to increase rotorcraft performance while greatly simplifying the entire rotor system mechanical hardware. The US effort was centered at the

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

29

Navy’s David Taylor Naval Ship R&D Center (DTNSRDC), where the approach taken was to develop lower-drag, high-lift rotor blade sections by converting the circular cylinder profile into a much thinner blown elliptic airfoil. These efforts also became the basis for fixed-wing efforts, and are presented here to clarify understanding of these CC pneumatic devices. Figure 4 shows several such single-slotted CC rotor elliptic airfoils, where the obtainable Cl is lower than for the cylindrical airfoil, but the required C, is a factor of 10-20 less. Note that this performance is all at angle of attack a = 0 deg, providing a nonpitching alternative to both the mechanical cyclic and collective angle of attack variation required of conventional rotor blades. Note the very high force augmentation, ACl/C, of 80, representing an 8000% return on the momentum invested. Also shown for comparison is a typical 30-deg jet flap applied to a 15% thick ellipse airfoil; the greatly reduced force augmentation of the jet flap is evident because the jet exits from the lower surface of the airfoil at a fixed angle. It should be clear that CC is not a jet flap, but achieves its high-lift capability because the stagnation stream line movement and resulting circulation can be controlled and increased well beyond that of a sharp trailing edge. Figure 5 shows the equivalent lift-to-drag ratio of sample elliptic CC airfoils, where the

YOMEMUM COEFFICIENT C

Fig. 4 Typical blown-lift capabilities of two-dimensional CC elliptic airfoils at a = 0 deg.

R. J. ENGLAR

30

equivalent drag in the denominator now includes a severe penalty for compressor power required as well as for intake ram pressure.lo9l1Maximum equivalent LID values roughly 6-7% greater than the conventional unblown rotor blade NACA 0012 airfoils (varying only a ) are seen for the 20% CC ellipse (at a = 0 deg), but at a lift coefficient 30% higher, at about 1.3. Furthermore, the Cl can be increased up to 6 or 7 if desired, but at a lesser LID,,. For additional comparison, if the equivalent drag is defined as merely adding C, to the measured drag (i.e., CDE= Cd C,), then LID,, values of over 120 at Cl = 2.5 are possible, all at a = 0 deg (almost three times the Cl of the 0012 airfoil at stall). The efficiency and simplicity of CC was obvious from these tow-dimensional airfoil results, and a serious effort to develop these CC airfoils was undertaken. Reference 12 summarizes much of this Navy effort at DTNSRDC for the years 1969 through 1983, as well as providing a summary of CC-related research conducted by other agencies (in the United States and abroad) outside the Navy from 1956 to 1983.

+

X

3'

a

Y

Pd

SECTIONAL LIFT COEFFICIENT, C/

Fig. 5 Equivalent efficiencies for CC and conventional two-dimensional airfoils.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

31

In 1979, a CC Rotor flight demonstrator based on a Kaman H-2 helicopter was flown with pneumatic aerodynamic and control systems replacing conventional mechanical cyclic and collective blade pitch. 13- l5 Whereas this flight vehicle was hindered by control system response phasing problems, which limited its flight test envelope, it did demonstrate the ability to substitute pneumatics for mechanical blade lift and control devices for hover and forward flight. It also led to the possibility of higher harmonic control of helicopters, where cyclic lift variations at frequencies higher than one per revolution were possible to eliminate rotor-induced vibrations. The absence of blade collective and cyclic pitch links is possible; they can be replaced by internal control cams or valves to vary blowing pressures.

C. Circulation Control Airfoil Development Considerable CC airfoil development was ongoing at this time, both experimentally and analytically. A number of CFD techniques using various Navier-Stokes codes have been developed and used to understand the relevant viscous flowfields. These will not be discussed here, but can be found summarized in much more detail in Refs. 12 and 16. A typical example of CFD-calculated streamlines and velocity vectors” is shown in Figs. 6 and 7 for a generic flat-sided semi-elliptic CC airfoil. Of particular interest here are the computed velocity vectors and streamlines downstream of the slot on the blown trailing edge (Fig. 7), where the stagnation point of the jet sheet appears to be turned nearly 130-140 deg from the jet exit. A considerable number of additional CFD analyses, both subsonic and transonic, have been conducted by various

investigator^.^^,^^,^^

Fig. 6 Computed streamlines for simplified CC airfoil (a= -2 deg, Cl = 4.6).”

32

R. J. ENGLAR

Fig. 7 Computed velocity vectors and streamlines," CC airfoil.

A number of experimental programs have also been conducted to understand the CC phenomenon and the details within the blown curved surface region. Two-dimensional laser-velocimeter measurements at Lockheed" for the same CC airfoil as in Figs. 6 and 7 showed mean velocities that confirmed the CFD results already presented. Again, jet flow turning to a separation point/stagnation streamline approximately 130-145 deg from the slot was seen (Fig. 8). Experimental investigations by the present author" of a very similar generic airfoil (Fig. 9), used surface static pressure, static pressure across the jet, and a rotatable hot-film shear stress probe to measure the actual separation point location (where

Fig. 8 CC velocity vectors recorded by Lockheed Laser Doppler Velocimeter."

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

33

G 2

w

0 U U

8

t 3

MOMENTUM COEFFICIENT, C, NONDIMENSIONAL CHORDMSE STATION. x/C

Fig. 9 Two-dimensional semi-ellipse CC model geometry, plus measured lift and static pressures as functions of C, and slot height.

shear stress = 0) as a function of blowing and slot height. The resulting C,and C, distributions are seen in Fig. 9. As Fig. 10 shows, jet turning as high as 170175 deg was measured for this airfoil. At a constant C,, greater turning occurred with a smaller slot height because the resultant jet velocity and entrainment are higher as jet area reduces. Figure 9 (left plot) shows that this greater velocity and jet turning clearly results in generation of higher Cl, where values close to nine are possible at a = 0 deg (although tunnel flow impingement occurs here). Figure 9 (right plot) presents associated static pressure distributions on the airfoil. These analytical and experimental data confirm the effectiveness of blowing to greatly deflect the entire flowfield and then strongly increase the circulation and lift on these very generic airfoils, to the point that very high lift is produced without wing flaps and slats and at Odeg angle of attack. Some additional information on generic CC airfoil performance is provided in Ref. 20. One last note on CC airfoil performance: as previously mentioned, smaller slot height yields a larger return in Cl at constant C, than does a larger slot height, primarily because of greater vj/Vm and extra flowfield entrainment. Figures 9 and 10 show this trend. However, if the static pressure coefficient just outside the slot exit (C,,,)is known or can be determined, a new parameter defined in Fig. 11 can be used (when vj is expanded to this local condition to yield CBLc)to collapse the different slot height results (left) into a single curve (right).

R. J. ENGLAR

34

Fig. 10 Blowing jet separation point location measured by hot-film shear stress probe.

c,

%LC

CONVENTIONAL MOMENTUM COEFFICIENTS {EWANDED TO fREESTREAM CONDITIONSI

LOCAL MOMENTUM COEfflClENlS IEXPANDED TO LOCAL WNDITIONS AT SLOT)

Fig. 11 Comparison with momentum coefficients based on local jet exit static pressure (right) and variation with slot height.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

35

D. X-Wing Aircraft An extraordinary use of unusual CC airfoils is the X-Wing VTOL ~ e h i c l e , ~ al - combined ~~ rotary/fixed-wing aircraft equipped with a fourbladed rotor, which was designed to take off and hover with the same nonmechanical cyclic and collective benefits as those already described. However, forward flight at speeds roughly twice the limit on conventional rotors could be achieved using a “reverse velocity” blown rotor/wing concept (Fig. 12). Typically, as vehicle speed increases, the retreating blade of a rotor sees a resulting velocity that is the difference between the vehicle forward speed and the blade rotational velocity; this can rapidly become a reverse flow at the blade trailing edge, an unacceptable region that moves further outboard as speed increases. Lift on that “stalled” blade segment can actually be negative; the rotor might not be trimable in roll, and drag increases dramatically. The X-wing avoids this problem at high speeds by employing CC on each end of the blade (Fig. 12), and a “clever” control system can blow whichever slot is currently on the airfoil’s trailing edge. Thus, the airfoil never experiences flow from the “wrong” direction. The entire system can be simplified even further by use of simultaneous blowing from both leading and trailing edges of the double-ended airfoils24 (Fig. 12, right). Note that even if the flow is coming from the wrong direction (dashed curve), the dual-slotted airfoil still yields 80-90% of the single-slotted airfoil’s lift, even when the leading edge is counter the conventional direction (compared to little, zero, or negative lift from a conventional airfoil). This allows rotorborne flight at a much higher speed until eventual conversion to a fixed wing in an X-configuration is achieved, with the representative TE slots on each blade of the now fixed wing being used for roll and pitch control without moving surfaces. This concept was actually “flown” full-scale in the NASA Ames 40 x 80 ft tunnel and successfully completed the transition from hover to stopped-wing using pneumatics. Two representative configurations from Refs. 16 and 25 are shown in Fig. 13.

DUAL PLENUM AIRFOIL SECTION

HELO DIRECTION

Fig. 12 Dual blowing on a reverse velocity rotor2’ and blown lift of dual-slotted CC airfoils.

36

R. J. ENGLAR

LIGHTLY LOAOEO LOW TIP SPEED TAIL F A N

Fig. 13 X-wing rotor configuration with rotor stopped, and control systems.

E. Circulation Control Wing (CCW) The high-lift capability independent of angle of attack, which was demonstrated by the CC rotor airfoils above, led to the application of CC as a simplified very-high-lift device for STOL aircraft. The airfoil in Fig. 1 is representative of this simplified pneumatic concept, where both the mechanical TE flap and the LE flap or slat have been replaced with nonmoving pneumatic systems. Primary development of the concept took place in conjunction with CC rotor development efforts at the Navy's DTNSRDC12,16,26-30in the late 1960s to early 1980s. Initially, the concept was modeled as a small add-on device" that would convert the wing flap's sharp TE into the round CC Wing (Fig. 14, right), which was tested at DTNSRDC in specialized two-dimensional high-lift test facilities." Compared to results from a family of more conservative blown flaps (Fig. 15, from Ref. 30), the CCW profiles showed two significant advantages. They could generate greater Cl than the blown flap, because of much greater streamline displacement, and had no sharp TE to limit streamline turning, or for the same chord-length device, could generate the same incremental Cl at much less C, required. An alternative 180 deg rotatable CCW TE is also shown in Fig. 14 (left), which, although it may be mechanically simpler, pays the penalty of losing wing area in the blown high-lift mode. Numerous two- and three-dimensional wind-tunnel evaluations and feasibility studies led up to the flight test of a fixed CCW device on an A-6/CCW STOL demonstrator a i r ~ r a f t ~in l - 1979. ~ ~ Flow visualizations in Fig. 16 show a full 180 deg of jet turning on a static 1/8-scale model of the test aircraft in the DTNSRDC tunnel, and Fig. 17 shows the CCW installation on the fixed flap of the A-6 flight-test aircraft. Because this was a proof-of-concept flight test, the CCW device was not retractable and the air supply lines were mounted externally and cross-ducted in the fuselage, where they connected to the high-pressure

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

37

CONVENTIONAL

Fig. 14 Retractable/storable CCW trailing edges.

bleed ports of the standard J-52-P8A turbojet engine. Results using only available bleed air from the engines confirmed maximum C, values 120% greater than the conventional Fowler flap, or, even more applicable, 140% increase in the usable lift coefficient at takeoff/approach angles of attack. Also confirmed were 3035% reductions in the takeoff and approach speeds resulting in 60-65% reductions in takeoff and landing ground roll distances, and yielding values as short as 600-700 ft. This full-scale confirmation of CCW also implied that

FLAP CHORD LENGTH

MINIMUM Cp REQUIRED

Fig. 15 Comparisons between CCW and blown flap airfoils at a = 0 deg.

38

R. J. ENGLAR

Fig. 16 CCW jet turning on the Ad/CCW wind tunnel model at DTNSRDC.

there was sufficient extra CLgenerated to increase the liftable payload by 75% if the conventional takeoff ground roll distance were used. Also shown was that the additional lift-induced drag resulted in much steeper glide slopes on approach, where higher engine power settings (which could also be used for quicker response during waveoff) were offset by this excess drag.

Fig. 17 A-6/CCW STOL flight demonstrator aircraft.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

39

Fig. 18 WVU STOL demonstrator CC airfoil?'

A smaller CCW demonstrator based on a prop-driven BD-4 general aviation aircraft had been flown earlier by West Virginia University (WVU).'6,35,36In the flight-tested configuration, the CC blown cylinder was mounted at the TE of a hinged flap that rotated 180 deg aft to increase the effective high-lift area by 20%, and included a boundary layer control (BLC) suction slot at the flap hinge upper surface (Figs. 18 and 19). The blowing air was supplied by an onboard 200 hp compressor (APU), which provided enough air for blown ailerons in addition to the CCW. The section lift coefficient on the blown CCW wing section was increased by a factor of nearly 2.5 with blowing. Wing downwash on increase to a factor of 1.92, but provided threethe tail reduced trimmed C,, dimensional lift augmentations of ACL/C, = 15.2, a significant increase should the required airflow be available from a general-aviation aircraft engine, for example, if using a supercharger or turbocharger.

Fig. 19 WVU BD-4 based STOL demonstrator air~raft.3~

40

R. J. ENGLAR

CORRECTED

mnm. F&, MS x 1

6

Fig. 20 Thrust performance of J-52-PSA turbojet engine with bleed.

Both of these fixed-wing flight programs demonstrated the feasibility of CCW as an operational STOL system in terms of high lift, short takeoff and landing, and simplicity, but also identified issues still to be resolved. Among these were the drag of the device in cruise flight (WVU solved this but at the cost of a mechanical 180-deg rotating flap that stowed in the aft wing cavity, and GTRI solved it with the dual-radius airfoil discussed in Sec. 111.F.) and, of course, the need for an onboard air source. Figure 20, which presents turbojet engine ground test data31 taken during the A-6/CCW program, show that the airflow acquired from highpressure compressor bleed ports could be increased up to three to four times that of the standard engine spec bleed limit without overheating, but obviously at the cost of takeoff thrust lost. Similar data for turbofan engines show that engine core bleed is much more costly in thrust loss (although lower-pressure fan bleed is possible), and thus the idea of an ejector to trade excess pressure for extra mass flow appears feasible. However, the need to reduce CCW drag in cruise is a necessity for operational aircraft.

F. Advanced CCW Airfoils DTNSRDC and Grumman took two approaches to the drag p r ~ b l e m , ~ ~ ’ ~ ~ - ~ one a fixed simple radius reduction and the second a very-small-chord deflectable CCW flap. From the nondeflectable ~ t a n d p o i n ta, ~supercritical-type ~ airfoil was employed as the baseline because it already had a bluff base thickness between 0.005~and 0.010~.CCW rounded and semiround (96 deg arc instead of 180 deg)

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

41

designs were tested, including a series of smaller radii, looking for reduced drag without loss of lift augmentation. The CCW/Supercritical airfoil was developed primarily from a low-speed standpoint, where a TE radius of 0.009~was found to produce very little drag penalty yet have superb lifting capability. Figure 21 shows its lift curves at constant C, compared with a family of mechanical multi-element flaps. Not only do the no-moving-parts CCW airfoils generate the same or greater lift as the maximum Cl of a triple-slotted-flap airfoil with mechanical slat, they also do so at a = 0 deg. Note that the large leading edge of the supercritical airfoil provides a natural nonmoving LE device, which generated similar stall angles to the mechanical slat. One further benefit is the cruise drag polar (Fig. 22), which is slightly higher unblown than the baseline supercritical airfoil, but with very slight blowing can reduce c d to less than the baseline while also increasing lift, both at constant a. For clarity here, measured c d includes C, because the wind-tunnel balance cannot easily separate blowing thrust from drag. That is why Fig. 22 shows negative drag recorded with blowing. This is accounted for in the equivalent drag term, LID,,, where C, is included (see Ref. 10 for a more detailed explanation). Additional benefits of blown CC airfoils at speeds up to transonic were shown in the compressible flow tests of Ref. 40, where blowing was seen to produce a very favorable boundary layer/shock interaction, drag reduction, and increased Cl (Fig. 23). The second approach to the drag problem was a simple CCW flap with a curved upper surface and a sharp trailing edge (Fig. 24, from Refs. 38 and 41). Here, a short-chord flap (less than 0 . 1 0 ~ pivots ) about a hinge on the lower surface and exposes a smaller-radius CCW surface downstream of the tangential slot. This radius is approximately the airfoil thickness at the slot location, less the slot height. The upper surface of this flap is a second arc of much larger radius, the radius being chosen to keep the arc close to the airfoil aft contour. As the small flap is deflected on this dual-radius CCW airfoil, the large radius produces an arced CC aft surface with a turning arc much larger than the flap deflection angle.

MULREWEHT MECHANICAL ~ W F T AIRFOILS

NDWVHG-PART CCWlsUPERCRlTlcAL AIRFOIL

Fig. 21 CCW/Supercritical, dual-radius CCW, and conventional mechanical flap airfoil comparisons.

42

R. J. ENGLAR

Fig. 22 Low-speed drag polars for CCW/Supercritical airfoil.

Freedream Mach Amber, Iy,

Fig. 23 Transonic lift caused by blowing for three pneumatic ellipse airfoils!'

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

43

Supercritical Contour

Fig. 24 Dual radius CCW airfoil with LE blowing.

For the 90-deg flap shown here, the jet turning angle is about 135 deg (compare to Figs. 8 and lo), limited by the TE comer. With flap retracted to 0 deg, the airfoil is in a sharp-TE cruise configuration. The slight mechanical addition provides unblown camber as well. The leading edge employs an inverted tangential slot to replace any mechanical flap there. Lift data for the 90 deg flap configuration are also shown in Fig. 21, where Cl increases of 35% over the CCW/Supercritical airfoil occur, with considerably greater increases over the conventional flaps. Figures 25 and 26 also show additional advantages of this configuration: the ability to dramatically interchange lift and drag as the small-chord CCW flap

2-DCCW SUPERCRITICAL AIRFOIL, DUAL- RADIUSFLAPS, DRAG POLARS, THE PENALTY FOR LIFT??

Fig. 25 Drag polars of CCW dual-radius airfoil at various flap angles.

R. J. ENGLAR

44

2-D CCW/SUPERCRITICAL AIRFOIL, DUAL-RADIUS CCW,

u, degrees

Fig. 26 Dual-radius 90-deg flap CCW airfoil lift as functions of

(Y

and C I * ~ ~ .

is deployed (Fig. 25) and the increased lift and stall a as LE blowing is activated (Fig. 26). The thrust/drag interchange in Fig. 25 implies the potential for high lift and drag for STOL approach (remember, induced drag due to high lift is not included in this data for two-dimensional airfoils) or high lift and reduced drag for takeoff. Figure 26 shows the capability of this nonmoving LE device to reattach flow, prevent stall, and dramatically increase C , . Figure 27 combines the above data in terms of Cl vs LID,, where the equivalent drag coefficient is defined as CDE= Cd C, to account for the blowing required to yield these drag changes. These data include four CCW flap angles and various LE and TE blowing values. Figure 27 includes a locus of achievable Cl vs the associated efficiencies in comparison to the clean cruise airfoil (flap = 0 deg, C,E = 0, , = 0). This plot confirms the ability of CCW airfoils to generate very C high lift and associated drag (reduced LID,) for approach, plus much higher L / D e at somewhat lower Cl for takeoff and climbout. Because of the latter, the 30-deg CCW flap at reduced C, appears to be an excellent configuration for takeoff. One additional benefit results for the 0-deg flap CCW case. When in cruise, drag is low because of the sharp TE, but should blowing be initiated without flap deflection (Fig. 28), significant lift is generated by the flap curvature, while drag reduction occurs due to thrust recovery. Note the comparison to the NASA Energy Efficient Transport slotted, flapped airfoil. Not only is lift

+

Next Page OVERVIEW OF CC PNEUMATIC AERODYNAMICS

45

CCW Dual-Radius Airfoil DRAG POLARS: the Penalty for Lift?

2-D Lift Coefficient, C,

Fig. 27 Lift and equivalent efficiencies of dual-radius CCW airfoil.

greater for the CCW cruise airfoil, the drag polars move into the thrust recovery region. From these results, one can also immediately realize the potential of this high-lift system as a nonmoving roll/yaw device. Blowing only the undeflected right wing’s flap will produce a lift (roll with right wing up) and favorable yaw (nose left), thus yielding favorable roll/yaw coupling from a nonmoving surface, instead of the usual adverse roll/yaw coupling from a conventional aileron. A study41 was conducted for NASA Langley Research Center to evaluate the effectiveness of applying this concept to an Advanced Subsonic Transport. Here, the dual-radius CCW of Fig. 24 was applied to a 737 wing characterized in Fig. 29. The typical 15 moving elements per wing were replaced with the CCW single element flaps and LE blowing, yielding perhaps a maximum of three components per wing (the outboard CCW flap became the aileron, and blowing differentially on the CCW flap replaced the spoilers for roll). Using only fan bleed air (and the associated lower thrust lost), replacing the conventional flaps with CCW was able to triple the usable lift at takeoff and produce the ground roll reductions shown in Fig. 30. For lighter aircraft weights, blown takeoff rolls of 400-500 ft are possible with 0 kn headwind, about a third that of the conventional aircraft; with a 20 kn headwind (wind over deck), 200-300 ft

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46

R. J. ENGLAR

Fig. 28 Comparison of cruise dual-radius CCW (Odeg flap) with mechanical flap airfoil.

737 WlNOlFLAPICONTROL SYSTEM

Fig. 29 Pneumatic airfoils simplify wing complexity.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

47

WODIO.0 kt.

Qrou WoIgM; lb

Fig. 30 737/CCW takeoff ground rolls, sea level, at 0 kn headwind.

rolls were predicted for the 737/CCW configuration. The tradeoff in increased liftable weight at constant ground roll is possible here as well. For the conventional 737’s ground roll of 1350 ft at W = 70,000 lb, the 737/CCW can increase the liftable weight to near 100,000 lb with the same runway length, a 43% increase in gross weight (fuel, payload, etc.). Climb angle over the 50-ft obstacle is slightly less for the CCW aircraft41than for the conventional one. These pneumatic benefits are not limited to subsonic transports. Reference 42 reports on the application of CCW to a generic High-speed Civil Transport (HSCT) configuration developed by GTRI for NASA Langley Research Center. Here the intention was to reduce takeoff wing area required (and thus increase cruise performance) by employing simplified CCW flaps and trading increased lift coefficient for reduced wing area. The extra lift would also allow a reduction in takeoff/approach angle, and eliminate the need for fuselage nose droop, aft fuselage upsweep, extended nose gear, and/or synthetic vision. Reference 42 reports 90-deg jet turning on the CCW flap, over 100% increase ,,,,C , and 45% increase in stall angle. High-lift generation from various in blown devices and blown canards is shown in Fig. 31. Note the effect of the blown canard on stall a. In-house feasibility studies43 have been conducted and confirm this potential for pneumatic wings/canards on high-speed aircraft.

R. J. ENGLAR

48

Angle of Attack,a,deg

Fig. 31 Lift augmentation on the GTRI HSCT/CCW semispan model with blown

IV. Powered Lift and Engine Thrust Deflection A. CCW/USB Mechanical flaps have been employed to entrain and deflect thrust from engines mounted on the wing upper surface (upper surface blowing, USB), and it was envisioned that the entrainment capabilities of CCW could do the same without the mechanical complexity; thus the CCW/USB c o n ~ e p t ~ ' , ~was ~-~* born (Fig. 32). Subsonic wind-tunnel investigation^^^ at DTNSRDC showed no-moving-part pneumatic capability to entrain and turn USB engine thrust well past the 60 or so degrees of a mechanical USB system, but also continuing through 90 deg, and then rotating the thrust forward as a thrust reverser through 165 deg (Fig. 33). The possibility then exists for high lift and thrust reversing all in one system just by varying the CCW blowing rate, with a possibility of VTOL in between (depending on installed thrust levels). Wind-on data (Fig. 34) show very interesting lift-drag polars at a = 0 deg, with the ability to vary lift and drag by blowing alone, independent of angle of attack. The enhanced lift capability is far more than mere thrust deflection (i.e., ACL = CT [sin a,, a]).It

+

WRIPBLEMUSTOEFW3lD+l UJETOCC.PLENUHPFESWW

U A R m

Fig. 32 USB and CCW/IJSB powered-lift concepts.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS T = ENQINETHRUST (LB) 01 * ANQLE OF ATYXK q = DYNAMIC PRESSURE

49

L! 0 (CCW ALONE)

P, = SLOT PRESSURE mV, = SWT MOMENTUM FT T + IhVi T. = RESULTAKTTHRUST

0 25.4 1,LB 0 48 6T L5

Fig. 33 CCW/USB model static thrust deflection by blowing only.

results from the increased velocity from the engine exhaust being entrained onto the blown lift surfaces, and the greatly increased circulation lift beyond the powered wing only. A full-scale ground test was performed by the present author at NASA Ames with the CCW/USB mounted behind one engine of the NASA Quiet Short-haul

THRUST

DRAG

Fig. 34 CCW/USB model lift-drag polars, (Y = 0 deg.

50

R. J. ENGLAR

Fig. 35 CCW/USB test assembly on the QSRA aircraft.

Research Aircraft (QSRA), with the aircraft mounted on a force b a l a n ~ e . ~ ~ , ~ ’ - ~ l Figure 35 shows the installation behind the left inboard engine of the QSRA. Thrust deflections as high as over 100 deg were recorded behind this single operating engine. These data are expected to improve if the two engines per wing are operated together and the two exhaust sheets converge for even better turning. As a result of this test, Navy feasibility studies46were conducted for a sea-based turbofan-powered STOL aircraft using both CCW and CCW/USB (Fig. 36). These studies, based on the preceding powered model wind-tunnel tests, showed takeoff ground rolls of 100-200 ft (Fig. 37), varying with weight, blowing, and thrust levels, and resulting from powered C, values of 8-9.

Fig. 36 Proposed CCW/USB Navy STOL aircraft.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

51

GROSS WEIQHT. LBllOm

Fig. 37 Takeoff ground rolls for proposed CCW/USB Navy STOL aircraft.

B. CC/Jet Deflection In a related effort,52CC entrainment was also applied to high-performance aircraft to yield thrust deflection for much higher engine exhaust velocities, where lesser jet turning could still provide excellent STOL potential due to higher thrust/weight ratio. An example is shown in Fig. 38. In-house unpublished experimental work by the present author provided similar studies, where we were able to deflect supersonic jets from rectangular nozzles by more than 80 deg, using blowing jet momentum values around 10% of the engine thrust.

BLOWING WY)ENTW,

&I,

US.

Fig. 38 Pneumatic thrust deflection of rectangular jet exhaust (left, from Ref. 52) and unpublished static test results of a similar configuration (right).

52

R. J. ENGLAR

C. Pneumatic Channel Wing A configuration using similar thrust deflection capability of a CC trailing edge has recently been under development by GTRI for NASA Langley Research Center. Called the Pneumatic Channel Wing (Fig. 39), it employs blowing at the TE of a 180 deg channel (similar to the much earlier but unblown Custer Channel Wing) to entrain the propeller’s thrust, augment the velocity in the channel, and thus generate high-powered lift. Figure 40 (from Refs. 53 and 54) shows typical GTRI wind-tunnel lift data as a function of both blowing and thrust compared to the baseline unblown channel wing configuration, where untrimmed CL,, is increased by a factor of over 7 to a value of 10.5- 11. Reference 54 shows predicted takeoff ground rolls of less than 100 ft on a hot day at 3000 ft altitude using wing angle of attack of only 10 deg. Further details of this concept’s capabilities and the associated data are found in the paper by Englar and Campbell in this volume, and also in NASA CP 2005-213509.

Fig. 39 Conceptual pneumatic channel wing and semispan model in GTRI MTF tunnel.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

53

a,degrees Fig. 40 Pneumatic channel wing lift from thrust and blowing?4

V. Other Aircraft Applications

A. CC Propeller In a manner somewhat similar to the CC rotor above, CC airfoils have also been incorporated into general aviation propeller designs to replace complex and expensive mechanical variable-pitch blades with fixed-pitch pneumatic blades that change aerodynamic and thrust characteristics through mass flow variation to each blade. Figure 41 shows a proposed application, where the propeller blade airfoil is the CCW/Supercritical type of Fig. 21. References 55, 56, and 57 discuss feasibility studies, and concluded that such a pneumatic variable-pitch propeller was possible and held interesting promise depending on the details and costs of an air compressor (such as an aircraft supercharger or turbocharger) to supply the blowing. The study also envisioned supersonic jet blowing to be a possible problem, but much of the CCW data already presented above have blowing pressures well above choked (sonic).

54

R. J. ENGLAR

Flow Control Valve

Compressor

Fig. 41 Circulation control propeller system.

B. Moment Control, Stability Augmentation, and Induced Drag Reduction The preceding data and applications show the ability to pneumatically augment or modify lift and drag without use of moving parts (except possibly very short chord dual-radius CCW devices) and with a high rate of return on input jet momentum. The application to a pneumatic rudder or even winglets can provide side force generation as well.58 It should be obvious that augmenting the aerodynamic force capability of any control surface by blowing can also either increase the control power or reduce the required area of the device, with associated benefits including maintaining stability levels but reducing cruise drag. A few further and less obvious examples of pneumatic control devices will now be mentioned; many others can be found in Refs. 12, 16, and 58. The aft suction peak downstream of the CCW slot (usually at 95% chord or greater) produces very large nosedown pitching moment, which, besides having to be trimmed, also produces greatly enhanced longitudinal pitch stability. In fact, the A-6/CCW flight demonstrator had such large negative values of dCM/dCl that the center of gravity of the flight-test aircraft was moved aft by an additional 10- 15% chord to aid in trim, and the aircraft still had greater longitudinal stability than the conventional A-6, flaps A clever application of CCW for moment control is shown in Fig. 42 on a forward swept wing.59 Previously, increasing blowing on an aft-swept trailing edge pulled the center of pressure (CP) outboard and aft, but during this tunnel e~aluation,~’ the CP was made to move outboard and thus forward with blowing. The amount of xcp movement and the resulting moment were controlled by which segments of the TE slots were blown, and by how much. Pneumatic roll control by differential wing blowing can produce phenomenal rolling moment increments (Fig. 43), where only one wing of a CCW configuration is blown. A second innovation is also shown here: letting the slot continue around the wing tip. Now, high suction peaks at the wing tip, having a maximum moment arm, can yield even greater rolling moment. For reference, a conventional 0.20-chord aileron deflected down 30 deg on the outboard 50% span of

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

55

WEIGHT FLOW-0.25 Iblsec

Fig. 42 CCW applied to forward swept wing for pitching moment reduction.

this wing produced an incremental hCrol1= -0.03. Figure 44 suggests one further advantage of the CC wing tip, where blowing down around the tip directly counteracts the tip vortex rollup and relocates it further outboard, creating an effective aspect ratio increase. Figure 44 shows the effective drag reductions due to tip blowing*l on an already high-aspect-ratio CC rotor blade. At the higher C, values where induced drag usually dominates, C , reductions of 1719% are seen, with greater percentage reductions at lower CL. Lower aspect ratio aircraft wings using this technique should yield even greater CD reduction. One can also alter the spanwise lift distribution with spanwise tapered blowing to approximate an elliptic distribution, and thus minimize induced drag both in cruise and during climbout. In 1986, the present author experimentally applied blowing from a tangential slot along the nose of a generic high-a vortex-lift configuration, thus turning the

Coefficients based on full span and area

Fig. 43 Roll due to CC wingtip blowing, (Y = 0 deg.

56

R. J. ENGLAR

WNG LIFT COEFFICIENT

Fig. 44 Tip blowing for induced-drag reduction.

fuselage into a side-force and yawing-moment generator. Other investigators have more recently tried similar schemes, but the results shown in Fig. 45 summarize these effects. At a = 35 deg, the conventional rudder was useless because of fuselage blockage and separated flow (see C , = 0 curve), but blowing on the right side of the nose restored directional stability when the vehicle was yawed to the left, and vice versa. Large side force values were also generated by blowing.

C. Microflyer and Pulsed Blowing A combination of all of the above force and moment control applications has been pursued recently at GTRI relative to a very small unmanned aerial vehicle (UAV), the Pneumatic Microflyer.60 Jet turning on a small-scale, lowReynolds-number wing is illustrated in Fig. 46. Pneumatic lift and control surfaces will be driven by gas generated by a GTRI proprietary engine powering the flapping wings of a 6-in.-span flying-insect-like UAV. The opportunity also exists here to take advantage of pulsed blowing, investigated in Refs. 61 and 6 2 for application to blown flaps and in Ref. 6 3 relative to CCW. Here, for properly shaped blowing wave forms, mass flow required was greatly reduced

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

t c

57

0.40

I 0.30

I!

a 0.20

CY SIDE FORCE

a. 10

COEFF.

0.00

-0.10 -O.’O

1 t

I!

O-I5 a 0.10 y1

1

I

0.201

I c CN YAWING MOMENT COEFF.

s I

cn 0.05

2

0.00

-0.0s -0.10

-loo

-5O

SIDESLIP ANGLE,

0

11

+50

+loc

NOSE RICHT-

Fig. 45 Tangential forebody blowing for yaw and side force control, a = 35 deg.

experimentally by up to 40 to 50% (Fig. 47), or conversely, greater lift could be generated by the same mean mass flow levels. Also, as a simplifying means, all pneumatic Microflyer control moments would be generated by differential blowing, rather than by very small moving mechanical parts.

VI. Nonflying Applications of Circulation Control A number of nonflying applications have been investigated, where the CC phenomenon was used to augment or modify flowfields for unique purposes. In order to provide pitch and/or yaw control for submarines without using mechanical stem planes, a dual-slotted “pneumatic” stem plane64 was designed for submerged applications (Fig. 48). Here, up or down pitch of the submarine (or right or left yawing moment) could all be provided by blowing the appropriate slot. Towing basin tests of this concept verified that blowing water from the slots when underwater was equally as effective as pneumatic devices (even if the power required might be higher), and provided the opportunity for smaller stem-, bow-, or sail-planes, or avoided the possible control plane jam problem of moving hydrodynamic surfaces.

58

R. J. ENGLAR

Fig. 46 Low-Reynolds-numbermicroflyer wing with CCW turning.

Applications of pneumatics similar to the CC rotor were both conceived as the CC fan (Fig. 49, from Ref. 65) and the CC windmill. Here, variation in blowing parameters through the individual blade slots could vary the output of the fan, or conversely, for a pneumatic windmill, vary the sensitivity of each individual blade to the incoming wind angle and strength, as well as the radial load distribution on the blades. For the windmill, blade pitch would not be required to change mechanically for maximum performance or avoidance of rotor overspeed. More recently, application of pneumatic concepts to improve the aerodynamic performance of automotive vehicles has been heavily pursued at GTRI. Tests on European Formula 1cars (Fig. 50) have verified that proper application of blowing can dramatically increase the download frequently required for higher-speed cornering of these cars, or reduce the required wing area and its associated drag (note the absence of the conventional inverted fore and aft multi-element wings). The high suction (negative static pressure) difference across a blown lifting wing inverted on a race car can also entrain sufficient flow to provide cooling through a radiator located therein. Figure 51 shows a Formula SAE car with an aft blown wing including a pneumatic radiator (unit developed and tested at GTR166 with assistance from the GT Motorsports team). It is now possible to have a multipoint aerodynamic race car design that had previously been prohibited by the “nonmoving-aerodynamic-components”rule. More details of these concepts and of testing of this device are found in another paper by Gaeta and Englar in this volume and in NASA CP 2005-213509,2005. A GTRI program originally intended only to reduce aerodynamic drag on production cars for increased fuel economy has recently led to additional benefits. Blowing on the curved aft panels67 of a generic streamlined car (Fig. 52) showed drag reductions of up to 35%, but also drag increases of over 100% by blowing different elements, which could be used as a form of aerobraking. Lift could also be increased by up to 170% over the unblown car, or conversely, a lower surface slot could also yield downforce if so desired. GTRI tests also showed that yawing and pitching moments could be dramatically changed by

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

8

m a a

I

59

60

R. J. ENGLAR Optional End Plate

Fig. 48 Blown model stern plane design, two-slotted.

FAN INFLOW

Fig. 49 Circulation control fan concept.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

61

Fig. 50 Pneumatic Formula 1 car model in GTRI tunnel.

blowing, and lateral and directional stability could be restored by blowing only one side of the slot. Interestingly, the blowing required for all of this could be provided by turbochargers or superchargers now being installed on highperformance cars. These experimental data for automobiles have now been extended to and adapted in a GTRI program for the Department of Energy68969to improve the aerodynamics, performance and economics of heavy vehicles (HV; i.e., large tractor/trailer trucks). Figure 53 shows blowing on all four aft comers of the trailer; this combination is able to reduce drag, turbulent separated flow,

Fig. 51 GTRI-patented pneumatic aerodynamic heat exchanger installed on formula SAE race car by GT Motorsports Team.

62

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Fig. 52 GTRI Pneumatic Futurecar model for drag reduction, showing jet turning.

spray, and aft suction on the back doors. Blowing the lower slot only can increase download and aid in braking or provide traction in wet/icy weather, while blowing the top slot only can generate lift and thus reduce effective weight on the tires and rolling resistance. Blowing either side slot can offset yaw due to gusts or sidewinds (which can yield a large component of increased highway drag), or can help to restore lateral/directional stability. Because the response of the blowing system can be virtually instantaneous (pressure of only 1314 psig can produce sonic jet velocity), safety of operation is very promising, including the ability to prevent jackknifing by generating opposite yawing moment for the trailer. Blowing on the trailer top leading edge also appears promising, because it can provide not only a boundary layer control device, but also can entrain flow up through the cab/trailer gap and eliminate strong separation and vorticity there, plus enhance cooling. Wind-tunnel investigations of this concept on a smaller-scale model of a blown pneumatic heavy vehicle (PHV)68,69 have shown drag reductions of up to 80% relative to a baseline generic HV

Fig. 53 Pneumatic heavy vehicle configuration with potential for 5 blowing slots.

OVERVIEW OF CC PNEUMATIC AERODYNAMICS

63

Fig. 54 Pneumatic heavy vehicle test rig undergoing road tests.

model, plus the ability to increase drag if needed for braking, as well as provide side forces and lateral/directional control in side winds. They have also confirmed that blowing only one vertical side slot at the rear of the trailer can eliminate the destabilizing yawing moments due to sidewinds and generate counteryaw in the opposite direction if needed. These tunnel tests have led to development of a full-scale test vehicle and on-road test program of a PHV test rig (Fig. 54), now ongoing for DOE. More data on this pneumatic groundvehicle program are found in another paper by Englar in this volume as well as in NASA CP 2005-213509,2005.

VII. Conclusions Capabilities The high flow-entrainment capability of tangential blowing over curved aerodynamic surfaces has been shown in the preceding discussions to yield augmentation and control of virtually all aerodynamic/hydrodynamic forces and moments by simplified means, which frequently require no moving external components. The capabilities of the CC devices demonstrated include the following: A.

1) Two-dimensional lift coefficients as high as 20 without moving parts and similar high Cl for download as desired in automotive applications. This extra high lift can also provide aircraft Super-STOL capability or the downsizing of wing area for more efficient cruise. 2) Lift augmentations ACl/C, of 80 and very effective boundary layer control.

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R. J. ENGLAR

3) Drag reduction due to flow reattachment and thrust recovery, or drag increase due to flow turning and lift-induced drag, and the ability to pneumatically activate these as needed by the pilot or driver-this is particularly applicable in automotive usage. 4) Aerodynamic moment increases from blowing or differential blowing to provide large control increases compared with those of mechanical devices, or to allow control surface downsizing. 5) Pneumatic engine thrust deflection to 165 deg or more without moving surfaces. 6) Pneumatic propellers or rotor blades to achieve variable thrust and control moment without mechanical cyclic pitch. 7) Automotive applications to vary all forces and moments, including racing vehicle download and drag, without moving parts, using only onboard air sources such as turbochargers. Also, a low-drag aerodynamic heat exchanger using pneumatic-generated pressure difference can cool the vehicle while controlling aerodynamic forces and moments. B. Future of Circulation Control The preceding capabilities offer the potential for aerodynamiclhydrodynamic vehicles simplified by pneumatic multipurpose sui$aces synergistically augmenting lif, drag, moments, control, stability, and propulsive functions without any moving mechanical parts. The force augmentation capability also offers the potential for reduction in wing and control surface areas for improved cruise performance, or multipoint designs with lift/control surfaces sized for optimal points of operation. Future investigations could include improved pulsed blowing to even further reduce the required input mass flows, or to simplify the operation of complex devices such as higher harmonic rotors. Application of CC pneumatics to automotive and hydrodynamic vehicles offers the use of aerodynamic surfaces for functions not currently employed, such as aerodynamic drag reduction or increase, download, heat exchange, thrust augmentation, and stability and control. The opportunity to incorporate all of these devices into a synergistic blown vehicle from the initiation of the design, rather than as an add-on, offers the potential for a very effective and efficient multipurpose vehicle, in which the pneumatic effectiveness, including the propulsion system air supply source and the control systems, is incorporated from the very beginning. A perfect example of how CC could be applied to a new and unique Super-STOL vehicle would be its application to the new NASA ExtremeSTOL concept aircraft, where desired goals include C, of 10, balanced field lengths of 2000 ft or less and, of course, the necessity to trim and control this vehicle at very low speeds, plus the ability to interchange drag increase and drag elimination between approach and takeoff operations, respectively.

References ‘“The Use of Slots for Increasing the Lift of Airplane Wings,” NACA Translation, PW 635, Aug. 1931 (Proceedings L’Aeronautique, June 1931). ’“Wings with Nozzle Shaped Slots,’’NACA Translation, TM 521, July 1929; Berichte Der Aerodynamischen Vereuchsenstalt in Wien, Vol. 1, No. 1, 1928.

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3Metral, A. R., “On the Phenomenon of Fluid Veins and their Application, the Coanda Effect,” AF Translation, F-TS-786-RE, 1939. 4Sproule, R. S., and Robinson, S. T., “Combined Intelligence Objective SubCommittee Report,” WF Document Library Item 5, File No. IX-1, X-2, XII-1, D52.420127, 1944. ’Voedisch, A., Jr., “Analytical Investigation of the Coanda Effect (Project No. FP188),” Air Material Command, Wright Field, Dayton, OH, Rept. F TR-2155-ND, April 1947. 6Cheeseman, I. C., and Reed, A. R., “The Application of Circulation Control by Blowing to Helicopter Rotors,” Journal of Royal Aeronautical Society, Vol. 71, No. 848, 1966. ’Cheeseman, I. C., “Circulation Control and Its Application to Stopped Rotor Aircraft,” AIAA Paper 67-747, Oct. 1967. 8Lockwood, V. E., “Lift Generation on a Circular Cylinder by Tangential Blowing from Surface Slots,” NASA Langley Research Center, Technical Note D-244, May 1960. 9 Englar, R. J., “Circulation Control Pneumatic Aerodynamics: Blown Force and Moment Augmentation and Modification; Past, Present and Future,” AIAA Paper 20002541, presented at AIAA Fluids 2000 Meeting, June 2000. “Englar, R. J., and Williams, R. M., “Test Techniques for High Lift Two-Dimensional Airfoils with Boundary Layer and Circulation Control for Application to Rotary Wing Aircraft,” Canadian Aeronautics and Space Journal, Vol. 19, No. 3, 1973 pp. 93-108. 11 Englar, R. J., “Two-Dimensional Subsonic Wind Tunnel Tests on a Cambered 30Percent-Thick Circulation Control Airfoil,” NSRDC, Technical Note AL-201, AD 91341 lL, May 1972. 12 Englar, R. J., and Applegate, C. A., “Circulation Control-A Bibliography of DTNSRDC Research and Selected Outside References (Jan. 1969 through Dec. 1983),” DTNSRDC-84/052, Sept. 1984. 13Wilkerson,J. B., Barnes, D. R., and Bill, R. A., “The Circulation Control Rotor Flight Demonstrator Test Program,” American Helicopter Society, Paper AHS 79-5 1, May 1979. 14Mayfield, J., “Aeronautical Engineering-Navy Sponsors Coanda Rotor Program,” Aviation Week and Space Technology, 31 March 1980, pp. 69-74. ”Wilkerson, J. B., Reader, K. R., and Linck, D. W., “The Application of Circulation Control Aerodynamics to a Helicopter Rotor Model,” American Helicopter Society, Paper AHS-704, May 1973. 16Nielson,J. N. (ed.), “Proceedings of the Circulation Control Workshop, 1986,” NASA Ames Research Center, NASA CP-2432, Feb. 1986. ”Shrewsbury , G., “Numerical Evaluation of Circulation Control Airfoil Performance Using Navier-Stokes Methods,” AIAA Paper 86-0286, Jan. 1986. 18Novak, C. J., and Cornelius, K. C., “An LDV Investigation of a Circulation Control Airfoil Flowfield,” AIAA Paper 86-0503, Jan. 1986. 19 Englar, R. J., “Experimental Investigation of the High Velocity Coanda Wall Jet Applied to Bluff Trailing Edge Circulation Control Airfoils,” DTNSRDC, Report 4708, Aero Report 1213, AD-A-019-417, Sept. 1975; also M.S. Thesis, Dept. of Aerospace Engineering, Univ. of Maryland, College Park, MD, June 1973. ”Wood, N., and Nielson, J., “Circulation Control Airfoils Past, Present, and Future,” AIAA Paper 85-0204, Jan. 1985. ”Williams, R. N., Leitner, R. T., and Rogers, E. O., “X-Wing: A New Concept in Rotary VTOL,” presented at AHA Symposium on Rotor Technology, Aug. 1976.

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”Reader, K. R., and Wilkerson, J. B., “Circulation Control Applied to a High Speed Helicopter Rotor,” DTNSRDC, Rept. 77-0024, Feb. 1977. 23Rogers, E. O., Schwartz, A. W., and Abramson, J. S., “Applied Aerodynamics of Circulation Control Airfoils and Rotors,” presented at 41 st Annual AHS Forum, May 1985. 240ttensoser, J., “Two-Dimensional Subsonic Evaluations of a 15-Percent Thick Circulation Control Airfoil with Slots at Both Leading and Trailing Edges,” NSRDC, Rept. 4456, July 1974. 25Williams, R. M., and Cheeseman, I. C., “Potential Acoustic Benefits of Circulation Control Rotors,” presented at AHS Meeting on Rotor Acoustics, NASA Langley Research Center, May 1978. 26Englar, R. J., “Investigation into and Application of the High Velocity Circulation Control Wall Jet for High Lift and Drag Generation on STOL Aircraft,” AIAA Paper 74-502, June 1974. 27Englar,R. J., “Subsonic Two-Dimensional Wind Tunnel Investigations of the High Lift Capability of Circulation Control Wing Sections,” DTSNRDC, Rept. ASED-274, April 1975. ”Englar, R. J., “Circulation Control for High Lift and Drag Generation on STOL Aircraft,” A I M Journal of Aircraft, Vol. 12, No. 5, 1975, pp. 457-463. 29Englar,R. J., Trobaugh, L. A., and Hemmerly, R. A., “Development of the Circulation Control Wing to Provide STOL Potential for High Performance Aircraft,” AIAA Paper 77578, June 1977. 30Englar, R. J., “Circulation Control Technology for Powered-Lift STOL Aircraft,” Lockheed Horizons, No. 24, Sept. 1987. 31Englar, R. J., Hemmerly, R. A., Moore, H., Seredinsky, V., Valckenaere, W. G.,and Jackson, J. A., “Design of the Circulation Control Wing STOL Demonstrator Aircraft,” AIAA Paper 79-1842, Aug. 1979; also published in Journal of Aircruft, Vol. 18, No. 1, 1981, pp. 51-58. 32Englar,R. J., “Development of the A-6/Circulation Control Wing Flight Demonstrator Configuration,” DTNSRDC, Rept. ASED-79/01, Jan. 1979. 33Mayfield, J., “Circulation Control Wing Demonstrates Greater Lift,” Aviation Week and Space Technology, March 19, 1979. 34Pugliese,A. J., and Englar, R. J., “Flight Testing the Circulation Control Wing,” AIAA Paper 79-1791, Aug. 1979. 35Loth, J. L., Fanucci, J. D., and Roberts, S. C., “Flight Performance of a Circulation Control STOL Aircraft,” AIAA Paper 74-994, April 1974; also published in Journal of Aircruft, Vol. 13, No. 3, 1976, pp. 169-173. 36Roberts, S. C., “West Virginia University Circulation Control STOL Aircraft Flight Test,” WVU Aerospace, Technical Rept. No. 42, July 1974. 37 Englar, R. J., “Low-Speed Aerodynamic Characteristics of a Small Fixed-TrailingEdge Circulation Control Wing Configuration Fitted to a Supercritical Airfoil,” DTNSRDC, Rept. ASED-81/08, March 1981. 38Englar,R. J., and Huson, G.G.,“Development of Advanced Circulation Control Using High-Lift Airfoils,” AIAA Paper 83-1847 July 1983; also published in Journal ofAircraft, V O ~21, . NO. 7, 1984, pp. 476-483. 39Carr, J. E., “An Aerodynamic Comparison of Blown and Mechanical High Lift Airfoils,” AIAA Paper 84-2199, Aug. 1984. 40Englar, R. J., “Two-Dimensional Transonic Wind Tunnel Tests of Three 15-PercentThick Circulation Control Airfoils,” NSRDC, Technical Note AL-182, AD 882-075, Dec. 1970.

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67

41Englar, R. J., Smith, M. J., Kelley, S. M., and Rover 111, R. C., “Development of Circulation Control Technology for Application to Advanced Subsonic Transport Aircraft,” AIAA Paper 93-0644, Jan. 1993; also published in Journal of Aircraft, Vol. 31, NO. 5, 1994, pp. 1160-1177. 42Englar, R. J., Niebur, C. S., and Gregory, S. D., “Pneumatic Lift and Control Surface Technology Applied to High Speed Civil Transport Configurations,” AIAA Paper 970036, Jan. 1997. 43Mavris, D. N., Kirby, M. R., Lee, J. M., Qui, S., Roth, B., Tai, J., and Englar, R. J., “Systems Analyses of Pneumatic Technology for High Speed Civil Transport Aircraft,” GTRI Final Technical Rept. A-5676, Oct. 1999. 44Ni~hols, J. H., Jr., Englar, R. J., Hams, M. J., and Huson, G.G.,“Experimental Development of an Advanced Circulation Control Wing System for Navy STOL Aircraft,” AIAA Paper 81-0151, Jan. 1981. 45Harris,M. H., Nichols, Jr., J. H., Englar, R. J., and Huson, G.G.,“Development of the Circulation Control WingIUpper Surface Blowing Powered-Lift System for STOL Aircraft,” Proceedings of the ICASIAIAA Aircraft Systems and Technology Conference, Paper ICAS-82-6.5.1, Aug. 1982. 46Yang,H. T., and Nichols, Jr., J. H., “Design Integration of CCW/USB for a Sea-Based Aircraft,” Paper ICAS-82-1.6.1, Aug. 1982. 47Englar, R. J., Nichols, Jr., Harris, J. H., Eppel J. C., and Shovlin, M. D. “Circulation Control Technology Applied to Propulsive High Lift Systems,” Society of Automotive Engineers, Paper 841497, Oct. 1984. 48Lowndes, J. C., “Aeronautical Engineering: Studies Show Lift Coefficient Tripling,” Aviation Week and Space Technology, Dec. 1, 1980. 49Englar,R. J., Nichols, Jr., J. H., Hams, M. J., Eppel, J. C., and Shovlin, M. D., “Development of Pneumatic Thrust-Deflecting Powered-Lift Systems,” AIAA Paper 86-0476, Jan. 1986. ”Eppel, J. C., Shovlin, M. D., Jaynes, D. N., Englar, R. J., and Nichols, Jr., J. H., “Static Investigation of the CCW/USB Concept Applied to the Quiet Short-Haul Research Aircraft,” NASA, TM 84232, July 1982. 51Shovlin, M. D., Englar, R. J., Eppel, J. C., and Nichols, Jr., J. H., “Large-Scale-Static Investigation of Circulation-Control-Wing Concepts Applied to Upper-Surface-Blowing Aircraft,” NASA, Technical Paper 2684, Jan. 1987. 52Bevilaqua,P. M., and Lee, J. D., “Design of Supersonic Coanda Jet Nozzles,” in Proceedings of the Circulation Control Workshops 1986, NASA, CP 2432, pp. 289-312, Feb. 1986. 53Englar, R. J., and Campbell, B. A., “Development of Pneumatic Channel Wing Powered-Lift Advanced Super-STOL Aircraft,” AIAA Paper 2002-2929; presented at AIAA 20th Applied Aerodynamics Conference, June 25, 2002. 54Englar, R. J. and Campbell, B. A., “Experimental Development and Evaluation of Pneumatic Powered-Lift Super-STOL Aircraft,” NASAIONR Circulation Control Workshop, March 2004; also published in NASA CP 2005-213509, 2005. 55Braslow, A. L., “Aerodynamic Evaluation of Circulation Control Propellers,” Bionetics Corp., NASA Contractor Report 165748, June 1981. 56Taback, I., Braslow, A. L., and Butterfield, A. J., “Circulation Control Propellers for General Aviation, Including a BASIC Computer Program,” NASA Contractor Rept. 165968, April 1983. 57Gamer, D., “No Moving Parts, The Circulation Control Airfoil and Fluidic Propeller,” EAA Sport Aviation, Vol. 37, No. 3, 1988, pp. 27-30.

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58Wilson, M. B., and von Kerczek, C., “An Inventory of Some Force Producers for Use in Marine Vehicle Control,” DTNSRDC-79/097, Nov. 1979. 59Wellman, L. K., and Jacobsen, C., “Wind Tunnel Investigation of the Application of Circulation Control to a Forward Swept Wing,” DTNSRDC/ASED-82/05, June 1982. 60 Englar, R. J., “Pneumatic High-Lift and Control Surfaces Applied to Micro-Aerial Vehicles,” Proceedings of GTRI International Conference on Emerging Technologies for Micro Air Vehicles, Feb. 1997. 610yler, T. E., and Palmer, W. E., “Exploratory Investigation of Pulsed Blowing for Boundary Layer Control,” North American Rockwell, Rept. NR72H-12, Jan. 1972. 62Walters, R. E., et al. “Circulation Control by Steady and Pulsed Blowing for a Cambered Elliptic Airfoil,” West Virginia Univ., Dept. of Aerospace Engineering, Rept. TR-32, July 1972. 63Jones,G. S., and Englar, R. J., “Advances in Pneumatic-Controlled High-Lift Systems Through Pulsed Blowing,” AIAA Paper 2003-3411; presented at AIAA 21st Applied Aerodynamics Conference, June 2003. 64Englar, R. J., and Williams, R. M., “Design of a Circulation Control Stem Plane for Submarine Applications,” NSRDC, Technical Note AL-200, March 1971. 65F~ry, R. J., and Whitehead, R. E., “Static Evaluation of a Circulation Control Centrifugal Fan,” DTNSRDC, Rept. 77-0051, AD A041-463, June 1977. %aeta, R. J., and Englar, R. J., “Pneumatically Augmented Aerodynamic Heat Exchanger,” Paper presented at NASA/ONR Circulation Control Workshop, March 2004; also published in NASA CP 2005-213509, 2005. 67Englar, R. J., Smith, M. J., Niebur, C. S., and Gregory, S. D., “Development of Pneumatic Aerodynamic Concepts for Control of Lift, Drag, Moments and Lateral/Directional Stability of Automotive Vehicles,” Society of Automotive Engineers, Paper 960673, Feb. 1996; also published in SAE SP-1145, “Vehicle Aerodynamics,” Feb. 1996. 68EnglarR. J., “Drag Reduction, Safety Enhancement and Performance Improvement for Heavy Vehicles and SUVs Using Advanced Pneumatic Aerodynamic Technology,” 2003 SAE International Truck and Bus Meeting and Exhibition, Society of Automotive Engineers, Paper 2003-01-3378, Nov. 2003. 69 Englar, R. J., “The Application of Pneumatic Aerodynamic Technology to Improve Performance and Control of Advanced Automotive Vehicles,” NASA/ONR Circulation Control Workshop, March 2004; also published in NASA CP 2005-213509, 2005.

Chapter 3

Exploratory Investigations of Circulation Control Technology: Overview for Period 1987-2003 at NSWCCD Robin Imber* Naval Air Systems Command, Patuxent River, Maryland and Ernest Rogerst and Jane Abramsont Naval Sur$ace War$are Center-Carderock Division, West Bethesda, Maryland

Nomenclature A = area of CC slot, or area of foil planform AR = aspect ratio C, = momentum coefficient of slot flow (rizvj/qS) CL = lift coefficient ( L / q S ) C, = drag coefficient ( D / q S ) C, = power coefficient CT = thrust coefficient c = chord length D = drag force d = diameter, or camber line offset h = CC slot exit height (gap) L = lift force rit = mass flow (pAV) PR = pressure ratio q = dynamic pressure (&pV2) S = planform area of lifting surface t = thickness of airfoil

*Aerospace Engineer. 'Aerospace Engineer, retired. This material is declared a work of the U.S.Government and is not subject to copyright protection in the United States.

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70

V = velocity a = angle of attack ACL/AC, = lift augmentation ratio (slope of CL vs C, curve) p = density u = rotor solidity, cavitation index

Subscripts

i = induced j =jet I. Introduction EGINNING in 1967, when the Naval Surface Warfare Center, Carderock Division (NSWCCD) was known as the David Taylor Model Basin (DTMB), researchers there were involved with many circulation control (CC) exploratory projects, including CC-airfoils, CC-centrifugal fans, dual-directional CC-airfoils, CC-fixed wings, including the A6 aircraft modification, CC-rotorcraft, including XH2-CCR and X-wing, CC-hydrodynamic applications, and valving systems for CC. The first Circulation Control Workshop (unpublished) was held at DTMB in 1971, and the second was held at the National Aeronautics and Space Administration, Ames Research Center, in 1986.' Papers, presentations, and reports of the research performed from 1967 to 1985 at what is now NSWCCD (there have been several name changes since 1967) are cited in the proceedings from the 1986 CC Workshop and more can be found in Ref. 2. This overview is intended as a brief summary of the highlights of six of the major CC experimental investigations that have taken place at NSWCCD since the second CC Workshop, specifically between 1987 and 2003, and was presented at the 2004 Circulation Control Workshop held in Hampton, Virginia.3 The following investigations are discussed: 1) The Dual-Slotted Cambered Airfoil, LSB; 2) The Self-Driven Rotary Thruster, TIPJET; 3) The Annular Wing, CC-Duct; 4) The Circular Wing, CC-Disc; 5 ) The Miniature Oscillatory Valve, CC-Valve; and 6) The Dual-Slotted Low Aspect Ratio Wing, CCHydrofoil. For further details regarding these investigations, the reader is encouraged to examine the publications that are referenced for each of the projects. Used throughout this summary is a frequently used measure of CC performance, the lift augmentation ratio. This ratio is defined as the ratio of the gain in lift (ACL)to the change in slot flow momentum (AC,). In this review, the ratio is determined from experimental data by assessing the slope of the lift response in the low blowing (low C ), range, where the response is usually linear.

B

11. Dual-Slotted Cambered Airfoil (LSB) The Dual-Slotted Cambered Airfoil, also referred to as the LSB (lower surface blowing), was designed and tested in 1987 by Abramson and colleague^.^,^ The inclusion of a lower surface slot along with the usual upper surface slot provides the ability to produce lift in both positive and negative directions. The presence of camber in this model, along with the objective of preserving the contour of a

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71

Fig. 1 Photograph of the dual-slotted cambered airfoil (LSB).

proven (parent) single-slot CC airfoil, means that the geometric properties of the lower-slot region are not the same as those of the upper-slot region. Additionally, when operating in the lower surface blowing mode, the CC section functions with negative camber, all with unexplored consequences at that time. A photograph of the LSB model is shown in Fig. 1 and a cross-section drawing of the model is shown in Fig. 2. The LSB has a 17% thickness ratio with 1.1% camber. It was constructed with a 12-in. chord and a 36-in. span. The upper slot is located at 96.8% chord and the lower slot is slightly further aft at 97.0% chord. The airfoil was experimentally evaluated in the NSWCCD 8 x 10 ft wind tunnel configured with two-dimensional wall inserts. Testing included three blowing modes: upper surface only, lower surface only, and dual blowing. The wind tunnel dynamic pressure ranged from 20 to 60psf, Reynolds number ranged from 0.8 to 1.4 x lo6, and angle of attack (AOA) ranged from - 10 to +10 deg. Two slot height-to-chord (h/c) ratios were set: 0.0013 and 0.0020. The maximum momentum coefficient (C), was 0.22. One of the main design goals was to have the dual-slotted model perform as well, when using only the upper slot, as the single slotted “parent” model. The dual-slotted model had the same cross-section as the parent model. Lift performance results from the single and dual slot models are shown in Fig. 3. The comparison shows that there was no detrimental effect in adding the second slot. The second design objective was to increase the control range so that force control in both directions was available. Figure 4 displays a plot of lift coefficient Upper Surface Blowing Slot Coanda Surface

Air Supply Ducts

Fig. 2 Cross-section of LSB airfoil!

Lower Surface Blowing (LSB) Slot

R. IMBER, E. ROGERS, AND J. ABRAMSON

72

ACr due to blowing

Momentum Coefficient, Cp

Fig. 3 Comparison of lift performance for dual-slotted LSB and single-slot parent airfoil?

against momentum coefficient, and reveals that the goal of doubling the control range was met. An unanticipated finding was that the performance of the lower slot, in terms of measured lift augmentation ratio, was noticeably better than that for the upper slot (80 compared with 60). This empirically unexpected performance enhancement illustrates the need for a well-validated computational code (computational fluid dynamics; CFD) to help guide future CC designs. Another finding was the effect of simultaneous blowing5 At the only ratio of dual blowing examined,

Lift Coefficient

Momentum Coefficient

Fig. 4 Control range increase demonstrated with upper and lower slot ~apability.~

INVESTIGATIONS OF CC TECHNOLOGY AT NSWCCD

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where the lower slot flow momentum level was 25% that of the upper slot, activation of the lower slot decreased lift. This dual-slotted airfoil helped pave the way for another dual-slotted CC investigation discussed later, the CC-Hydrofoil. Summarizing the key findings from the Dual-Slotted Cambered Airfoil investigation, it was found that 1) incorporation of a lower slot did not affect performance of the upper slot; 2) the available lift control range was doubled, as expected; 3) a lift augmentation ratio of 80 for the lower slot was obtained; and 4) simultaneous blowing can be used to decrease (control) the lift increment produced by single-slot operation.

111. Self-Driven Rotary Thruster (TIPJET) Experimentally investigated in 1991 by several NSWCCD engineers, the SelfDriven Rotary Thruster was the first integrated lift/reaction-drive rotor system combining Cbanda CC aerodynamics with cold cycle reaction drive technologies.6-8 The rotor was developed as part of the TIPJET unmanned air vehicle, shown conceptually in Fig. 5 . The design involves a stoppable two-bladed rotor concept where, after lifting off vertically in rotary mode and accelerating forward, the rotor transitions to a fixed wing to enable high-speed flight. A “cold cycle” gas generator, such as the fan stage of a turbofan engine, supplies the compressed air for both the circulation control and the tip jets that provide rotor drive torque. Figure 6 shows a sketch of the completely pneumatic rotor. Circulation control slots are located along most of the rotor blade span on both the leading and trailing edges. Reaction drive nozzles are located at the rotor tips. Thus, a single source of air pressure provides flow for the CC slots to augment rotor lift (vertical Y

Fig. 5 TIPJET vertical takeoff and landing unmanned air vehicle with circulation controlled stoppable rotor.6

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R. IMBER, E. ROGERS, AND J. ABRAMSON

No drive shaft, unarticulated, flat-pitch blades

NOZZLE DRIVE FORCE

AIR SUPPLY

Fig. 6 Sketch of TIPJETcompletely pneumatic rotor?

thrust), while at the same time providing the rotor torque drive via the tip-jet nozzles. (The full-scale application concept called for in-flight controllable slot gap settings.) A detailed investigation of the TIPJET rotor in hover took place in 1991.7 The primary objective of the hover experiment was to evaluate the interactions between the lift and drive systems. Drawings of the aluminum rotor model are shown in Fig. 7 and specifications of the model are listed in Table 1. The 80in. rotor blade is tapered, with no twist and zero pitch angle. The thickness and camber varies linearly with radius from the 25% to 95% span locations. Figure 8 is a photograph of the blade tip region, showing the CC slot along the span and the tip-drive nozzle. During the hover test, the rotor could be driven by either an electric drive motor that enabled the rotor to be operated at selected rpm settings while investigating specific performance attributes, or by the tip-jet reaction drive. To better understand the performance of the integrated lift/drive system, a detailed investigation was conducted with the tip nozzles closed and the rotor mechanically driven. Figure 9 shows the experimental data for the measured rotor thrust coefficient as a function of momentum coefficient for several slot height settings. The slope of the curve is 29 at the lower values of CJu. This ratio is higher than that of any previously tested CC rotor. As shown in Fig. 9, this measure of efficiency was independent of the four slot heights tested. To determine if the level of understanding of the performance of the fully pneumatic rotor system was sufficient for successful incorporation into a flight vehicle, numerical calculations were developed and compared to the experimental performance. The results of this comparison, shown in Fig. 10, indicate excellent correlation for both the rotor thrust developed and drive power required. The ultimate goal of the experimental rotor investigation was to determine the aeromechanics of the model rotor in self-drive mode. The behavior of the rotational speed in response to pressure input was unknown at the beginning of

INVESTIGATIONS OF CC TECHNOLOGY AT NSWCCD

SLOT HEIGHT ADJUSTMENT SIDE-BY-SIDEOPPOSING SCREWS

Fig. 7 Drawings with details of TIPJET pneumatic rotor?

75

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R. IMBER, E. ROGERS, AND J. ABRAMSON

Table 1 TIPJET rotor specifications’ Blade Fixed pitch angle (OJ, deg Rotor diameter, ft Number of blades Chord, in. 25% span 93% span Solidity ratio Geometric twist, deg Airfoils Thickness ratio ( t / c ) Camber ratio (d/c) Trailing edge radius (rt,/c) Slot height ( h / c ) Tipjet nozzles (rectangular) Area/nozzle, in.*

0 6.67 2 7.95 5.40 0.110 0 25% span 0.213 0.053 0.05 variable

93% span 0.170 0.011 0.03 variable

0.764

the test. When a slot height and tip nozzle area are set, the blade pressure input is the only determining factor of the operating condition. Shown schematically in Fig. 11, as pressurized air is introduced into the rotor, the rotor rotates in reaction to the flow from the tip nozzles. As the nozzle flow increases, the rotor lift will increase as a result of using a cambered airfoil and, most especially, as a result of the increased circulation due to the CC slots. This additional lift increases the required drive torque, which then limits the rotational rate for a given air pressure setting. In essence, pressure input simultaneously influences the lift and produces the torque drive. Identifying the behavior of a system coupled by these two effects was one of the objectives of the experiment.

Fig. 8 Tip region of the fully pneumatic rotor model.

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77

c

f SLOT MOMENTUM COEFF. / SOLIDITY, C ~ / O Fig. 9 TIPJET rotor thrust performance when mechanically driven at constant rpm.’

It was discovered that, in full self-drive mode, the rotational speed is stable and exhibits a self-limiting maximum for a given ratio of slot area to nozzle area. The data in Fig. 12 reveal the nature of the self-limiting rotational speed. Rotational tip speed is shown as a function of blade pressure for several slot height settings. At each of the slot settings, the pressure input response of the rotor is to increase the rotational rate until a limiting tip speed is reached. It was demonstrated that the value of the limiting tip speed is a function of the slot height setting; increasing the slot height results in a lower limiting tip speed. However, because CC-based lift is not solely dependent on local velocity, the lift response is not limited to a tip speed maximum. Figure 13 shows that the rotor lift is essentially always the same linear function of the applied pressure, independent of actual rotational rate (compare Figs. 12 and 13). A major finding was that a non-shaft-driven completely pneumatic rotor inherently seeks a rotational rate that results in lift being a near-linear function of the blade pressure input, and this linear lift is easily controllable by pressure throttling. In addition to the specific TIPJET vehicle application, this capability can be applied to systems that require mechanically simple, easily controlled thrusters. Summarizing some of the key findings from the Self-Driven Rotary Thruster investigation, it was found that 1) a lift augmentation ratio of 29 was obtained when in the mechanically driven mode; 2) the pneumatic rotor inherently seeks equilibrium and the self-limiting rotational rate is a function of slot-to-drivenozzle area ratio (the resulting rotor lift is a near linear function of the blade pressure); 3) there is a significant impact on induced power efficiency because of the non-lifting tip nozzle region; and 4) the presence of the tip nozzle jet has no discernible impact on the external aerodynamics of the lift system.

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. D

c

Fig. 10 TIPJET performance numerical analysis correlation with experimental results.'

cc slot area

Slot flow

1

t

r3

I

Internal air

Rotor

Drive

rotor drive torque

Fig. 11 Conceptual schematic of TIPJET rotational rate equilibrium.

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d a, a,

Q

v)

Blade Root Pressure Ratio, PR ,oo,

Fig. 12 TIF'JET rotor characteristics when self-driven via tip-jet nozzles?

IV. Annular Wing (CC-Duct) In the mid-1990s the performance characteristics of an annular wing, having both inner and outer trailing edge circulation control slots, were explored. The focus of this investigation was to apply full, or partial, perimeter trailing

Blade Root Pressure Ratio, PR,,,,

Fig. 13 Relationship of blade pressure to thrust developed.' (Data and symbols are the same as in Fig. 12.)

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R. IMBER, E. ROGERS, AND J. ABRAMSON

Fig. 14 Annular wing model installed in free jet wind tunnel?

edge CC fluid ejection to a propulsor duct to enhance maneuvering control on watercraft via a thrust vectoring capability. An existing CC-Duct model with inner and outer trailing edge CC slots had been borrowed from West Virginia University. The model has been used in the 1970s to investigate the attributes of variable diffusion for ducted fans on aircraft.’ The CC-Duct model is shown in Fig. 14 as it was installed in the Atlantic Applied Research Corporation open-jet acoustic tunnel in 1993. The model originally had a motor housing and stator that were removed for the CC-Duct investigation discussed here. The objective for removing the motor housing and stator was to have a simple configuration in which to establish an understanding of the performance attributes, and to provide a data set for correlation to basic ring-wing theory. No propeller was present in any of the test series. A cross-section of the top of the CC-Duct is shown in Fig. 15 and its geometry is presented in Table 2. The model is 18411. in diameter with a 10-in. chord and a 20% thick uncambered foil section. The inner and outer slots are located around the full trailing edge circumference at 97% chord. During the Duct investigation, metal foil tape was applied and burnished well to the external surface of the model, over the slot region, in order to temporarily block off portions of the slot. This commonly used technique for exploratory CC research provides the ability to control the distribution of slot flow. By using the tape, several configurations, focusing on potential attributes of the Duct, were tested.

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Dimensions in inches

Fig. 15 Annular wing model cross-section (at top).’

The model and experimental arrangement provided the opportunity to examine many interesting flow control configurations. Icons representing six basic configurations, or modes of operation, and a brief explanation of each, are displayed in Table 3. The dotted line on each of the icons represents the trailing edge of the CC-Duct. The solid lines represent the portion of either the inner or outer slot that is open and where the fluid ejection occurs. The author suggests reviewing Table 3 before reading further. As a simple illustration of force vectoring capability, a long strand of yam was positioned in the center of the CC-Duct. The photograph on the left of Fig. 16 shows the configuration with no CC blowing. The yam is aligned with the tunnel free stream velocity, along the longitudinal axis of the Duct. The photograph on the right of Fig. 16 was taken for a complimentary halves configuration, where the outer lower slot is active and the upper inner slot is active. The yam strand is now at an angle to the free stream, indicating the wake deflection due to the force vectoring brought about by the active flow control. Quantitative data from the configuration on the right in Fig. 16 are shown in Fig. 17 as a plot of force developed as a function of C ., The reference area used

Table 2 Annular wing model specifications Model geometry Outside diameter Inside diameter Chord Slot gap hlc Slot position d/c (16.2110) AR (effective)

Dimensions, in. 18.2 14.2 10

0.009 0.0009 0.970 c 1.62 2.1

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82

Table 3 CC annular wing modes of operation as a marine propulsor duct Aft view of ducta

Slot ejection configuration

Effect

Operation benefit

Inner slot only

Increased duct flow-through: accelerating nozzle

Higher prop efficiency

Outer slot only

Decreased duct flow: diffusion

Reduced cavitation

Complementary quadrants

Side-force: yaw

Steerage

Complementary quadrants

Side-force: pitch

Depth keeping

Alternating

Vortex generation: very high drag, no side force

Braking: crash-back

Both slots

Drag reduction, auxiliary thruster

Cruise efficiency, dock side positioning

aDashed line is trailing edge; solid lines represent active slot.

for the force coefficients is duct length times duct diameter (c x d). The lift force developed for this configuration, even at zero model pitch angle, is more than twice that available on a passive duct. When interpreting the CC performance of ring wings, in comparison to that of flat wings or airfoils, it is important to be observant of how the performance parameters are nondimensionalized. In Fig. 17, note that the CL versus C, performance curve for the duct is close to that of some CC airfoils (e.g., see Fig. 4). This result occurs even though the finite wing effect of shed wake vorticity causes a downwash that should decrease the net lift response to roughly half that of a two-dimensional CC foil. The explanation for the apparent two-dimensional-like performance level has to do with the reference areas used for the coefficients. Consistent with the practice for CC airfoils and wings, the duct C, is based on the full slot length (Tx d x c). At the same time, as is standard practice, CL is defined based on projected area (c x d ) , without consideration that there are two lifting surface areas involved: the upper and lower halves of the ring-wing. Therefore, the CL that should be compared to that of an airfoil is half the CL shown in Fig. 17, and thus

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No Blowing

83

180 I 180 deg blowing

Fig. 16 Lateral force capability; wake deflection with asymmetric trailing edge CC blowing?

matches what would be expected from a wing having an as ect ratio (AR) of 2.1, which is the equivalent AR for the geometry of the duct."(See the later discussion and performance of the CC-Hydrofoil "flat" wing, which has about the same aspect ratio.)

Slot Flow Momentum Level CF = mV,/( % pVm2ndc)

Fig. 17 CC annular wing lift and drag performance' (the drag force is in the direction of a negative propulsive force).

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Test results met expectations including that the drag (plotted as negative thrust) is a linear function of C , as shown in the experimental data, and derived as in the following equations. From the experimental data,

CL = 1 o G From Hoerner and Borst"

AR = 2.1

(2)

and for lift-induced drag

then

One of the findings from the CC-Duct investigation was the ability for braking, or control of induced drag, without a change in net lift. Alternating the active inside and outside slots every 90deg creates two pairs of counterrotating vortices. Figure 18 shows the measured performance demonstrating this capability. The drag is about the same as it was when lift was being developed. Figure 19 represents a VSAERO potential-flow solution of surface pressure and wake filaments for a CC-Duct braking configuration. (A discussion of

Slot Flow Momentum Coefficient (Cp) Fig. 18 Braking configuration measured data; lift-induced drag without net lift?

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Wake Filaments looking upstream

Fig. 19 Braking configuration computational solution."

potential flow solution techniques related to this work, and CC applications in general, can be found in Ref. 11.) The key findings from the CC-Duct investigations include 1) lift and side force can be generated using specific blowing segments; 2) at zero angle of attack (AOA), forces of almost 2.5 times a conventional ring-wing are possible; 3) a braking force via induced drag is available without the development of lift; and 4) the performance met expectations and it was found that the performance can be predicted using a potential flow code, although the slot flow requirements have to be estimated from the historical CC airfoil database.

V. Circular Wing (CC-Disc) In 1995, Rogers and Imber created a circular wing model with a full perimeter circulation control capability. l 2 The purpose was to investigate the effectiveness of CC on very low aspect ratio wings and to explore the attributes of a CCenhanced omnidirectional type of control surface or vehicle. The CC-Disc, also known as the Coanda Disc, was tested in the 8 x loft Subsonic Wind Tunnel at NSWCCD, employing a six-component external balance for force and moment measurement. Figure 20 is a photograph of the anodized aluminum 2-ft-diam. model and Table 4 lists the model specifications. The Disc has a 19% thick cross-section with 2.4% camber. Figure 21 shows a drawing of the

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Fig. 20 Circular wing model with full perimeter circulation control.

centerline cross-section with surface pressure tap locations. The baseline slot height was 0.032 in. The major configurations investigated are shown in Fig. 22. The circular icons represent a planform view of the Disc with the shaded sections representing the perimeter sections where fluid ejection occurred. In Fig. 22, the free-stream flow would be directed from the top to the bottom of the page. The three groups of configurations were 1) increasing area centered about the trailing edge; 2) constant area at variable azimuth; and 3) lateral and asymmetric variations. The same slot tape-over technique used in the CC-Duct test was employed for flow control configuration changes on the CC-Disc. Results from the first configuration group are shown in Fig. 23 for the model at zero pitch angle, gradually increasing the flow ejection circumference region. Lift is shown as a function of the region of blowing, starting with unblown and then, centered about the trailing edge, increasing the perimeter region blown until there was full 360-deg fluid ejection. The lines on the plot are for constant

Table 4 Specifications of circular wing Diameter (chord) Reference area (S) Aspect ratio (AR) Thickness ( t l c ) Camber Coanda radius Slot position Slot lip thickness Slot height ( h ) hlc hlrs Pressure tap diameter

2 ft 3.14 ft2 1.27 19% 2.4% rs/c = 0.050 rte/c = 0.040 3.2%c from edge 0.026 in. 0.032 in. 0.0013 0.027 0.040 in.

\

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Surface Pressure Taps

Centerline Cross-section

Fig. 21 Circular wing model with full perimeter circulation control.'*

blowing coefficient C,. The optimum, or highest lift, configuration varied somewhat with the C, level. The overall highest lift was obtained using a 225-deg perimeter of fluid ejection centered about the trailing edge. It is notable that high lift performance was obtained even with full perimeter blowing, showing that the omnidirectional configuration is viable. Figure 24 reveals lift performance results for the 225-deg perimeter configuration at several AOA and C, levels. The maximum C, was more than twice that available from a disc without circulation control. Both with and without blowing, the slope of C,/a closely matches the slope calculated from inviscid theory. The same configuration and data collection series is shown in Fig. 25 plotted as lift

lncreasina Area Centered About the Trailina edae:

Fig. 22 Circular wing test configurations: planform view (sections with blowing shown in black).

aa

R. IMBER, E. ROGERS, AND J. ABRAMSON

Configuration

Fig. 23 Lift as a function of azimuthal mass ejection coverage.'*

coefficient squared against drag coefficient. The measured induced drag matches the prediction of lifting surface wing theory. These matches suggest that, for low aspect ratio wings, there are no basic effects unique to lift development by means of the Coanda form of CC.

Angle-of-Attack (deg) Fig. 24 Influence of angle of attack?

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89

CD Fig. 25 Aerodynamic efficiency: induced drag with incidence?

As discussed in more detail in Ref. 12, the circular wing with CC has two aerodynamic centers: 1) lift due to angle of attack and 2 ) lift due to circulation control. The pitching moment map in Fig. 26 demonstrates the results of these two centers. The pitching moment was established about the center of

moment

Fig. 26 Pitching moment map resulting from two aerodynamic centers?

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R. IMBER, E. ROGERS, AND J. ABRAMSON

CP Fig. 27 Sensitivity to slot gap setting?

the disc and the plot can be used to determine the combination of pitch angle and level of circulation control that will provide a specific desired maneuvering effect. During part of the CC-Disc investigation, the slot height around the perimeter was changed to determine the sensitivity of wing performance to slot gap setting. For the range of C, shown in Fig. 27, there was essentially no change in lift performance for a 4:l slot gap change. The lack of performance change with the large change in slot gap is considered a desirable attribute, because it allows flexibility in the selection of a slot-flow supply system. For many of the configurations investigated, upper and lower surface pressures were measured every 10 deg in azimuth. A sample of the upper surface pressure distributions of four configurations is shown in Fig. 28. The area of fluid ejection is easily seen as the darker shaded regions around the perimeter, which have significant low pressure values and steep pressure gradients. The radial lines that appear in Fig. 28 are a product of the plotting software and not a true characteristic of the overall pressure distribution. There were many key findings from the Circular Wing test. The investigation demonstrated that 1) circulation control is effective on very low aspect ratio lifting surfaces and, for a circular planform, can provide an omnidirectional capability when full perimeter blowing is applied; 2) with at least 225 deg of flow control around the perimeter the lift produced was more than double that of an unblown circular wing (the limit to augmented lift is believed to be the result of excessive wall jet turning); 3) roll control was demonstrated using asymmetric blowing; 4) lift control without change in pitching moment was demonstrated when blowing only the lateral edges; and 5 ) sensitivity to the 4:l change in slot gap is minimal.

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91

High

pressure

Low pressure Full perimeter blowing

Fig. 28 Upper surface pressure results showing four configurations?

VI. Miniature Oscillatory Valve (CC-Valve) for Unsteady Wing Load Reduction There were two main objectives for this project. The first objective was to demonstrate that mass ejection in the trailing edge region of a hydrofoil could be used to cancel periodic unsteady hydrodynamic loading. The second objective was to show that a practical closed-loop control system could be devised and that the required oscillatory valving could be miniaturized and incorporated into the trailing edge region of a hydrofoil. The focus of the miniature valve design and control demonstration was to develop the capability to cancel unsteady foil forces and be automatically adaptive to upstream disturbances. Fry and Jessup designed and tested the slot control valve in 1993.13-15 An overall sketch of the 15-in. chord, dual-slotted hydrofoil used in this demonstration is shown in Fig. 29. The flow ejection for this application was normal to the surface, as in a jet flap. (The concept could be adapted to the production of tangential mass ejection). Figure 30 shows a trailing edge section-cut of the

dual-slotted, surface-normal mass ejection into the boundary layer

Fig. 29 Dual-slotted hydrofoil used for miniature valve proof of ~ 0 n c e p t . l ~

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R. IMBER, E. ROGERS, AND J. ABRAMSON

electromagnetic actuator

14

Fig. 30 Trailing edge of dual-slotted hydrofoil: rotor pivots to open/close upper lower slots. Constant fluid pressure results in high-response-rate,efficient system.

model and Fig. 3 1 depicts the actuator mechanism construction. A small rocker valve embedded in the trailing edge uses an electromagnetic actuator attached to a permanent magnet assembly to produce a high-frequency response of the rocker valve. As shown in Fig. 30, there is an optional nonmovable trailing edge tail section. The main feature of the actuator was that it controlled the slot exit area and not the fluid pressure. A schematic of the water tunnel installation is shown in Fig. 32. The hydrofoil was attached to one side of a 24-in. water tunnel test section. A freewheeling propeller provided a periodic upstream flow disturbance. An external pump and fluid lines delivered fluid to the rocker valve assembly region. A controller was employed to send signals to the magnet assembly installed in the trailing edge body of the hydrofoil.

Permanent Magnet Assembly

Actuator

Coil Assembly

Rotor Valve Assembly

Fig. 31 Actuator mechanism constr~ction.'~

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24-inch water tunnel

Fig. 32 Operational schematic of actuator test.

During operation, fluid pressure to the trailing edge of the foil is not throttled; it is simply redirected as needed. This low-rate perpendicular ejection of mass into a boundary layer, from a slot at the trailing edge, may mean that the flow effect responsible for any change in foil lift is the same as the flow effect attributed to a Gurney flap. (The term “low-rate’’ ejection is used to make a distinction from the high momentum level of a true jet flap.) As an example of the experimental results, two plots of force against frequency are shown in Fig. 33. The top plot shows the periodic force spectrum produced by the hydrofoil due to the upstream flow disturbance. The bottom plot shows the force variation with the active flow control system operating. As shown, the targeted hydrofoil load spikes were successfully eliminated by the system, with three frequencies simultaneously reduced. The key findings from the Miniature Oscillatory Valve project include the following: 1) functional model-scale actuators can follow steady or time-varying input signals up to 500 Hz;2) hydrofoil forces were successfully varied up to 110 Hz;and 3) alleviation of a high-frequency periodic hydrofoil loading is feasible.

VII. Dual-Slotted Low Aspect Ratio Wing (CC Hydrofoil) In 2002, Rogers16 was the principal investigator for an in-depth low aspect ratio hydrofoil investigation that employed dual-slotted trailing edge CC. The experimental investigation took place primarily in the Navy’s 10-ft-Large Cavitation Channel (LCC) and is extensively documented in Ref. 16. The question as to the effect of cavitation on the performance of a CC-foil had been pondered for many years and was finally answered during this research.

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R. IMBER, E. ROGERS, AND J. ABRAMSON

Frequency (Hz)

Fig. 33 Load cancellation effectiveness. Two plots show three frequencies of unsteady lift simultaneously reduced to the broadband fl00r.l~

There are several compelling reasons to incorporate circulation control fluid dynamics on underwater vehicles. The buoyancy of waterborne vehicles means that they can and do operate down to extremely low speeds, where conventional control surfaces have very limited force generation capability. Because the forte of CC is to leverage the momentum flux from a slot to make a planar surface produce much greater force than otherwise possible, it becomes an attractive consideration for low-speed maneuvering enhancement. Furthermore, CC augmentation has its best pumping-power efficiency at low speed, in terms of the control force advantage over a conventional surface. Photographs of the model installed in the water tunnel on a reflection plane, and of the dual-slotted cross-section, are shown in Fig. 34. A more inclusive view of the test section is provided in Fig. 35. The half-span model of aspect ratio 2.0 has an uncambered 20% thick elliptical profile essentially identical to a previously tested CC airfoil, thus allowing a comparison of two- and threedimensional (finite wing) performance. The slot-height-to-chord ratio of approximately 0.0018 was maintained over the full 2-ft span of the tapered planform. The fluid pressure in the dual plenums could be individually regulated and model loads were measured by a multicomponent balance. Fundamental lift performance is shown in Fig. 36 with lift coefficient as a function of C, for the model at 10 deg AOA. At low C, levels, the wing performed slightly better than expected, producing a lift augmentation ratio of 36 in the initial linear portion of the curve. Transition from a linear to a squareroot-like response to C, occurred, as expected, at the higher blowing levels. An important discovery was made early in the test. With only the upper slot blowing, lift roll-off occurred at a much lower C, than expected, as shown on

INVESTIGATIONS OF CC TECHNOLOGY AT NSWCCD

Fig. 34 Dual-slotted low aspect ratio circulation control hydrofoil.'6

Fig. 35 Dual-slotted hydrofoil installed in the Large Cavitation Channel.

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R. IMBER, E. ROGERS, AND J. ABRAMSON

96 3.5

AOA = 10°

3

with 2nd

slot flow

2.5 2 CL 1.5

single slot (max conventional)

1 0.5

with 5% lower slot Cµ no lower slot flow

R52

0 0

0.1

0.2

0.3

0.4

0.5

0.6

Cµ Cp total (upper + + lower)

Fig. 36 Lift performance and benefit of dual-slot activation.16

the curve in Fig. 36 labeled “single slot”. It was concluded that excessive turning of the wall jet was causing the loss in lift. The lower slot was then employed to produce a very small counter flow, no larger than 5% of the upper flow, to see if it would prevent the excessive turning. This dual-flow configuration produced the greatly improved performance shown in the upper line in Fig. 36. Note that there was no performance penalty at low C, for the dual-flow (where the benefit was not needed) and the investigated C, range extended to 0.5, a very high value for CC. (Recall that the prior dual flow experiment on the LSB airfoil had used a much higher percentage of second slot flow, 25%, with an accompanying decline in lift.) The comparison of actual to expected performance for the hydrofoil, shown in Fig. 37, shows excellent agreement. For the three-dimensional foil the average value of C, vs C, is about 50% of that seen in the corresponding twodimensional airfoil data. This is the same ratio of three- to two-dimensional performance as found for the C, vs AOA on conventional wings of the same aspect ratio as compared to an airfoil. Also, similar to the circular wing discussed earlier, the induced drag performance of the hydrofoil matched predictions based on conventional lifting line theory. One of the major test objectives was to determine where the minimum pressure occurs on the model and what the impact of subsequent cavitation would be on the ability of the jet to induce circulatory lift, or even to remain attached. Cavitation occurs when the minimum pressure reaches the value corresponding to the vaporization of water, about 0.5 psia depending on temperature. is the term for the absolute value of the pressure The cavitation index, sigma (a), coefficient that will result in vaporization and is a function of the test section static and dynamic pressures.

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3.5 AOA = 0o w/ 2nd -slot

3

airfoil data

2.5

CL

2

CD

1.5

wing data

prediction

prediction

CD

1

drag pred.

0.5 0

G3

0

0.05

0.1

0.15

0.2

0.25

0.3

Slot Momentum Momentum Coefficient Coefficient (total), (total), Cµ Cy

Fig. 37 Comparison of actual and expected performance.16

The data plotted in Fig. 38 show that after the onset of cavitation, the lift continued to increase in response to increasing duct pressure. Eventually the lift began to roll over, but not abruptly. At no time did the Coanda jet detach prematurely from the trailing edge due to cavitation. For this particular model 2 AOA = 0° single slot

σ = 13.5

1.5

10.2 6.6

CL

1 onset of cavitation on Coanda surface for σ = 6.6

0.5

sigma = 13.5 sigma = 10.2 sigma = 6.6

σ = Cpmin at which water would vaporize 0 0

0.05

Cµ CP

0.1

0.15

Fig. 38 Lift response to Coanda surface cavitation developrnent.I6

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R. IMBER, E. ROGERS, AND J. ABRAMSON

Fig. 39 Cavitation induced by decreasing tunnel static pressure at moderate lift coefficient.I6

design, cavitation initiated on the nozzle lip face; see the sketch in Fig. 39. The photograph in Fig. 39 shows some interesting flow visualization on the CCHydrofoil, compliments of the cavitation that caused the white bubbles of vaporized water. Cavitation is not likely to occur operationally, but if it does, it is not catastrophic to the fundamental CC effect. Another advantage of the dual slots is the ability to vector the jet thrust. In fact, in static conditions, as representative of extremely low speed operations, the direction of jet thrust can be vectored essentially through a full 360deg because the two jets merge to form a free planar jet. The photographs in Fig. 40 show qualitatively the results obtained when sequencing through a range of pressure differentials between the upper and lower slots. The tests were conducted in air, and air pressure was used in the model. The wall jet vector directions are visualized with yam tufts. In the photograph at the top left, only the lower slot is active, and the ejected air follows the curved trailing edge and vectors out the leading edge of the wing. In the top center photograph, a small amount of upper surface slot flow is introduced, which results in lifting the wall jet off the wing surface. In the top right image, the upper flow is increased to produce a vertical thrust vector effect from the planar jet formed by the merger of the two slot flows. The bottom set of images represent a continued increase in the upper slot pressure until it equals the lower pressure

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single slot blowing (lower) 180 deg redirectionof the wall jet

Fig. 40 Model checkout in air and dual-slot operation with no freestream, 0.2 psig. The two wall jets merge to form a steerable planar jet.16

(bottom right photograph) and the jet flow direction is now 180 deg from the direction shown in the top left photograph. Verification and quantitative data for this thrust vectoring capability were measured in water using a load cell and revealed a thrust efficiency of 70-80%. Among the many findings from the CC-Hydrofoil test, the investigation demonstrated that 1) cavitation has a benign effect on the Coanda wall jet and there is no performance detriment with the onset of cavitation; 2) wake velocity profile filling is viable with dual slots; 3) a low flow rate from the second slot can eliminate one form of the CC lift limit; and 4) dual slots permit 0 to 360 deg static thrust vectoring and this merged-dual-jet mode may be viable as a jet flap for lift augmentation at extremely low speeds, where the coefficient of momentum would be too high for a viable CC mode.

VIII. Status of Design Capability To date, design implementation of CC technology has been based on the historical CC airfoil database and potential flow solutions (two-, three-dimensional) where local increment in lift is specified directly as an empirical relationshi between slot momentum flux and two-dimensional lift augmentation. R Whereas potential-flow-based techniques can readily address the “what-if” of a proposed CC application, they cannot guide the “how-to” in terms of trailing edge design details and the subsequent mass flow requirements, nor can they identify the performance boundaries. To support future CC applications, there exists a need for viscous-based computational codes to guide the subtle design details of the Coanda trailing edge region, as well as the upstream contour. Until a CFD code is available that has been properly and thoroughly validated for sensitivity to CC contour changes

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(airfoil data exists for this), design details will have to continue to be based on the historical database and intuition. Reliance on past practices as the only guide will result in a level of conservative design that may fail to uncover the full performance potential of lift augmentation by active flow control. Another benefit of future CFD will be the ability to develop additional insight into the fluid dynamics of circulation augmentation. For that insight to be valid, the challenge is to ensure that the numerically modeled flow physics is correct and not just fortuitous in producing correlation with a given set of experimental data. At NSWCCD, and other organizations, there are ongoing CFD CC validation efforts.

IX. Conclusions There have been many diverse experimental CC investigations at NSWCCD since 1986, the time period reviewed in this summary. Each of the projects built on lessons learned from previous experiments dating back to the 1970s for fixed and rotary wing, air and water applications. The experimental results increased insight into the fundamentals of CC and were used to correlate computational codes for the evaluation of various proposed applications. The six exploratory investigations revealed many new findings and, although similar in many ways, all were unique in basic geometry. Of the four dual-slotted models, one was a two-dimensional foil, two were three-dimensional foils (one with tangential mass ejection, the other with perpendicular ejection), and the fourth was an annular wing (duct). The dual slots provided either increased control range, extension of lift limit, or increased maneuvering steerage compared with a single-slot configuration, or, as in the case of the miniature oscillating valve, unsteady load reduction. Two of the investigations took place in water and four in air; however, each would be viable in either fluid regime. Of relevance to hydrodynamic applications, there had been the question of what would happen if cavitation occurred in the Coanda-slot region, which is where minimum pressure occurs. In a special series of experiments, it was revealed that the onset of cavitation did not have a disruptive effect on performance. Lift continued to be augmented in response to increased slot flow with an eventual smooth roll-off in lift as the cavitation became more extensive. Cavitation resulting from incorporating CC is not foreseen as an issue for presently contemplated hydrodynamic applications. Four of the investigations were low aspect ratio geometries and these helped to extend the understanding of the viability of CC on short-span surfaces. The successful omnidirectional demonstrations of the CC-Disc suggest future application to very maneuverable low AR vehicles or to appendages. It is helpful to know that analysis of the low AR investigations determined that there are no basic effects unique to wing lift developed by means of the Coanda form of CC, as compared to the classical lift development approaches. The TIPJET was the only rotary device reviewed and is unique in that it is driven in the rotary mode by the same air source that provides the blade lift augmentation. The findings from the TIPJET hover test were significant in furthering rotary-wing CC knowledge and in demonstrating a novel application of CC.

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In conclusion, the six diverse exploratory applications presented all met basic performance expectations, suggesting a certain degree of maturity of this technology. These experimental results, plus those reported by other organizations, now make for a rather extensive body of knowledge on the subject, ready to encourage additional creative applications and to support a variety of full-scale applications. Complementing this broad range of empirical knowledge, the anticipated emergence of thoroughly validated viscous-based two- and three-dimensional computational codes for CC will contribute to the achievement of the full performance potential of active flow control by efficiently allowing design refinement of how the circulation control effect is implemented.

References ‘Nielsen, J. N. (ed.), Proceedings of the Circulation-Control Workshop 1986, NASA Ames Research Center, NASA/CP-2432, Feb. 1986. 2 Englar, R. J., and Applegate, C. A., “Circulation Control-A Bibliography of DTNSRDC Research and Selected Outside References (Jan. 1969-Dec. 1983),” DTNSRDC-84/052, Sept. 1984. 31mber, R. I., “Exploratory Investigations of Circulation Control Technology: Overview for Period 1987-2003 at NSWCCD,” NASAICP-2005-213509, Proceedings of the 2004 NASAIONR Circulation Control Workshop, compiled by G. S. Jones and R. D. Joslin, March 2005. 4Abramson, J. S., and Rogers, E. O., “Design of a Circulation Control Airfoil Having Both Upper and Lower Surface Trailing Edge Slots (Model LSB17),” DTNSRDC/TM16-86/03, Sept. 1986. ’Abramson, J., “Characteristics of a Cambered Circulation Control Airfoil Having Both Upper and Lower Surface Trailing Edge Slots,’’NSWCCD-50-TR-2004/030, April 2004. 6Reader, K. R., Abramson, J. S., Schwartz, A. W., and Biggers, J. C., “Tipjet VTOL UAV Summary: Volume 1-1 200-Pound Tipjet VTOL Unmanned Aerial Vehicle,” DTRC/AD-89/01, Jan. 1989. ’Schwartz, A., and Rogers, E., “Hover Evaluation of an Integrated Pneumatic Lift/ Reaction-Drive Rotor System,” 30th Aerospace Sciences Meeting and Exhibit, AIAA Paper 92-0630, Jan. 1992. ‘Schwartz, A. W., and Rogers, E. O., “TIPJET VLAR UAV: Technology Development Status,” Presented at the 20th Annual Symposium and Exhibit of the Association for Unmanned Vehicle Systems, June 1993. 'Waiters, R. E., and Ashworth, J. C., “Experimental Investigation of a Circulation Controlled Shrouded Propeller,” West Virginia Univ., Morgantown, WV, TR-39, Feb. 1974. “Hoerner, S. F., and Borst, H. V. (ed.), Fluid-Dynamic Lift, Hoerner Fluid Dynamics, Vancouver, WA, 1985. “Rogers, E. O., and Abramson, J., “Selected Topics Related to Operational Applications of Circulation Control,” NASA/CP-2005-213509, Proceedings of the 2004 NASAIONR Circulation Control Workshop, compiled by G. S. Jones and R. D. Joslin, March 2005. ‘’Imber, R., and Rogers, E., “Investigation of a Circular Planform Wing with Tangential Fluid Ejection,” 34th Aerospace and Sciences Meeting and Exhibit, AIAA 96-0558, Jan. 1996. 13Fry,D. J., and McGuigan, S., “Hydrofoil Circulation Control Via a Miniature Valve for Alternating Flows Between Two Exit Slots,’’ CDNSWCISHD-1401-02, Dec. 1993.

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Fry, D. J., Louie, L. L., and Jessup, S. D., “A Water Tunnel Evaluation of a Novel Actuator and Active Control System to Cancel Unsteady Foil Forces,” CDNSWC/ SHD-1401-04, Dec. 1993. ”Louie, L., Fry, D. J., and Jessup, S. D., “An Active Control System to Cancel Unsteady Foil Forces,” DE-Vol. 75, Active Control of Vibration and Noise, American Society of Mechanical Engineers, New York, 1994. 16Rogers, E. O., and Donnelly, M. J., “Characteristics of a Dual-Slotted Circulation Control Wing of Low Aspect Ratio Intended for Naval Hydrodynamic Applications,” 42nd Aerospace Sciences Meeting, AIAA 2004-1244, Jan. 2004.

1I.A. Experiments and Applications: Fundamental Flow Physics

Chapter 4

Measurement and Analysis of Circulation Control Airfoils F. Kevin Owen* Complere Inc., Pacific Grove, California and

Andrew K. Owent University of Oxford, Oxford, England, United Kingdom

Nomenclature c = airfoil chord CL = lift coefficient CL, = infinite aspect ratio lift coefficient C, = blowing momentum coefficient U = mean axial velocity U, = edge velocity u' = rms velocity fluctuations x = streamwise position y = crosswise position a = airfoil angle of attack aeff= airfoil effective angle of attack ai= induced flow angularity

*Consultant. 'Research Assistant. Department of Engineering Science. Copyright 0 2005 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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I. Introduction IRCULATION control (CC) airfoil concepts have been studied extensively for more than four decades. These studies have included low-speed airfoil, helicopter rotor, and flight demonstrator configuration^.'-^ In these, and other studies of CC, the sharp trailing edges of otherwise conventional airfoils are replaced with rounded or bluff surfaces, typically with either circular or elliptic cross-sections, with thin tangential blowing slots located on the aft upper surface. These rounded trailing edges allow the rear stagnation point to move. This movement is controlled by the relative blowing momentum of fluid injected through the slots, and by the properties of the external flow field. By blowing through the slot, a jet sheet is issued, which, as a result of the balance of centrifugal force and subambient static pressure within the jet, remains attached to the airfoil. At low blowing rates, this Coanda effect entrains upper surface boundary layer flow and prevents trailing edge separation. As the blowing momentum is increased, the rear stagnation point is moved further around the trailing edge and the wake deflection angle is increased. An effective camber is introduced, and the lift is increased. Blowing rates can be adjusted until the airfoil static pressure distribution is that predicted by inviscid potential flow. With increased blowing, the jet controls the location of the airfoil stagnation points, and therefore the circulation and lift. However, eventually there comes a point where there is no longer a balance between the static pressure and centrifugal force and jet blow-off occurs, with a corresponding dramatic decrease in lift. Lift values greater than those predicted by inviscid potential flow theory are generated in the CC regime. Pneumatic camber similar to a mechanical high-lift system can be obtained. However, CC lift augmentation is far more efficient than conventional high-lift devices, because they only have to overcome the viscous losses in the flow. By compensating for the viscous losses, the flow field more closely resembles the ideal inviscid case. Accordingly, lift augmentation several times that attainable with jet flap or blown devices has been achieved. Unfortunately, the precise determination of CC airfoil performance for design and computational fluid dynamics (CFD) assessment purposes is difficult to achieve. The most serious problem encountered in testing these high-lift devices is the interference produced by wind tunnel test section wall separation. Owing to the strong adverse pressure gradients on the airfoil upper surface, strong secondary flows can be generated in the sidewall boundary layers. The problem is further compounded by significant spanwise circulation gradients, because circulation must decrease toward the wall. Even at moderate lift, these factors can generate trailing edge vorticity more characteristic of a three-dimensional than an infinite span wing. A great deal of research and analysis is still required in order to properly establish a reliable database for full-scale model development and CFD code validation. To address this shortfall, a wind tunnel investigation has been conducted of a two-dimensional CC airfoil section equipped with trailing edge blowing. The tests were conducted in the NASA Ames 2 x 2 ft Variable Density Transonic Wind Tunnel over a range of freestream Mach number and unit Reynolds numbers. Detailed nonintrusive flowfield measurements of the mean flow

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and turbulent properties were obtained in the airfoil wake for a number of different blowing coefficients. In this paper, some of these results have been related to the CC airfoil performance obtained from direct surface pressure measurements. The analysis shows that wind tunnel wall interference can have significant influence on high-lift test results. This influence must be accounted for before wind tunnel test data can be used for design extrapolation or for turbulence modeling and CFD assessments. Corrections have been made for finite aspect ratio (AR) wind tunnel wall interference in order to provide interference-free benchmark data for turbulence modeling and CFD code development and validation. 11. Experimental Details

The work described in this report was conducted in the NASA Ames 2 x 2 ft Variable Density Transonic Wind Tunnel at a freestream Mach number of 0.5 and at a unit Reynolds number of 3.2 x 106/ft. The test model spanned the test section and was held at zero angle of attack for the present work. The model was a symmetric 6-in. chord airfoil, 20% thick, 3% camber ellipse with a nominally circular arc trailing edge. An adjustable, nominally 0.010-in. tangential blowing slot was located on the upper surface, 1-2% before the usual upper surface separation point, at the 96% chordwise location. Transition strips were attached to the airfoil section at the 17% chord on both the upper and lower surfaces. The 1.25-mm-wide strips consisted of 0.13-mm nominal diameter glass beads. Transition effectiveness was verified by the sublimation technique. A regulated 3000psig air system was utilized to supply the internal plenum of the model, and a maximum internal pressure of 60 psig was attainable. The resulting high internal contraction ratio ensured adequate two-dimensionality of the jet exit flow. The jet exit velocity was calculated from isentropic relationships referenced to tunnel static conditions. There were a total of 91 pressure taps on the model, 59 of which were positioned along the centerline. Of these taps, 24 were on the upper surface and 35 on the lower surface. The airfoil performance data were obtained by direct integration of these centerline pressure taps.4 The flow-field measurements were obtained using a two-component laser velocimeter with conditional sampling ~apability.~ The effective sensing volume approximated a cylinder with a 200 diameter and 3 mm length, with its axis aligned with the cross-stream direction. Detailed measurements of the mean axial and vertical velocities, turbulent intensities, and turbulent shear stress distributions were obtained.

111. Sample Results Examples of laser velocimeter wake measurements at 5% chord downstream of the trailing edge for a zero angle of attack airfoil case are shown in Figs. 1 and 2. These results show the effects of jet blowing on the near-wake axial velocity profiles. In the zero blowing case, there is a small wake displacement due to the airfoil camber that produces lift at zero angle of attack. There is also a large region of reversed flow typical of a blunt body recirculation zone. With a small amount of blowing (C, = 0.024), there is a significant downward wake

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0.1 Wake Location, Y/C

,. . >

0.05

U

0 −0.2 0 −0.05

−0.4

s

B

0.2

0.4

0.6

0.8

1

1.2

−0.1 −0.15 Wake Velocity,U/U U/Uee

Fig. 1 Wake velocity profile (C, = 0, x/c = 0.05).

displacement and a fuller profile wake that has been energized by the jet. These data, and other profiles obtained at various distances downstream of the airfoil trailing edge, can be used to determine wake deflection angles, wake deficit recovery, and mixing and shear layer growth downstream. It is evident that the Coanda effect of the flowing jet relocates the aft stagnation point upstream along the airfoil lower surface, resulting in a downward displacement of the wake. This trend of increased wake displacement continues with higher blowing rates until there is no longer a balance between local static pressure difference and centrifugal force required for continued jet attachment. At this point we get what is referred to as jet blow-off. Jet effectiveness is destroyed, and there is a rapid drop in wake displacement and in the measured lift coefficient. However, in the present case, Fig. 3 shows that blowing can produce effective angles of attack, determined from measured wake displacement, of almost 20 deg in the present 0-deg airfoil angle of attack (AOA) case before blow-off occurs. From these data we can calculate the infinite AR lift coefficients from inviscid potential flow theory and thus assess airfoil lift performance. These results are shown in Fig. 4, where the lift at zero blowing agrees well with zero offset

Wake Location, Y/C

Y >

0.3 0.25 0.2 0.15 0.1 0.05 0 −0.05 0 −0.1 −0.15

0.2

0.4

0.6

0.8

1

Wake Velocity, U/Ue UlUe

Fig. 2 Wake velocity profile (C, = 0.24, x/c = 0.05).

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01

0

0.005

0.01

0.015

0.02

0.025

0.03 0.035

Blowing Momentum Coefficient, C,

Fig. 3 Measured wake angles.

camber predictions. However, these results are significantly higher than the lift computed from the measured airfoil surface pressure distributions. However, as expected, we have seen seed particle deposits on the test section windows, which suggest that strong secondary flows are generated in the wind tunnel sidewall boundary layers. This shed vorticity will induce unknown flow angularity in the freestream flow ahead of the model, thus changing the airfoil’s effective AOA. However, from the wake measurements, we are able to calculate the induced flow angularity as a function of jet blowing momentum coefficient. These results, calculated assuming a semi-elliptic lift distribution, are shown in Fig. 5 . With this information, we are able to compute the finite AR lift coefficients that are shown in Fig. 6 . These results are in excellent agreement with the surface pressure, direct lift measurements shown in Fig. 7. This comparison shows that sidewall effects are indeed significant, because agreement is not reached until an induced freestream downwash for a fully three-dimensional wing is introduced, that is, CL = 2m,f

2.5

01

0

0.005

0.01

0.015

0.02

0.025

Blowing Momentum Coefficient, C,

Fig. 4 Infinite AR lift coefficients.

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F. K. OWEN AND A. K. OWEN

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0

0.005

0.01

0.015

0.02

0.025 Blowing Momentum Coefficient, C,

0.03

0.035

Fig. 5 Induced flow angularity.

1.2

,

I

01

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

Blowing Momentum Coefficient, C,

Fig. 6 Calculated finite AR lift coefficients.

01

0

0.005

0.01

0.015 0.02 0.025 0.03

Blowing Momentum Coefficient, C,

Fig. 7 Measured lift coefficients.

0.035

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0.1-

40

-nit -".

I

I

18.

50

WakeTurbulence Level, u'/Ue%

Fig. 8 Wake turbulence profile (C, = 0, x/c = 0.17).

where

Wake turbulence measurements indicate that large-scale fluctuations are introduced by jet blowing and that wake unsteadiness may well be present at the higher blowing rates just before jet detachment. In the no blowing case shown in Fig. 8, small-scale turbulence dominates, and local RMS turbulence intensities are related to the local mean velocity gradients as in a planemixing layer. Thus, using the measured local turbulence levels and the measured local mean velocity gradients, we can calculate the effective mixing length for this flow. There is good agreement between this calculated mixing length to wake width ratio of 0.2 compared to the nominal value of 0.18 for a plane-mixing layer. However, once jet blowing is initiated, as shown in Fig. 9, a wide highly turbulent core develops that is indicative of high turbulent kinetic energy production in the blown jet wake. Turbulent length scales are increased by a factor of three, an indication of large-scale turbulent mixing andor wake unsteadiness.

0.1

2

$

-1 t

I

0.05-

.Q

-

Ll ~1

Y

i!

0 0 -0.05-

10

20

-0.1",

I"

WakeTurbulence Level, u'/Ue%

Fig. 9 Wake turbulence profile (C, = 0.009, x/c = 0.17).

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IV. Conclusions New CC test measurements and analysis have been presented that show the need for caution when attempting to use wind tunnel test results for CFD code validation, or for design purposes. In particular, the results have identified the quantitative extent wall influence can have on CC test results; for example, lift augmentation reduced from 68 to 42. The results also suggest that turbulence models must be modified to account for the effects of unsteady, large-scale turbulent mixing. The agreement between the measured and the calculated finite AR lift coefficients suggests that if we know the effective angle of attack, then simple inviscid theory may well be adequate for lift coefficient predictions. In turn, the analysis suggests that two-dimensional CFD computations could well be meaningless unless the airfoil effective angle of attack is known. Full three-dimensional calculations will probably be required to account for wall interference; that is, effective angle of attack and effective camber, especially at high lift. Estimates of the errors caused by non-uniform flow due primarily to wall boundary layer separation are essential. These initial investigations suggest that angle of attack corrections of at least - 1.5 CL will be required. Clearly, this can be a substantial correction factor, because lift coefficients well in excess of 2.0 can be expected for high-lift systems. Effects on the estimated drag coefficient are even more acute. Typical drag coefficients show errors of over 100% at induced angles of less than 1 deg. Indeed, at lift slopes typical of those at transonic speeds, angle of attack errors of 0.01 deg can lead to drag measurement uncertainty of more than one drag count. Clearly, in any high-lift experiments, accurate estimates or measurements of induced flow angularity must be made before useful design estimates or meaningful comparisons with CFD calculations are undertaken. A detailed review and analysis of finite AR CC experiments must be conducted to assess wind tunnel wall effects on experimental data previously reported in the literature. Although induced flow angularity is a fundamental consequence of the flow around finite AR lifting wings, our experiments and calculations show that these problems could be ameliorated to some extent by testing higher AR wings, and by measuring the induced flow angularity upstream. References ‘Kind, R. J., and Maull, D. J., “An Experimental Investigation of a Low Speed Circulation Controlled Airfoil,” The Aeronautical Quarterly, Vol. 19, 1968, pp. 170- 182. ’Cheeseman, I. C., and Seed, A. R., “The Application of Circulation Control by Blowing to Helicopter Rotors,” Journal of the Royal Aeronautical Society, Vol. 71, 1966, pp. 451-464. 3Englar, R. J., “Development of the A-6 Circulation Control Wing Flight Demonstrator Configuration,” DTNSRDC Rept. ASED-79/01, Jan. 1979. 4Wood, N. J., and Conlon, J. A., “The Performance of a Circulation Control Airfoil at Transonic Speeds,” AIAA Paper 83-0083, Jan. 1983. 50wen, F. K., “Application of Laser Velocimetry to Unsteady Flows in Large Scale High Speed Wind Tunnels,” International Congress on Instrumentation in Aerospace Simulation Facilities, Inst. of Electrical and Electronics Engineers Publ. 83CH1954-7, September 1983.

Chapter 5

Some Circulation and Separation Control Experiments Dino Cerchie,* Eran Halfon,+ Andreas Hammerich,* Gengxin Han,s Lutz Taubert,* Lucie-Trouve? Priyank Varghese,* and Israel Wygnanski** University of Arizona, Tucson, Arizona

Nomenclature c = chord length

C , = drag coefficient ( D / q c) CDp= form drag coefficient ( s ( p - p,) dy/q c) 2 1/2 C , = integrated force coefficient (Ct CO,) C, = lift coefficient ( L / q c) CMac= moment coefficient about the aerodynamic center C, = pressure coefficient ( p - p , ) / q CQ = steady volume flow coefficient (Q/SU,) C, = steady momentum coefficient [(2 h / c ) ( U s l o t / U,)’] (c,) = oscillatory momentum coefficient [(h/C)(USlotMax/ d = reference length, diameter f = frequency of excitation F+ = nondimensional frequency (f d l U,) h = slot height J =jet momentum q = dynamic pressure ( J p U k )

+

u,)~I

*Research Associate. ’Research Assistant. Currently at Tel-Aviv University, Ramat-Aviv, Israel. *Research Assistant. gPostdoctoral Fellow. TResearch Assistant. Currently at L’Ecole Nationale Supdrieure de Mdcanique et d’Adrotechnique, Poitiers, France. **Professor. Copyright 0 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Q = volume flow through the slot Ree = Reynolds number (U,8/u) U , = freestream velocity UJ = slot velocity UJ = maximum slot velocity a = angle of attack or slot location on a circular cylinder Sf = flap deflection 8 = angular distance from the leading edge of a cylinder

I. Introduction THIN jet being emitted tangentially from a slot milled in a circular cylinder or other convex, highly curved surface, alters its direction and wraps itself around the surface. A circular cylinder can turn a jet around and alter its direction by more than 180 deg. The centrifugal force acting on the deflected jet is balanced by the pressure difference between the surface of the cylinder and the ambient fluid. Integrating this pressure results in a force that is approximately equal to twice the jet momentum emitted at the slot (Fig. 1). Blunting a trailing edge of an airfoil and blowing over its upper surface will deflect the fluid downward, changing the “Kutta condition,” and provide a powerful means of increasing the usable lift. This is loosely referred to as supercirculation. One may divert the flow around a blunt trailing edge by using suction, as it was aptly demonstrated by Prandtl,’ who removed the boundary layer from one side of a circular cylinder and attached the flow on the side of the suction slot and generated lift. This idea

A

m

B

Fig. 1 Streamlines representing a wall jet flowing around a cylinder.

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was applied by Schrenk to thick airfoils that were otherwise plagued by early separation.2 Steady suction has been characterized historically by a volume flow coefficient CQ, because its primary aim was to remove low-momentum fluid from the boundary layer of a given freestream. Excessive suction could also provide circulation control (CC), which implied an increase of lift above and beyond the expected value generated by incidence and camber. The use of conformal mapping correctly predicted the lift generated by a strong, slot s ~ c t i o n , ~ which was directly proportional to the sink strength associated with the suction and depended on the location of the slot on the airfoil. The suction contribution to lift is given by ACL = 2 c Q cot(c$/2), where 4, in this case, represents the location of the slot in the mapped “circle plane”. The drag penalty associated with suction is very large (ACD= 2cQ), and it was theoretically predicted and experimentally verified by this model. Slot suction for the purpose of lift enhancement (CC) did not withstand the test of time because of the associated drag increase and the large ducts that were required to remove the low-pressure, external fluid. As the thickness of airfoils diminished with the quest to increase speed, they could not accommodate large internal ducting. Nevertheless, surface suction and multiple slot suction is still considered to be useful for drag reduction and for delaying transition to turbulence. The integration of propulsion with lift generation is a long-sought dream advocated by many researcher^.^ The advent of jet propulsion seemed to offer such an opportunity, but it quickly became apparent that materials withstanding the heat were too heavy and too costly for aeronautical applications. In most instances (the application to MIG-21 is an exception), only the compressed air generated prior to combustion by turbojet engines was ducted to slots and blown over flaps to augment their lift. A number of production aircraft used this form of lift augmentation (e.g., Lockheed F104 Starfighter, Blackburn NA39 Buccaneer, Dassault Etandard-IVM). In the application of blowing, a distinction is made between boundary layer control (BLC) and circulation control (CC). The first function of the jet, as it blows over the surface, is to increase the mean kinetic energy of the fluid within the boundary layer so that the latter may advance without separation into a region of rising pressure, for example, over the upper surface of a highly deflected trailing-edge flap. An adequate jet momentum is expected to generate a lift coefficient that is approximately predicted by a potential flow solution. In this regime of boundary layer control, the lift increment is roughly proportional to the first power of the jet momentum (ACL oc C,). An increase of jet momentum augments the lift further, but this augmentation is only proportional to the square root of the jet momentum (ACL oc JC,). This is the regime of supercirculation, where the jet departs from the trailing edge with sufficient downward momentum to increase appreciably the circulation around the wing. Poisson-Quinton4 is credited with establishing these criteria, as well as the critical value of (C,),,, that empirically determined the momentum required to pass from one flow regime to the other over an airfoil with a deflected flap at arbitrary angle 8, Circulation control may also be obtained by blowing the jet obliquely from the trailing edge of the wing, as was done on pure “jet-flap” experiments; however, there

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C, =0.24 Calculated using row of sinks

Fig. 2 Calculated and measured streamlines around a cylinder.

is a substantial gain in lift when the jet is blown over a suitably designed solid flap.4 A number of theoretical methods have been developed for predicting the ACL resulting from supercirculation. Stratford attempted to calculate the lift by assuming that the “jet-flap” was equivalent to a physical flap.5 More realistic assumptions were made by Helmbold,6 S p e n ~ eLegendre,* ,~ and Woods,’ who replaced the jet by a vortex sheet originating at the trailing edge. Woods used the hodograph method, whereas Spence7 and Malavard” linearized the problem, assuming small incidence and small jet deflection. In all the theoretical models, the mixing of the jet with the ambient flow is neglected. In reality, the jet entrains fluid from its surroundings and that entrainment is well represented by placing a suitable distribution of sinks along its path” (Fig. 2). When a strong jet flows over a curved flap or the upper surface of an airfoil, this distribution of sinks contributes to circulation,’2 which is also proportional to JC,. When the jet is emitted from the trailing edge of bluff bodies (e.g. circular or elliptic cylinders), the entrainment that takes place on both sides of the jet contributes to form drag.” Some aspects of the ideal flow models are controversial and they have not been entirely resolved to date, for example, the prediction that the entire jet momentum should be recovered as thrust regardless of the jet’s initial inclination angle relative to the oncoming stream. This result was proven experimentally up to a flap deflection of 60 deg, at which approximately 90% of the jet momentum was recovered as thrust as long as the value of C, was quite large. At larger flap deflections, the C, required to overcome separation and other “real flow” effects (mixing) became excessive, and the thrust recovery almost entirely vanished when the flap deflection exceeded 90 deg. The effects of steady blowing, steady suction, or periodic excitation on circulation and drag are assessed presently. This report represents an ongoing research with the purpose of improving our understanding of each technique and to sorting out the leading parameters that affect, control, and manipulate the flow. We shall start by examining the flow over a flapped, conventional, symmetrical airfoil, the NACA 0015 (Fig. 3a). The Kutta condition is fixed and the impact of the increased circulation is easily recognized when compared to the standard airfoil performance. Thereafter, we have replaced the normal, 26% chord simple flap with a stubby, 8% chord flap consisting of a circular cylinder that blends into a wedge having an included angle of 40 deg at its

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a)

b) Stream

Fig. 3 Airfoil models used.

trailing edge. This configuration was extensively studied at S t a n f ~ r d in ' ~ conjunction with strong steady blowing. The circular trailing edge facilitates the generation of supercirculation, but the trailing edge wedge predetermines the Kutta condition provided the flow over the flap is attached (Fig. 3b). The flow over the small, blunt, and concave trailing edge brought into focus the need to investigate the controlled flow over a concave surface in the presence of adverse pressure gradient more extensively. Such flows were investigated over wall-mounted humps, started by S t r a t f ~ r d , 'who ~ coined the concept of a boundary layer that is maintained on the verge of separation over an extensive distance. When periodic excitation was applied to such a boundary layer,15 the skin friction was increased while the shape factor was reduced, and it thinned and stabilized the boundary layer and enabled it to better overcome the imposed pressure gradient. If the pressure recovery region at the rear of the hump is made steeper, the boundary layer separates, but it has to reattach farther downstream due to the presence of the long flat surface that extends beyond the trailing edge of the hump. The control of this flow is reduced to control of a separation b ~ b b 1 e . l ~Because ~ ' ~ the hump used in Refs. 16 and 17 is based on Glauert's GLAS I1 airfoil (GLAS stands for Glauert's Laminar Airfoil Section), it was investigated in the present context (Fig. 3c). The flow around a thick elliptical cylinder was later examined. Its maximum thicknessto-chord ratio is 30%, and its leading and trailing edges are circular. This geometry easily lends itself to a change in the actuation location and in the slot width. The pressure gradient near the leading edge resembles the pressure gradient experienced by a standard airfoil, whereas the flow near the trailing edge is complicated by the fact that the Kutta condition is not well defined. The circular cylinder was the last test article to be examined, because it is the most widely researched flow, but it might be the most difficult one to control. The Kutta condition is not determined and the parameters affecting flow reattachment interact and affect the circulation in a more complex manner than on previous configurations due to the strong coupling between the flows near the leading and trailing edges. It is believed that by increasing the complexity and the number of degrees of freedom that are associated with the various configurations selected, the dominant variables controlling the flow will be identified. Typical questions to be answered include the following:

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1) What is the best method or a combination of methods to increase lift? 2) Is C, the unique parameter that governs BLC and CC control and are they occurring sequentially as C, is increased beyond a prescribed threshold level ( CJcrit?

3) Is separation effectively controlled by suction? 4) Is C, displaced by CQ when suction is used for LBC or CC? 5) When and why is periodic excitation (active flow control, AFC) more effective than blowing or suction? 6) How sensitive is each method to the location of the actuation, and how is it affected by the configuration on which it is employed? The present chapter focuses on some of these questions, in an attempt to categorize the effects of the leading parameters in a rational manner.

11. Discussion of Results

A. Flow Control over an Airfoil with a Conventional Flap Most aerodynamic control of lift experiments begin with a standard NACA airfoil and then either progress in the direction of more custom lofting, lift augmentation devices or flow control to achieve not only the desired loads, but more favorable distribution of the load along the airfoil surface. We will discuss the impact of the total load and distribution of the load on a standard airfoil using both a trailing edge flap and flow control. Data were collected using a NACA 0015 airfoil with a simple 26% chord flap at Re < 5 x lo5. A schematic drawing is included in Fig. 3a, showing a cross-section through the airfoil model. Some early observations carried out by Greenblatt and Wygnanski indicate that the flow over a deflected flap at Sf= 20 deg separates around a = -2 deg." Even at a = 0 deg, both steady blowing and periodic excitation are beneficial. Consider injection of momentum at C, = 3% (Fig. 4). For a flap deflection of Sf = 20 deg, both steady blowing and periodic excitation at very low frequency generate a lift increment of ACL = 0.5 relative to the baseline airfoil performance, whereas periodic excitation at F+ = 1.1 generated an inferior lift increment of only ACL = 0.35. Repeating the same experiment at a lower C, of 1.2% shows slightly lesser periodic excitation performance at F+ > 1.1, and even poorer steady blowing performance (see Fig. 5 for Sf= 20 deg). At Sf= 35 deg and at the high C, of 3%, both steady blowing and low frequency excitation (F+ = 0.3) peaked out by generating ACL = 0.4 and 0.52 relative to the baseline flapped airfoil, respectively. An increase in the flap deflection beyond this angle caused a reduction in the lift increment generated by steady blowing until, at Sf> 50 deg, the injection of steady momentum became detrimental to the generation of lift (i.e., the baseline CL exceeded the value obtained by using steady blowing). The efficacy of the low-frequency periodic excitation at C, = 3% did not deteriorate with increasing flap deflection beyond Sf= 35 deg, whereas the excitation at the higher frequency of F+ = 1.1 improved with increasing flap deflection until the two curves crossed over around Sf= 65 deg. At the lower level of C, = 1.2%, the increase in flap deflection beyond Sf= 35 deg rendered the steady blowing

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1.61.41.2-

Re = 300K -#-Baseline AFC = 3%, F+= 0.3

1.o-

I*

+AFC=3%,F -#- Blowing C = 3%

0.8-

+=1.1

P

0.6 !

I

20

I

I

60

40 Flap deflection Sf(')

Fig. 4 Effect of blowing and AFC at C, = 3%, as a function of flap deflection on a NACA 0015 at a = 0 deg.

ineffective, if not detrimental, whereas even higher frequency excitation remained effective (Fig. 5 ) . The pressure distribution associated with the three modes of flow control at C, = 3% and Sf= 35 deg is plotted in Fig. 6 . The constant, low pressure on the upper surface of the flap ( x / c > 0.75) indicates that the baseline flow was

'L

1.81 1.6

1

1.4 121.o -

Re = 300K f

AFC = 1.2%, F+= 1.1

P

"." .

+

+AFC = 1.2%, F = 2.5

0.8 -

--C Blowing C I

20

I

40 Flap deflection

P

= 1.2% I

4(")

60

Fig. 5 Effect of blowing and AFC at C, = 1.2%,as a function of flap deflection on a NACA 0015 at a = 0 deg.

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120 C

P

-3

-2

Re = 300K α = 0°, δ f = 35° Baseline + AFC = 3%, F = 0.3 + AFC = 3%, F = 1.1 Blowing C µ = 3%

-1

0 0.0

0.5 X/C

1.0

1

Fig. 6 Pressure distribution over NACA 0015 with 26% chord deflected flap.

totally separated over the deflected flap. The flow was partially attached by the periodic excitation at F+ = 1.1 and completely attached by the low-frequency excitation at F+ = 0.3 and by the steady blowing. The reattachment of the flow over the flap changes the circulation around the airfoil and has a far-reaching effect on the upstream pressure distribution all the way to the leading edge of the airfoil. The acceleration of the flow upstream of the slot is of particular interest in this case. It seems reasonable to examine various flow control mechanisms providing identical circulation and C ., This approach is of practical interest, because a potential designer may be required to generate a prescribed lift by various techniques available and should be familiar with the consequences, such as drag, pitching moment, momentum input, and so on, associated with generating the required lift. During the experiments discussed here, the flap was deflected at two angles; Sf= 20 and 40 deg, with periodic excitation being applied through a 0.06 in. slot at the interface of the main element and the flap shoulder. Figure 7 includes two pairs of angle of attack (AOA) sweeps with and without AFC (periodic excitation) at the two different flap deflections discussed previously. Some features are immediately apparent in this figure. All of the configurations share the same dCL/da when a > 0 deg. The deflection of the flap on the model increases the effective camber of the model, even if the flow over the flap is separated, causing a shift upward (or to the left) of the C, vs a curves. However, when the flow over the flap separates (see curve corresponding to Sf= 40 deg that uses AFC in Fig. 7),there is a shift to the right with a dCL/ d a % 0 in the range -4 < a < -2 deg. This effect is not seen as clearly for the baseline airfoil sweep with Sf = 20 deg because of the low Re of the experiment, although the flow separates partially from the flap a x - 2 deg. The other two curves, plotted in Fig. 7, are not expected to have a discontinuity in dCL/da. The baseline flow over the flap that is deflected at Sf = 40 deg is separated over the entire range of a, considered, and the periodically excited flow at Sf = 20 deg is attached over the flap until a = astall. The excitation level for Sf = 20 deg (F+ = 0.9, (c,) = 2.2%) was specially selected in order to overlap the lift

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121

CL 2

1

+

Re = 200K F = 0.9 δ f = 20° Baseline δ f = 20° = 2.2% δ f = 40° Baseline δ f = 40° = 2.2%

0 -10

0

α (°)

10

20

Fig. 7 NACA 0015 airfoil performance with AFC.

curve generated when the flap deflection of the basic airfoil was Sf = 40 deg. This enables a detailed comparison to be made between the effect of flap deflection and periodic excitation. In fact, the curve for Sf= 20 deg with AFC falls on top of the curve for S,= 40 deg without AFC until the occurrence of stall. Although the stall angle is somewhat higher for Sf= 20 deg in conjunction with AFC, the resulting CLma,is approximately the same for both cases. When the same AFC is applied while S,= 40 deg, the maximum lift coefficient, C,, = 2.25 at a = 10 deg. At negative angles of attack (-8 < a < -4 deg), the ACL generated by the application of AFC to S f = 40 deg is commensurate with the ACL observed at 20 deg flap deflection for a < astall, because the flow over the flap is attached for both 8, values in the respective range of a. Inspite of the flow separation from the upper surface of the deflected flap at S,= 40 deg, the airfoil continues to generate a higher lift than for Sf= 20 deg, primarily due to the deflection of the flow by the lower surface. In the discussion that follows, we examine pressure distributions measured on the surface of the airfoil model, which produced three different lift coefficients (CL= 1.0, 1.35, 1.5). These “sectional” cuts through the (CL vs a ) curves show the different approaches that a designer could select to produce a specific lift and are marked in Fig. 7 to aid the reader. We consider this to be an important technique to evaluate different flow control strategies, rather than simply look at the relative benefit in performance that the control can provide at a fixed geometric configuration. Figure 8 shows four pressure distributions that generate c, = 1.0. In the absence of AFC and with the flap deflected to 40 deg, a slight pitchdown attitude ( a = -2 deg) generates a small suction peak near the leading edge (LE) and mild adverse pressure gradient along the upper surface of the entire main element. The flow over the flap is undoubtedly separated and, consequently, there is a drag penalty associated with this configuration. When Sf= 20 deg for the baseline airfoil, the incidence must increase to a = + 2 deg in order to

D. CERCHIE ET AL.

122

Cp

-2

Re = 200K F+ = 0.9 δ f = 20° α = 2° Baseline δ f = 20° α = -2° = 2.2% δ f = 40° α = -2° Baseline δ f = 40° α = -8° = 2.2%

-1

0 0.0

0.5

1.0

X/C 1

Fig. 8 Pressure distribution that develops at C, = 1.0 for the NACA 0015.

maintain the same lift, and a larger suction peak is created at the LE, while separation on the flap is pushed slightly farther downstream. When the appropriate level of AFC is applied while maintaining Sf = 20 deg, the AOA can once again be returned to a = - 2 deg, resulting in a more uniform pressure distribution over the upper loft and reducing the suction peak near the LE. Because a is the same for the two flap deflections as the total circulation, the flow near the LE is identical over the upper surface, as is the pressure distribution in the range 0 < x / c < 0.4 for the two cases considered (Sf= 40 deg and Sf= 20 deg with AFC). At x/c > 0.5 and in the absence of AFC, the pressure remains constant over the upper surface of the airfoil and the deflected flap, because it is dominated by the “base pressure” (C p RZ - 0.7) of the recirculating region downstream. On the other hand, when AFC is applied and the flap is only deflected at 20 deg, the flow accelerates upstream of the slot (i.e., for 0.6 < x/c < 0.74). The flow over the flap is fully attached with a pressure coefficient at the trailing edge being positive ( C p x 0.25), suggesting that the flow downstream of the trailing edge continues with its downward momentum, generating perhaps a “jet flap” effect. Increasing the flap deflection to 40 deg while maintaining the AFC results in an attached flow with C p x 0 at the trailing edge (TE), while heavily loading the aft region of the flapped airfoil. This is achieved while maintaining a favorable pressure gradient over the entire upper surface of the main element by placing the airfoil at a = - 8 deg. In this case the flow acceleration upstream of the slot is magnified. This behavior would be especially advantageous for laminar flow applications where delay of transition is important. It is easy to identify the upstream influence of the AFC along the upper surface of the main element. Figure 9 shows the pressure distributions over the model that generated CL = 1.35. The two cases that share the same angle of attack and lift (a= 2 deg, S - 20 deg, C, = 2.2% and Sf= 40 deg, C, = O%), indicate that fthe pressure distributions on both the upper and lower surfaces over the upstream half of the airfoil are almost identical. The case with AFC shows the flow over the

CIRCULATION AND SEPARATION CONTROL EXPERIMENTS C p -4 CD

123

Re = 200K F+ = 0.9 δ f = 20° α = 8° Baseline

-3

δ = 20° f δ = 40° f δ = 40° f

-2

α = 2° < C > = 2.2% µ α = 2° Baseline α = -4° < C > = 2.2% µ

-1 0 0.0 1

0.5

1.0

X/C

2

Fig. 9 Pressure distributions that develop C, = 1.35 for the NACA 0015.

flap is attached with a pressure coefficient close to zero at the TE. The heavily deflected flap in the absence of AFC has the same C p % -0.7 at the TE as it did at CL = 1.0, indicating that in both cases the recirculating wake region has approximately the same dimension. It implies that the flow over the flap was completely separated at both angles of incidence and that the additional lift was generated by the main element. Once again, the flow with AFC accelerates before reaching the slot. This reduces the adverse pressure gradient on the upper surface, making the airfoil less susceptible to stall at this value of CL.The basic airfoil with the flap deflected at 20 deg also generates CL = 1.35, but at a = 8 deg. Under these conditions, the flow is still separated over the flap, but the wake is narrow as evidenced by Cp % - 0.2 at the TE. One may reattach the flow to a deflected flap at 40 deg through AFC enabling CL = 1.35 at a = -4 deg, due to the increased suction on the aft portion of the main element upstream of the slot (i.e., at 0.4 < x / c < 0.74). In conclusion, the results described in Fig. 9 are similar to those associated with C L = 1, but with the effect of AFC being accentuated. Therefore, not only does the AFC prevent flow separation downstream of the actuation location, thereby increasing the circulation, but it also lowers pressure on the upper surface upstream of it, enhancing the lift. The prime benefit is in the form-drag reduction, which was reduced by a factor of four in the range 0.5 < CL < 1. The total drag was reduced only by a factor of two and there is some uncertainty in the drag estimate. Nevertheless, the effect of AFC is significant and will be discussed in full in Section D, p. 144. The pressure distributions at a given CL,with AFC being applied, suggest that a can be significantly reduced depending on the deflection of the flap needed to produce the same lift coefficient. For the lower flap deflection case, the suction peak is reduced by approximately 40% and the flow over the flap is fully attached with a TE pressure coefficient close to that of the freestream. At the higher flap

124

D. CERCHIE ET AL.

deflection, the suction peak is further attenuated when sufficient lift is generated, even at zero AOA. Although the flow over the flap may not be fully attached, the upstream effect of the AFC is strong enough to load the main element sufficiently to generate the necessary lift. This behavior is not unique to this airfoil. One can conclude from the data presented that AFC contributes through three distinct mechanisms to airfoil performance: first by preventing flow separation on a deflected flap (this mechanism was investigated by Nishri and Wygnanski” and Darabi and Wygnanski2’); secondly, by enhancing circulation through the inviscid jet flap effect; and, thirdly, associated with turbulent entrainment of the flow and the reduction of the static pressure both upstream and downstream of the slot location. Poisson-Quinton, in Ref. 4, identified the first two mechanisms using steady blowing. Acceleration due to entrainment, while being present in those cases, is more prominent when oscillatory excitation is used. The demarcation between (BLC) and (CC) was quite well defined when steady blowing or suction were used to control the flow, because CC implied “an artificial increase in circulation over that which could be expected from incidence and camber in unseparated flow”.21 Because the lift over streamlined bodies (over which the flow is totally attached) is predicted by inviscid solution, a comparison of pressure distribution both measured and calculated was essential. The pressure distributions calculated from viscous and inviscid solutions using the “Xfoil” program, assuming that the flow is entirely turbulent in the viscous case, are plotted in Fig. 10, together with the measured results with and without the use of AFC. When Sf = 20 deg, a = 2 deg and, in the absence of actuation, the experimental results agree quite well with Xfoil’s viscous prediction, with the exception of the base pressure observed over the separated flap. The experimental data for the forced flow suggest that the flow over the flap was attached as a result of the excitation and, as a consequence, the pressure over the entire upper surface was reduced (Fig. 10). The measured pressure distribution resulting from excitation at (CY) = 2.2% at F+ = 0.9 fell short of the expected inviscid values, suggesting that this level of excitation is below the (CJcrit that separates the BLC and the CC regimes. Similar conclusions may be drawn for the results obtained for Sf= 40 deg and a = -4 deg, except that the ideal flow solution overpredicts the forced results by a larger amount, and the viscous solution does as poorly in predicting the base-flow pressure distribution. Both examples (Fig. 10) confirm the suggestion that periodic excitation, at the level and frequency used, keeps the flow attached (i.e., controls separation), but does not enhance the circulation above the normal inviscid limit. A complimentary example where the enhancement of circulation was achieved without reattaching the flow over the flap was provided by H. Nagib (personal communication, 2003) who examined the control of the flow over a three-element airfoil with a slot located at the shoulder of a highly deflected, simple flap. In this case, periodic excitation affected the pressure distribution on the main element and on the leading edge slat without causing reattachment to the flap itself (Fig. 11). Circulation was increased and the pressure upstream of the actuation was lowered, in spite of the fact that the pressure over most of the flap remained unchanged. Because the upstream effect of AFC may be more significant than the downstream effect, it is possible that the demarcation

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125

NACA 0015, a=2', &m=200 CP

Fig. 10 A comparison between pressure distributions calculated using viscous and inviscid solutions and the measured pressures on the basic airfoil and after using AFC (all data taken on the NACA 0015).

between BLC and CC is artificial and that the effect of AFC on circulation occurs concomitantly with BLC.

B. Control over a Truncated NACA 0015 Flapped Airfoil Typical CC airfoils have circular trailing edges in order to make use of the Coanda effect and generate maximum circulation. The addition of a cusp provides the ability to control the angle at which the jet departs from the trailing edge and control the circulation without altering the blowing emanating from the slot. The truncated NACA 0015 airfoil shown in Fig. 3 was widely used in the verification of jet-flap concepts.13 In the absence of strong blowing, the airfoil does not perform very well. Its maximum lift coefficient is, CL % 1.2 in

D. CERCHIE ET AL.

126

Rec = 0.75e6 Flap = 40 deg. alpha = 13 deg.

ADVINT/ATT 5% Model in NDF at IIT Nagib & Kiedaisch; 2002

Baseline, slat 2 F = 120 Hz, Uj/Uinf = 2.8

Cp CP Slot Location for AFC

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

x/c

Fig. 11 Pressure distributions over a three-element airfoil with and without CC (H. Nagib, private communication).

the absence of flap deflection (Fig. 12), but the rounded trailing edge generates a large form drag (Fig. 13) and a wider wake than is generally expected from the NACA 0015 at a given Re. In fact, by deflecting the flap to 15 deg, the C ,,, as well as the C, attained at small angles of incidence, is slightly reduced, but this reduction does not affect the form drag or even the total drag. Strong blowing approaching C, % 1 has been used in previous experiments for CC.13 In the present investigation C, 5 0.1, in order to use the upper limit of C, = 0.1 for comparison with data acquired by Hynes.13 For C, = lo%, CLmaxis increased to 1.5 in the absence of flap deflection, and it attained C , = 2.5 for S,= 60 deg (Fig. 12). In the absence of blowing, the minimum form drag is attained at incidence 4 < a < 6 deg, regardless of the flap deflection (Figs. 12 and 13). The total drag was determined from wake surveys and corrected for buoyancy (both Betz's and Jones's corrections were used; however, there was no difference between the two methods of correction). It behaves in a similar manner to CDp provided 8, > 30 deg (Fig. 14). It is interesting to note that the total drag is always lower than the CDpand the difference between the two increases with increasing SF Because the skin friction drag is generally positive, there has been a search for experimental error and uncertainty. It is possible that the number of pressure taps near the leading edge is insufficient, but it is equally plausible that vortices shed from a lifting airfoil over which the flow is partially separated (either due to high incidence or large Sf) increase CDp,making CDp> C ,. This excess is more noticeable whenever AFC is used. The constant presence of vortices downstream of

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127

C, vs. a,Baseline &Steady Blowing

Fig. 12 C L - a curves on the truncated NACA 0015, for C, = 0% and C, = 10%.

a bluff body induces low “base pressure” near the base of such a body and contributes to form drag. This possibility should be examined more closely in the future, particularly near bluff bodies where the skin friction drag is negligible relative to CDp. We shall now focus on the effect of increasing C, on the characteristics of the airfoil when Sf= 30 deg, at a constant representative C, = 1. In the absence of blowing, CL = 1 is attained at a x 7 deg, but at C, = 0.1 it is achieved at a x -0.7 deg (Fig. 12). There is a coupling between a and the C, necessary to provide the required lift. This relationship is not linear (Fig. 15a), although it is explored in the region where dCL/da is constant (Fig. 12). The highest effect on reducing the incidence required to generate the necessary lift corresponds to 0.025 < C, < 0.075. The moment coefficient about the quarter chord location (the aerodynamic center) behaves in a similar manner, implying

D. CERCHIE ET AL.

128 C, vs. C,,,

Baseline, Uh+=12m/s, Re=2.P105

Fig. 13 Form drag polars for C, = 0% and C, = 10%.

that a desired pitching moment can be obtained at a prescribed lift by trading incidence with C, (Fig. 15a). It is interesting to note how the form drag increases with increasing C,, (Fig. 13), whereas the total drag turns to thrust with increasing C, (Figs. 14 and 15b). The increase in C D p is attributed primarily to the low pressure generated on the convex surface downstream of the flap shoulder due to the Coanda effect. The concave surface resulting from the presence of the cusp generates positive pressure, but the surface is too small to affect the CDp in a meaningful way (Fig. 16). The increase in incidence necessary to generate the proper CL at lower values of C, also results in an increased suction at the LE and a reduction in C D p . One may now examine the lift increment generated by increasing C, while maintaining a and afconstant (Fig. 17). It is interesting to note that at low levels of C,, ACL cc C;, and only at higher C,, it becomes linearly dependent on this parameter. This is contrary to the accepted n o t i ~ n ~ ” ’ ~ ’ ’ ~

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129

CL vs. CD,Baseline, Ui,=12mls, Re=2.7*105

Fig. 14 Drag polars for C, = 0% at four flap deflections and for 8f = 30 deg, but for O
that in the BLC regime (i.e., at small C,) ACL cc C,, and it only becomes proportional to JC, in the CC regime. The difference may stem from the low values of C, that are presently considered. The critical C, distinguishing between the two regimes is 3.5 < C, < 5.5% in the range of flap deflections investigated. Suction at equivalent C, = 10% generates lower lift than blowing, although the stall angle increases somewhat by using suction. A comparison between the two methods is shown in Fig. 18 for S f = 30 deg. Whereas, for blowing, most of the added momentum coefficient is manifested as thrust (the CD of the baseline airfoil is 0.04 while with C, = 10% the CD x -0.06), for suction CD x 0.02, implying that some 80% of the suction momentum generated drag. It is interesting to note that, based only on CDp,suction generates apparent thrust, whereas blowing increases the drag. The explanation for this can come from comparing the pressure distributions corresponding to C, = 1 (Fig. 19). In one case there is a negative pressure peak over the deflected flap, whereas in the other it occurs near the leading edge.

D. CERCHIE ET AL.

130

a& Pikhing M o m e n t a C14 vs. Cp Variation of a & Pitching Moment for fixed CL=l FIap=3Oo

a) 9 8 7 6

..............

5 a 4

3 2 1 0

-1 -2

..............

.

J

.

.

I

4

0

6

> 4.5 10

8

c, [",'.I CDp & CrJ vs. C, @ C p l for various C,, %hap=300

b)

0.15

...... ...... ...... ................. ......A

........... , .......,......7 ............

0.1J 0

1

2

3

4

5

. . . I.

6

+CDp ..; .. - +- CD __.I

- - - -

.....:.

c-

......,......,

7

8

9

i 10

c, [%I Fig. 15 Dependence of a) (Y and CM,, and b) CD and CDpon C, at CL = 1.

C. Controlling the GLAS I1 Airfoil The Glauert Laminar Airfoil Section I1 (GLAS 11) has a maximum thicknessto-chord ratio of 31.4% and was designed to operate with massive suction through a slot located at 69% of the chord. The designer intended to have favorable pressure gradients over most of the upper and lower lofts, and maintain

CIRCULATION AND SEPARATION CONTROL EXPERIMENTS

131

C, profiles for different C,,,CL-l, qlSp=3O0

Fig. 16 Pressure distributions for various values of C, at C, = 1.

laminar flow over a large fraction of the airfoil surface. Suction provided a pressure discontinuity across the slot that led to a positive pressure along the entire concave recovery ram terminating with stagnation pressure at the trailing edge. With adequate suction!’ the measured L I D varied between 250 and 550 for C, > 1 and Re x lo6. In the absence of suction, L I D > 12 for the same Re, but C, was reduced to C, x 0.6. The L I D increased to approximately 30 at C,Sweep, Steady Blowing for different Flap Angles

Fig. 17 Dependence of CL on C, at (Y = 3 deg.

D. CERCHIE ET AL.

132

CL vs. a for various C,

CL vs. Cm and C,,

Fig. 18 Comparison between the performance of the airfoil using strong suction or blowing at C , = 10%. C,

- Comparison between Steady Suction & Blowing fixed C,-I ,C,=lO%,~,,,=3O0

Fig. 19 Representative pressure distributions on the truncated 0015 airfoil using suction at C , = 10%.

CIRCULATION AND SEPARATION CONTROL EXPERIMENTS

133

Re % 3 x lo6. It appeared that the flow was intermittently reattaching to the ramp just upstream of the TE, resulting in large drag oscillations. Blowing appeared to be less effective than suction, requiring larger mass flux to forcibly reattach the flow. Either way, the C, required to keep the flow attached was in excess of 20%. Many additional research papers followed, culminating in test flights on a glider. If comparable performance could be achieved using periodic excitation at considerably smaller C, levels, the use of GLAS-type airfoils could be revived. Furthermore, there has been a considerable interest in a hump that was placed on the wind-tunnel wall and whose shape represents the upper loft of a GLAS I1 airfoil for validation of CFD codes. The differences between the flow over such a hump and over the airfoil should be fully understood and properly documented. The dependence of CLon a is plotted in Fig. 20 for the baseline configuration at Re < 0.5 x lo6. dCL/da is not constant at these low Reynolds numbers, but the discontinuity in the slope diminishes with increasing Re. At Re = 1.17 x lo5 there is a sudden increase in the lift at a % 20 deg. With increasing Re, the kinks in the CL-a curves occur at lower incidence (e.g., a % 16 deg at Re % 1.7 x lo5) and they become more moderate. Also, the maximum lift experienced by the airfoil decreases from being CL,, = 1.7 at Re=1.17 x lo5 to C , = 1.05 at Re = 4.8 x lo5. The data acquired at the Compressed Air Wind Tunnel of the W L Z 3agrees fairly well with the present results. The CL-a curve reproduced from Ref. 23 was taken at Re = 4.06 x lo5 and could be obtained by interpolating the present data taken at 3.5 x lo5 < Re < 4.8 x lo5 (Fig. 19). No wind tunnel corrections were applied to either set of data. Figure 21 shows the typical pressure distributions measured on the surface of the baseline airfoil at 1.2 x lo5 < Re < 4.8 x lo5, corresponding to CL % 0.5, suggest that a laminar boundary layer could be maintained over the lower loft

C,

YS.

a - Baseline

Fig. 20 Dependence of C, on (Y and on Re for the baseline airfoil.

134

D. CERCHIE ET AL.

Baseline C, profiles at q 4 . 5 for different Re

Fig. 21 Pressure distribution measured on the baseline airfoil at constant C, but different values of Reynolds number.

of the airfoil up to x/c ranging between 0.5 and 0.7 due to the favorable pressure gradient existing on that surface. On the upper loft, however, the location where the C p is minimum depends strongly on Re as it changes from x/c = 0.07 at the lowest Re used presently tox/c = 0.33 at Re = 4.8 x lo5. The C p measured near the TE indicates that the flow is mostly separated in this region, although the base pressure increased with increasing Re suggesting that the mean size of the separated region was reduced. Applying the strongest available suction (C, x 19%), blowing (C, x 22%), or periodic excitation (at (C,) x 2.1%) to this airfoil at x/c = 0.62 and Re x 1.2 x lo5 results in a tremendous increase in lift and the straightening of the CL-a curves (Fig. 22). The actuation location was moved upstream because the levels of actuation mentioned above were unable to affect the flow at small angles of incidence at the original slot location suggested in the literature. The Australian researchers faced similar difficulties and they, too, moved the suction location. The chosen location of the slot corresponds to the separation line predicted by CFD and it will be discussed separately. Contrary to the observations of Glauert et a1.,22 blowing was more effective than suction, approximately doubling the CL attained at a < 10 deg. The maximum CL was obtained in both instances at a,,, = 24 deg; however, in the case of blowing, CL,, = 5, whereas by using suction, CLm,, = 3.5 only. Periodic excitation at F+ = 0.7 and approximately 1/10 of the steady momentum input for blowing performed almost as well as the steady blowing up to a x 14 deg. At higher incidence, the difference in C, between the steady blowing and the periodic excitation became noticeable, yielding CLm,, = 3.5 for the periodic excitation. Because the C,, obtained for the basic airfoil was only 1.7, suction and periodic excitation represent 100% increase in CLm,,, whereas the much stronger steady blowing represents almost 200% increase in this coefficient.

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135

Fig. 22 Dependence of C, on a and on the control parameters used and the corresponding drag polars for Re = 1.17 X lo5.

The drag polars for the corresponding cases are also plotted in Fig. 22. They reveal very interesting features associated with each method of boundary layer and circulation control. The baseline drag at CL < 0 is approximately C , % 0.1 1. This number agrees very well with the National Physics Laboratory (NPL) results at comparable Re and C, values.23 The baseline pressure drag CDp= 0.086 was almost constant for all C, < 0, and it implies that the skin friction drag is approximately 0.024 for the negative lift coefficients. The application of suction at a < 6 deg generated a small dCL/da and it possessed relatively large C ,. For C, > 0.4, suction reduced the drag to C, x 0.02. This represents a substantial drag reduction relative to the baseline airfoil that attains C, x 0.4 around a x 20 deg. The form drag in this case is negative, indicating that there is a low pressure at the leading edge of the airfoil and a high pressure on the rear ramp. Strong, steady blowing generated thrust in addition

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to lift enhancement (Fig. 22). The maximum thrust measured at CL % 1.4 is C, = -0.27, and it exceeds the total momentum input of C, % 22.6%. According to these results, blowing at high levels of C, is much superior to suction. It is well known that there is a drag penalty associated with the removal of boundary layer flow through the airfoil’s surface. According to Poisson-Quinton4this drag is equal to CDsink= 2 c Q = 0.0312 and, were it not for this penalty suction, would have generated thrust as well. When massive blowing is used, the jet momentum is recovered as thrust, but in addition there is a source flow that should contribute to thrust. In this case the thrust attributed to the source CD,,,,, = - 2 c Q % 0.034. The maximum thrust recovered may be CT = C, 2 c Q % 0.26. Because the energy spent to generate suction is similar to that of blowing, the superiority of steady blowing is clear in this case. The drag associated with periodic excitation is larger than either steady blowing or suction, but because the (C,) spent in this case is but a small fraction in comparison to the steady cases, the comparison is inappropriate. Repeating the same experiment at Re = 2.35 x lo5 reduces the respective C values by a factor of 4. In the case of periodic excitation, it also affected F+Y lowering it to F+ = 0.35. The baseline CL-a curve is now much more normal, particularly for a > 10 deg (Fig. 23), reaching a CL,, of 1.45. The CLma,generated by steady blowing dropped from CLmax= 5 to 3.1, whereas suction attained a CLmaXof 2.5 only. The efficacy of periodic excitation was further reduced as a result of the concomitant change in F+ yielding a CLm, of 2.25, in spite of the fact that for a < 14 deg it generates a higher lift than the steady suction does at a C, that is an order of magnitude larger. The drag polar of the basic airfoil suggests that the flow may be attached to the TE ramp at a = 18 deg corresponding to CL = 1.2 (Fig. 23). For this C,, the L I D of the basic airfoil is 15. Steady suction at C, = 4.7% generated identical total drag at CL = 1.2. There is a major difference in the form drag of these two cases. Whereas the CDpgenerated by the suction is so small and perhaps even negative, the CDpassociated with the basic airfoil is larger than the total drag (CDp= 0.113, whereas C , = 0.085). Active flow control (AFC) generates a higher total drag at this C, (C, = 0.1 l), whereas steady blowing reduces the C, to 0.015. In fact, at CL = 0.8 the drag associated with the steady blowing at C, = 5.5% vanishes, implying that a wake generated by this C, resembles a wake of a self-propelled, two-dimensional body. The actual C, required to propel an aircraft at this CL is higher because of the added induced drag. In contrast to the results accumulated on the NACA 0015 and its truncated derivative, CDpis usually smaller than C , except near the stall angle; it is possible that the concave curvature affects the result as well. The pressure distributionsover the airfoil at C, = 1 and CL = 2.1 and Re values of 1.17 x lo5 and 2.35 x lo5 are plotted in Fig. 24 for the three control mechanisms used. Because the same CL was obtained at different incidence angles as a result of differences in C, and the specific method of control used, a comparison of pressures near the LE is inappropriate. Nevertheless, the efficacy of the control scheme reflects on the incidence angle for which a given CL is achieved. Consider the case for C, = 2.1 and Re = 1.17 x lo5 (Fig. 24b). It appears that the flow is accelerating toward the slot for all three of the control schemes used. Suction, however, provides the least maximum acceleration, but its effect is felt

+

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Uinf=I4mls,Re=2.35*1 O5

Fig. 23 The variation of C, with a and the corresponding drag polars at Re = 2.35 X lo5 and the various control parameters used.

farther upstream from x / c = 0.5 up to the slot located at x / c = 0.62. At the slot itself, suction brought the flow to stagnation (C, x l), whereas on the reminder of the ramp (0.65 < x / c < l), an almost constant, slightly negative base pressure is maintained (e.g., C= ., -0.2). Active flow control accelerates the flow upstream of the slot, but it also maintains a good pressure recovery downstream, culminating with C, = +0.2 near the TE. Strong blowing (at C, that is an order of magnitude larger than the AFC applied) provides a favorable pressure gradient just upstream of the slot and a very-low-pressure bubble immediately downstream. The bubble is generated by the oblique injection angle of the jet,

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P 0

rn

n

0

Fig. 24 Pressure distributions measured at CL= 1 and 2.1 for two values of Re and various control mechanisms.

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139

because the slot is inclined at 30 deg to the downstream surface. After reattachment the C p is still negative because the curvature of the surface is convex; however, at x/c % 0.68 the curvature of the surface changes sign and also the sign of the Cp, which becomes positive downstream of this chordwise location. The large positive pressure on the TE ramp contributes to the thrust generated by the steady blowing (Fig. 22). The measured C p can be larger than unity because the high-speed jet that emanates from the slot has a larger total pressure than the freestream (Cp,,, = 5). To attain CL = 2.1 at Re = 2.35 x lo5 requires a much larger incidence than for the lower Re, partly because of the lower C, available. In this case the same C, is obtained for AFC and for steady suction (in spite of the large difference in the level of actuation) at a = 18 deg, enabling a direct comparison between the pressure distributions upstream and downstream of the slot. It is clear (Fig. 24d) that steady suction is more effective in accelerating the flow upstream of the slot than AFC is; however, the latter is more effective in the pressure recovery region on the ramp. The maximum C p due to steady blowing measured around x/c = 0.7 was reduced from C p = +5 to C p = +1.25, because the total pressure of the freestream was increased by a factor of 4. It suggests that the regularly defined pressure coefficient so widely used on normal airfoils is not adequate in the case of blowing. The LE edge radius of curvature of the GLAS I1 airfoil is very small, pointing to a potential problem with this design. There is a large separation bubble on the lower surface of the airfoil that reduces greatly the favorable pressure gradient on this surface. This is most obvious when the flow over the ramp is separated (Figs. 24b, 24c, and 24d). Laminar flow on this airfoil is probably achieved on the upper surface around CL % 1 by a proper combination of incidence, BLC and Re. The application of AFC at low values of C, < 2% is considered in Fig. 25 with particular emphasis being placed on C, < 1% at incidence a = 0 and 6 deg. At a = 0 deg there is an increase in lift of ACL = 0.4 at C, < 0.4%, whereas at a = 6 deg the same increment in lift requires 0.4 < C? < 0.9% depending on the reduced frequency F+ used. The largest increase in CL for the smallest input in C, corresponds to a reduced frequency of 0.7 < F+ < 1. For F+ = 1.8 the sudden increase in CL requires a lower input of momentum, but the ACL is somewhat smaller. The pressure distributions taken in the region of transition from the “low to high” CL suggest that even a partial attachment of the flow downstream of the slot changes the circulation affecting the entire pressure distribution on the upper surface of the airfoil. Active flow control enables the boundary layer to overcome a very severe adverse pressure gradient existing on the convex part of the ramp (i.e., 0.6 < x/c < 0.7), even if it does not succeed in attaching the flow all the way to the TE (Figs. 26 and 27). The data presented in Fig. 26 correspond to C, = 0.25 and 0.75% for a = 0 and 6 deg, respectively. It is noted that the LE stagnation point occurs on the upper surface around x/c % 0.01 for a = 0 deg, and almost precisely at the LE for a = 6 deg. The existence of a bubble near the LE of the lower surface is observed. The acceleration of the flow over the leading 50% of the chord on the upper surface is followed by a smooth deceleration towards the slot and towards the inflection point on the surface of the TE ramp, provided the C , is close to the threshold value at

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Effect of F’ and C, on C, at GO*,

Ui,pl4ml~

Effect of F’ and C, on C, at a=$,Ui,,=14mls

Fig. 25 Increase of C, with C, for a variety of F’ at two values of a.

which CL increases (Fig. 26). A further increase in C, results in a “spike” in the observed C pjust upstream of the slot and in a higher pressure recovery over the ramp (Fig. 27). These results seem to contradict the previous concept that required a threshold value of CPcritto overcome separation before circulation can be increased. In this case the control of separation (downstream effect) and the increase in circulation (upstream effect) occurred simultaneously. The effects of AFC on drag are shown in Fig. 28 for a = 6 deg, Re = 235 x lo5 and for three values of F+ ranging from 0.36 to 1.8. For F+ > 1, the drag is reduced to approximately one-third of its value in the absence of excitation, whereas for F+ = 0.36, the reduction is only two-thirds. Although the high-frequency periodic excitation reduced the total drag at all

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C,-0.26%, GOO, U,,,+4mls

Fig. 26 Effects of F + on the pressure distributionfor C, = 0.25% and 0.75% at two values of a.

amplitudes (C, > 0), the excitation at F+ = 0.36 increases the total drag for 0 < C, < 0.6%. The highest frequency of excitation experimented with to date, F+ = 1.8, reduced the drag at the lowest amplitudes imposed. The response of CDpto the imposed excitation is quite different. First, CDpwas not increased by excitation at F+ = 0.36; secondly, this frequency was able to lower the CDpas effectively as F+ = 1.8; and thirdly, higher CDpresulted from excitation at F+ % 1. The present results are in general agreement with the parametric

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F'=0.36,a=0°, Uim=14mls

F'=0.36,a=6', U i e l4mls

Fig. 27 Effects of C, on the pressure distribution for two values of F + and a.

study of Nishri and Wygnanskilg on the reattachment of flow to a generic flap by using AFC. The high CDpassociated with excitation at F+ x 1 results possibly from the strongest eddies that are consistently present over the ramp and that are generated by this frequency. Excitation at lower frequencies results in stronger eddies that are not always present over the surface, because their wavelength exceeds the length of the surface, whereas excitation at F+ >> 1 generates weaker eddies, so their constant presence has a lesser effect on the flow. Nishri and colleagues also observed that to maintain the flow attached requires a much smaller momentum input than to force a separated flow to reattach. In other words, a hysteresis should be present whenever the level of actuation

CIRCULATION AND SEPARATION CONTROL EXPERIMENTS a=6', Pand C, effect on CD, CD,,

143

at Ui,,,=14mls

Fig. 28 Effects of C, on the drag for three values of F + at a = 6 deg.

is increased until the flow reattaches or when the fluctuation level is decreased until the flow separates. Such a hysteresis was observed and is plotted in Fig. 29. A comparison among all three methods of control discussed is shown in Fig. 30 for a prescribed incidence of 6 deg. Active flow control results in a sudden increase in ACL % 0.6 when (C,) % 0.5%; suction requires a C, = 1.5% to obtain the same result, and blowing requires a higher threshold. Seifert et al. made similar observations on the NACA 0015 airfoil.24 At larger values of C,, blowing may surpass AFC, but the practical implications of this are questionable in view of the large momentum required. a=6*,FreqrSSOHz, F'11.8, U,,,=14mls

Fig. 29 Hysteresis effect caused by an increase or a decrease in C,.

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144

a=6', C, sweep at U,,,=l4m/s CL

Fig. 30 Effects of suction, blowing, and AFC on lift at small values of momentum input.

D. Flow Around an Elliptical Airfoil The elliptical cylinder represents an aerodynamic body that has no definite Kutta condition at its trailing edge, but it generates pressure distributions that are similar to an airfoil at an angle of attack. The modified 30% ellipse has both leading and trailing edge cylinders that have adjustable slot widths and exhaust locations (measured as included angles either from the LE or TE; (Fig. 3). The focus in the present paper is only on fluidic actuation emanating from the TE cylinder. Because of the undefined Kutta condition, the elliptical airfoil has the adjustable circulation characteristics of a cylinder when flow control is introduced. Steady suction, steady blowing, and oscillatory excitation have been tested using the model. The effectiveness of the three flow control categories at a = 0 deg has been quantified (Fig. 31). The slot location c$ was varied around the TE of the model in order to generate the plotted data. The three flow types were run at a blowing coefficient of 1.9%. The oscillatory flow produced the best lift results, followed by suction and then blowing. The slot was inclined at 25 deg to the local surface to enhance mixing with the boundary layer. This may provide the reason for the poor performance of the blowing relative to the other control methods. If the exhaust angle had been tangential, blowing might have extended the attached flow region farther around the TE and increased the lift. The angle where the applied flow control provided the best lift moved toward the TE until the control was insufficient to counter the adverse pressure gradient upstream of the slot. As the momentum of the applied flow control was increased, the location maximizing the lift moved towards the TE, regardless of the method of control used. Only C, values generated by AFC are shown in Fig. 31, indicating that an increase in (C,) from 1.9 to 2.38% increases both C, and by at least 5 deg, to 130 deg.

,,+,

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/>

0.41

Suction (C@= 1.go/,)

-@-

( f = 130Hz, = 1.9p)

+AFC w.w-

145

I

I

100

I

110

io

120

1

140

TE slot exhaust angle (")

Fig. 31 Comparison among three different types of flow control on a 30% thick ellipse.

The pressure distributions at C, x 0.44 give some indication where the lift is generated for the three different control types (Fig. 32). The pressure distribution over the elliptical body is nearly constant for the three cases from 15 to 85% chord. The real difference is at the leading and trailing edges. Active flow

CP

-2.0

AFC CL = 0.448 = 1.9% TE = 110°/0.030" Blowing CL = 0.436 Cµ = 1.9% TE = 110°/0.030" Suction CL = 0.441 Cµ = 1.9% TE = 120°/0.030" 110° Slot Location

-1.5 -1.0 -0.5

120° Slot Location 0.0 0.0 0.5

0.2

0.4

0.6

0.8

1.0

x /c

1.0

Fig. 32 Comparison among different flow control techniques at constant C, 30% thick ellipse.

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D. CERCHIE ET AL.

control and steady suction produced the same increase in Cp at the LE when the control was applied at the TE. These two control approaches also produced the same pressure distribution along the lower surface near the TE. Active flow control and blowing looked similar along the upper surface near the TE where the higher velocity slot flows were entering the flow field. The lowpressure peak associated with the slot flow was not present at the TE for the suction case, as was expected. As a test note, the AFC data plotted represent the time-averaged pressure data. The data were gathered at 600 Hz, and 700 samples were gathered for each static port. Based on the number of static ports on the model, the typical data runs were just over a minute in duration. The effect of (C,) on Cp is plotted in Fig. 33. The AFC “off’ (baseline) condition is also plotted for reference. This baseline pressure distribution agrees well with the CFD results generated using NASA’s CFL3D program. The three AFC magnitudes increased the lift by increasing the velocity along the entire upper surface of the ellipse. The lower surface velocity is hardly affected by AFC. The application of AFC near the TE increased the circulation along the entire span of the airfoil, not just locally at the TE. A closer look at the Cp at the trailing edge region for various AFC magnitudes is presented in Fig. 34. The baseline condition shows symmetrically separated flow on the upper and lower surfaces from around 97% chord to the TE. Application of AFC resulted in a large, time-averaged increase in the velocity on the upper surface. The separated region shrank on the upper surface, but it was hardly affected on the lower one, until, for (C,) = 2.38%, the flow over the entire upper surface is attached.

Fig. 33 Supercirculation on a 30% thick ellipse using AFC (CFD courtesy of A. Hassan, Boeing).

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cp-1 .o

130" Slot Location

ff=

147

0"

+-----

\

-0.6

I -0A -02 0.o

02 C

-m-

-dP>

AFC C;=0.619 = 2.38% Baseline CFL3D Re =340K Fully Turbulent

-

Fig. 34 Zoom into TE region of Fig. 33.

The effectiveness of AFC applied to the TE element of the ellipse as a function of a is plotted in Fig. 35. The data show a nearly constant benefit of AFC until a nears stall. The increase in C, is approximately 40%. Angle of attack sweeps were performed with steady suction being applied from the TE cylindrical cross-section (Fig. 36). Rather than increasing the blowing coefficient, the slot angle was altered for each of these sweeps. The intent was to assess the influence of a on the optimal slot location and compare these data to the results shown in Fig. 31 at a = 0 deg. The initial run was at C$ = 120 deg, or 5 deg from the peak performance point shown on Fig. 31. The lift benefit ACL from steady suction was nearly constant for a < 4 deg. The slot angle C$ was then reduced by 10 deg, but a,,, (i.e., a corresponding to CL,,,) increased only by 3 deg. Another 10 deg reduction in C$ only increased a,,, by 1 deg. From these data it is clear that the zero angle of attack data provide a good insight into the optimum performance location. They also shows that the upstream pressure gradient has a very big influence on the performance of control applied at the TE region. In fact, the first decrease in lift (stall) occurs for the flow separating from the TE cylinder; thereafter the lift increases again with increasing a until the flow separates from the LE of the ellipse (Fig. 36). The same test was repeated for a larger slot angle and larger suction coefficient (Fig. 37). The trends seen at the smaller slot-width were validated. The larger .,,,C , However, the a at which CLm,, was suction coefficient produced larger realized was not directly proportional to the change in the suction slot location

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CL

1.6 F* = 0.35,TE = 130"/0.030", U = 1Oms-l

1

+Baseline A =l.2%

t=l.8%

1.2

0.8

0.4

0.0

1

0

5

10

15

20

a (") Fig. 35 Angle of attack sweep using AFC.

or its width. Figure 38 shows the effect of steady suction on the C, at a = 0 deg where the basic ellipse provides no lift. The lower surface velocity is decreased only slightly by the suction and almost at the same increment along the entire lower surface. The strongest effect occurs on the upper surface where the increase in velocity is constant from the LE to 90% of the chord. The change in the TE

CL

1

--W-

.2 a = 0",h = 0.030",Cp= 1.9%

Baseline

+120" 110" 100"

0.8

0.4

0.0 .d/

0

I

I

I

5

10

15

Angle of Attack a (")

Fig. 36 Suction on 30% ellipse: C, = 1.9% and h = 0.030 in., U = 10 d s .

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Angle of Attack a (")

Fig. 37 Suction on 30% ellipse at C, = 3.5%.

static pressure is very evident on this plot and it is similar to the AFC data shown in Fig. 33. Comparing the pressure distribution at C, = 0.8 with and without the suction necessitates a comparison between a = 4 deg and a = 10 deg (Fig. 37), which are very differently loaded along the chord and generate different moment around the aerodynamic center. Active flow control tests at the TE were then conducted for a variety of slot angles, excitation frequencies, and amplitudes.

Fig. 38 Pressure distribution at C, = 3.5%.

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ACL

Fig. 39 Scatter in data when C, used as parameter.

Data acquired at two slot locations are plotted in Fig. 39 against (CJ. The data did not collapse onto a single curve, indicating that some other parameter or parameters need to be included for this type of control and geometry. There are two broad clusters of points depending on the slot location. There is also a dependence on frequency with the lowest frequency of excitation generating the lowest ACL. Empirical correlation trials resulted in the data plotted in Fig. 40. This set of data appears to collapse all the results onto a single curve as a function of the product of (C,), the square root of F+, and the angle denoting the distance from the TE. The length scale used in the definition of F+ depends on the same angle, as it represents the length from the slot to the theoretical TE of the ellipse. In this case the data collapse onto a single curve whose generality is yet to be proven.

E. Controlled Flow Around a Circular Cylinder and Reexamination of Some Old Results The flow around the circular cylinder is discussed last because it represents the highest degree of complexity. Similarly to the ellipse it has no defined TE separation location or imposed Kutta condition. Unlike the previous four geometries discussed, the LE and the TE flows are closely coupled because of their immediate proximity. The flow is also sensitive to transition location, which is, in turn, sensitive to a variety of inflow parameters.

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Fig. 40 Collapse of data based on empirically derived flow control parameter when AFC is used.

In the absence of an external stream, a wall jet created by steady blowing wraps itself around a convex surface of the cylinder, following it up to, and sometimes beyond half of its circumference (Fig. 1). In the example shown, a jet of momentum J emanating to the right from a slot located on top of the cylinder encircles it before separating to the left from its lower surface. The change in the direction of the flow generates a low-pressure region on the right-hand surface, which, when integrated, yields a side force whose magnitude is almost equal to twice the jet momentum. This force multiplier makes some applications of wall jets over curved surfaces very attractive, arousing interest in improving the understanding of this flow. One of the unique characteristics of the curved wall jet is its phenomenal rate of growth from the surface, and its high turbulence level, which is attributed to the streamwise vortices generated by a centrifugal in~tability.'~ The cylinder over which these measurements were made was carefully designed with the jet emerging tangentially to the surface after passing through a smooth contraction. The resulting flowfield is a consequence of instability. Therefore, it is important to know how sensitive this flow is to the detailed jet characteristics leaving the nozzle and to the initial width of the jet relative to the radius of the cylinder (Fig. 41). There is a large region around the surface of the cylinder where the pressure is constant in spite of the rapidly thickening jet flow. The pressure distribution is not seen to be very sensitive to the jet-related Reynolds number.

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Fig. 41 Pressure distributions on a cylinder for various Reynolds numbers.

Two additional cylinders were constructed that were more suitable for wind-tunnel tests in which an external stream of variable velocity could be applied. One of the cylinders (2 in Fig. 42) was machined from two parts and has a continuous nozzle along its span. The figure shows that the nozzle design is a critical feature in determining the external flow characteristics. The best tangential design, nozzle 1, maintained attached flow at least 40-deg further around the cylinder than the other two slot designs. Nozzle 3, which had a segmented slot cut through the cylinder’s wall, demonstrated the worst performance. This poor tangential flow characteristic was previously discussed as a possible reason for the lower blowing control performance near the TE of the ellipse relative to suction and to AFC control. Using slot design 2, the pressure distributions around the cylinder, as a function of slot location relative to the freestream, were generated for a given-jet-to freestream velocity ratio of 14.5 (Fig. 43). The data possess the same trend as for the ellipse, where the maximum surface velocity continues to increase until the slot has rotated far enough from the natural baseline separation point (around 60-70 deg for that Reynolds number) that the added momentum can no longer hold the flow attached. The maximum Cp generated is -22 for an injection angle 120 deg measured from the LE of the cylinder. It is interesting to note that the region of separated flow on both sides of the cylinder is reduced as the performance of the cylinder is increased. For the smallest angle of 30 deg, the separated region extends from approximately 140 to 290 deg, whereas of extends only from 240 to 290 deg for the performance slot angle of 130 deg, which already exceeds the most effective location of 120 deg. A comparison of the pressure distributions for the three flow control approaches over a single slotted cylinder is plotted in Fig. 44. The slot location was rotated from the LE toward the TE and the data plotted

+=

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Comparison of Pressure Distribution of C o d a Flow ReM=llOOO

Fig. 42 Pressure distributions on a cylinder for various slot geometries.

correspond to the best-performance location. The data are presented in a manner similar to airfoil data in percent of chord, using the cylinder’s diameter as the chord, rather than degrees around the cylinder. The two blowing cases illustrate a standard observation: the stronger the blowing, the farther Cp distribution for different slot locations Re=26000, U,/Um=14.5, b/R=O.Oll

Fig. 43 Pressure distributions on a cylinder for various slot locations.

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D. CERCHIE ET AL. Comparison for three Cases

Fig. 44 Pressure distributions on a cylinder for three different flow control approaches.

downstream into the adverse pressure region the jet can be introduced for maximum performance. Note also that the suction and oscillatory (AFC) flow control have nearly the same pressure distribution from the LE to the slot, while the AFC input magnitude is less than one-third the suction magnitude. The AFC and steady blowing pressure distributions differ from the steady suction case aft of the slot location. The steady suction pressure distribution has a sharp increase in pressure whereas the steady blowing and AFC possess a more gradual change in pressure toward the TE. Judging from the pressure distribution, the flow is fully attached for the steady blowing at C, = 39%, but it is separated downstream of the slot when comparable suction is used. Active flow control manages to attach the flow over 95% of the chord at C, = 9%. The integrated force coefficient CF was measured and plotted in Fig. 45a as a function of slot angle relative to the LE for different slot velocities and heights. In general, a higher C, generates a higher force. An attempt to normalize this data with respect to the jet momentum (as was done in the absence of an external stream) is plotted in Fig. 45b. The data nearly collapse to a single curve for small jet injection angles relative to the ideal LE, but the scatter is still large downstream of the slot location corresponding to .,,,C , The pressure distributions along the cylinder for three different slot velocities are plotted in Fig. 46. In the first case the pressure distribution is normalized by the freestream dynamic pressure q, whereas in the second it is normalized by the cylinder diameter and by the jet momentum J . Again, the trend is a proportional relationship between slot velocity and the freestream, but neither q nor J normalize the pressure distributions correctly. In the first case the higher the ratio between the jet and the freestream velocity, the higher the negative C ,, while

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155

a)

Fig. 45 Total force generated on cylinder for different blowing conditions: a) C,, b) CFIC,.u

in the second the roles are reversed. Some other parameters are required for better correlation of the data, although self-similarity in C, may not be possible for the entire range of velocity ratios. A possible parameter would include C, and a correction to that term that would be flow configuration dependent. One such approach was

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D. CERCHIE ET AL.

Fig. 46 Pressure distributions on a cylinder for different blowing velocities and normalization.

proposed by Englar26:

where Us is the local velocity measured just upstream of the slot. This local velocity is not always available from past tests and should scale with freestream velocity. Therefore, using a simplified assumption, the equation can be written in

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157

the following form for discussion purposes:

In fact, when a jet of kinematic momentum J and volumetric flow q is injected parallel to an infinite stream of velocity U,, in a way that retains constant pressure on the surfaces bounding a control volume, the application of the momentum theorem yields an invariant

which is proportional to Englar’s simplified version of C , . The two derivations lead to the same result when J is calibrated in the absence of an external stream, which is normally the case. The last equation is applied to some early data in boundary layer flow and circulation control to determine its applicability. The blown flap data from Poisson-Quinton and Lepage2 is adjusted using the modified flow coefficient and the results are replotted in Fig. 47 together with the original results that use just C,. Similar reprocessing was carried out to the win data of Williams21 (Fig. 48) and to the NACA 84-M airfoil data of Attinello (Fig. 49). In all these cases the data scatter is no greater using the modified parameter, and in some of the more critical, low-momentum blowing tests the inclusion of CQ seems to fit the data better. Large C, values in testing are typically the result of large slot velocities. For these conditions, the impact of the correction term diminishes, allowing for reasonable performance predictions to be made for a given configuration. In an effort to minimize the momentum input required for effective flow control, this modified parameter provides a better insight into the measured data for smaller flow coefficients. In Fig. 49 the data collapse better using the modified parameter for the small corrected flow coefficient conditions, even matching the measured data while it shows a negative lift increment when the flow coefficient is negative. However, it is not suggested that these data justify a new term for design purposes. What this does show is that in order to optimize flow control using a minimum input, a flow-related correction parameter would have to be added to C,. It is unknown whether a correction factor can be introduced that would collapse all flow geometries and control strategies to a unique line without the introduction of terms that account for the pressure gradient and boundary layer characteristics in the region of the slot. An interesting characteristic of flow control is shown in Fig. 50. The cylinder exhibits a lift hysteresis around its maximum lift condition due to changes in slot

9

158

D. CERCHIE ET AL. CL vs. Cp or (0.5*CpCQ) (from Polsson-Qulnlon& Lepage, 1961)

Fig. 47 Blowing over a wing section using data from Ref. 2 (Poisson-Quintonand Lepage). CL vs. Cp or (0.5*Cp-CQ)

Fig. 48 Blowing over a wing section using data from Ref. 2 (Williams).

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159

CL vs. Cp or (0.5*CpCQ) (from John Attinello, 1961)

Fig. 49 Blowing over a wing section using data from Ref. 2 (Attinello).

location, under steady blowing. This phenomenon is similar to an airfoil when the Therefore, the initial flow condition is angle of incidence is altered around astall. also important to the performance of BLC and CC, as was discussed by Nishri and Wygnan~ki.'~ Previous tests on a Wortmann airfoil27 have shown that this hysteresis can be reduced or eliminated by using AFC. The pressure distribution around the cylinder with a slot located at 120 deg from the LE indicates the flow can be either attached, yielding a maximum lift, or completely detached, providing no lift at all (Fig. 50b). The latter pressure distribution is very close to the baseline data. The significance of a single slot location to the performance of the prescribed control mechanism has been amply demonstrated, but very few attempts have been made to add another slot with another array of actuators or jets. A second slot was thus added to cylinder 3 (Fig. 42) to determine its possible impact. Figure 51 shows the dual-slot configuration, where both slots are on the same side of the cylinder and they inject flow in the same direction. The performance of this configuration relative to the single-slot performance for the same blowing coefficient is compared. The dual-slot arrangement provides 30% more lift relative to the single slot by improving the downstream flow conditions. It is interesting to note that the second slot provides only a limited control of separation downstream, but it alters mostly the pressure distribution upstream, contributing to supercirculation. This emphasizes the significance of the location of the fluidic actuator and suggests that a distributed actuation may be more effective, as the designers of the NOTAR helicopter demonstrated.

160 a)

Pressure Comparison, Cp=O,57, or=190°

Fig. 50 Lift and pressure hysteresis for changing blowing location on a cylinder.

Control of lift is not the sole purpose of BLC, because drag reduction for cruise might be as important. By blowing from two symmetrical slots located at a = & 110 deg away from the lead stagnation point, the pressure or velocity over the cylinder can be changed without introducing lift (Fig. 52). The inviscid value of the pressure coefficient of - 3 is realized at 50% of the chord position. The cylinder’s drag drops from approximately CD > 1.0 for the natural case (not blown) case to a CD x 0.4 for the symmetrically blown case. Proper accounting should be made in order not to waste momentum, because too much flow control may be applied for given conditions. Figure 53 shows that for symmetric blowing on the cylinder at C, = 0.46, the total drag reduction is 0.61 and all of the injected flow is recovered as thrust. However, when C, is increased to 1.02,

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Comparison of Pressure Distribution slot located at u=9O0 Cylinder # 3, Re=39000, Cp=1.31

Fig. 51 Pressure distributions on a single- and dual-slot cylinder.

the total drag reduction is only 0.69 and some of the injected flow is wasted. The latter case may provide other benefits such as enhanced stability, but the performance suffers. It is also interesting to note the increase in the wake's vortex shedding frequency, which is associated with the reduction in drag (Fig. 53) and is equivalent to the momentum thickness in the wake. This is Double slotted Cylinder 1loo, 0' sweep back, Re135000, Steady Suction, C, per slot

Fig. 52 Pressure distributions on a dual-slotted cylinder for different suction C,.

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C,

0.00 0.46 1.02

C D ~1.02 0.26 0.18 Co

1.04 0.43

0.35

0.00 -0.14 0.32

Fig. 53 Cp distribution and wake profiles associated with a dual-slotted cylinder.

consistent with observations suggesting that a dominant Strouhal number associated with vortex shedding is constant, requiring a frequency increase resulting from a decrease in the momentum thickness. 111. Conclusions

Flow control tests over five two-dimensional, aerodynamic, and bluff bodies provided some new insight into the parameters governing active control of separation and circulation. Suction, blowing, and periodic forcing enable one to tailor the pressure distribution over the airfoil surface in a similar fashion to lofting or the introduction of passive devices such as flaps or control surfaces. This chapter questions some of the accepted concepts associated with separation and circulation control, although it is not able to provide definitive answers to these questions. For example, the enhancement of circulation can be achieved without attaching the flow; however, by forcing separated flow to reattach to the surface, circulation is generally enhanced. Flow reattachment and the control of circulation do not have to occur sequentially, as it is widely believed, leading to some critical value of an input parameter that separates the two.

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The use of steady blowing is ineffective at low C,, but its ability to contribute to thrust while increasing the circulation make it attractive at C , > 5%. Because 0.05 represents a typical ratio of thrust to lift on a civilian airplane in cruise, steady blowing might become the technique of choice whenever the integration of propulsion and aerodynamics are considered. Suction and periodic excitation are much more effective in reattaching the flow at low C , values and enhancing the lift, but the input momentum is usually not recovered. Steady suction, in particular, contributes to drag. The low levels of C , required to attach the flow by periodic excitation make it the most attractive technique for lift enhancement. The momentum coefficient that is widely used to correlate the data is not as universal a quantity as it is believed to be. It is mostly deficient at low-level inputs that characterize AFC, requiring consideration of mass addition and perhaps other variables (e.g., changes in the Kutta condition and pressure gradient at the slot). A corrected flow coefficient term was discussed and applied successfully to some well-known flow control cases. The modified control parameter may be especially useful for correlating data using lowintensity, pulsed blowing or suction. A better understanding of the performance of slot design, its width, and velocity for various geometries is required. The drag of a two-dimensional bluff body is dominated by pressure drag because the skin friction contribution to drag is small in comparison. However, in the presence of fluidic control the reliance on CDpas being the major contributor to drag is wrong. In most cases, as C , increases, so does CDpwhile the actual drag decreases. This is most obvious when the TE is blunt and the flow is attached due to fluidic control. One wonders how to normalize the pressure distribution around an airfoil when C , is of order 1 because then qc % J . The controlled flows over the ellipse and the circular cylinder have many features in common. On the ellipse, however, one may change the pressure gradient upstream of the slot by changing the location of the slot relative to the center of the ellipse or by changing the thickness ratio of the ellipse. The TE performance becomes less predictable as the pressure gradient along the forward section of the elliptical airfoil is increased. An empirical parameter was introduced to scale these results and make a rational comparison with other geometries that are creating large circulation at the TE. The circular cylinder data demonstrated the importance of a good slot design as well as the superior efficiency of oscillatory flow control relative to steady blowing or suction. It also showed that there is a limit to blowing that should be observed if the overall system performance is to be objective. The curvature of the surface downstream of a slot is an important variable, as was demonstrated on the truncated NACA and GLAS I1 airfoils. A concave surface generates a positive pressure in the TE region that reduces the drag; however, it is susceptible to centrifugal instability and may generate streamwise vortices whose effect on separation and CC is not well understood. Periodic excitation emanating from a two-dimensional slot utilizes the inherent convective instability of the flow that is either separated or is on the verge of separation. Curvature renders another mode of instability that can be used to delay separation. It may enable the replacement of slot blowing by compact individual nozzles whose spanwise separation triggers the centrifugal instability associated with curvature.

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Acknowledgments This work was supported in part by a grant from ONR that was monitored by R. Joslin. The authors are indebted to A. Hassan for his help in providing the CFD results that guided some of the experiments on the ellipse and the GLAS I1 airfoil. The authors wish to acknowledge H. Nagib for providing them with an important figure and for many helpful discussions related to AFC.

References ‘Prandtl, L., “The Generation of Vortices in Fluid of Small Viscosity,” Journal of the Royal Aeronautical Society, Vol. 31, 1927, p. 735. 2Lachmann, G. V. (ed.), Boundary Layer and Flow Control: Its Principles and Application, Pergamon Press, New York, 1961. 3Goldschmied, F. R., “Integrated Hull Design, Boundary Layer Control and Propulsion of Submerged Bodies,” Journal of Hydronautics, Vol. 1, 1967 p. 2. 4Poisson-Quinton, Ph., “Recherches Theoriques et Experimentales sur le Controle de Couche Limites,” VII International Congress of Applied Mechanics, 1948. ’Stratford, B. S., “Early Thoughts on the Jet Flap,” Aeronautical Quarterly, Vol. VII, 1956, p. 45. 6Helmbold, H. B., “The Lift of a Blowing Wing in a Parallel Stream,” Journal of the Aeronautical Sciences, Vol. 22, 1955, p. 341. 7Spence, D. A., “The Lift Coefficient of a Jet-Flapped Wing,” Proceedings of Royal Society Series A, Vol. 238, 1956, p. 46. ‘Legendre, R., Influence de 1’Emissiond’un Jet au bord de Fuite d’un Prof1 sur 1’Ecoulement autour de ce Profil, Comptes Rendus, AcadBmie des Sciences, Paris, 1956. ’Woods, L. C., “Some Contributions to Jet-Flap Theory and to the Theory of Source Flow from Aerofoils,” A.R.C. Current Paper 388, 1958. “Malavard, L., Sur une Thkorie Lineaire du Souflage au bord de Fuite d’un Profil d’Aile, Comptes Rendus, AcadBmie des Sciences, Paris, 1956. “Wygnanski, I., “The Effect of Jet Entrainment on Loss of Thrust on a TwoDimensional Jet-Flap Aerofoil,” Aeronautical Quarterly, Vol. 17, 1966, pp. 31 -51. ‘2Wygnanski, I., and Newman, B. G., “The Effect of Jet Entrainment on Lift and Moment for a Thin Airfoil with Blowing,” Aeronautical Quarterly, Vol. XV, 1964, p. 122. 13Hynes,C. S., “The Lift, Stalling and Wake Characteristics of a Jet Flapped Airfoil in a Two Dimensional Channel,” Stanford Univ., Stanford, CA, SUDAAR No. 363, 1968. 14Stratford, B. S., “An Experimental Flow with Zero Skin Friction Throughout its Region of Pressure Rise,” Journal of Fluid Mechanics, Vol. 5, 1959, pp. 17-35. ‘’Elsberry, K., Loeffler, J., Zhou, M. D., and Wygnanski, I., “An Experimental Study of a Boundary Layer that is Maintained on the Verge of Separation,” Journal of Fluid Mechanics, Vol. 423, 2000, pp. 227-261. ‘%eifert, A., and Pack, L. G., “Active Separation Control on Wall Mounted Hump at High Reynolds Numbers,” A I M Journal, Vol. 40, No. 7, 2002, pp. 1363- 1372. 17Greenblatt, D., Paschal, K. B., Yao, S. C., Harris, J., Schaeffler, N. W., and Washburn, A. E., “A Separation Control CFD Validation Test Case,” AIAA Paper 2004-2220, June 2004. 18Greenblatt, D., and Wygnanski, I., “The Control of Flow Separation by Periodic Excitation,” Progress in Aerospace Sciences, Vol. 36, 2001, pp. 487-545.

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‘’Nishri, B., and Wygnanski, I., “Effects of Periodic Excitation on Turbulent Flow Separation from a Flap,” AIAA Journal, Vol. 36, No. 4, 1998, pp. 547-556. ”Darabi, A., and Wygnanski, I., “Active Management of Naturally Separated Flow Over a Solid Surface, Part 11: The Separation Process,” Journal of Fluid Mechanics, Vol. 510, 2004, pp. 131- 144. ”Williams, J. “British Research on Boundary Layer Control for High Lift,” Boundary Layer and Flow Control: Its Principles and Applications, edited by G. V. Lachmann, Pergamon Press, New York, 1961. ”Glauert, M. B., Walker, W. S., Raymer, W. G., and Gregory, N., “Wind Tunnel Tests on Thick Suction Airfoil with a Single Slot,” Aeronautical Research Council R & M, No. 2646, Oct. 1948. 23Salter,C., Miles, C. J. W., and Owen, R., “Tests on GLAS I1 Wing Without Suction in the Compressed Air Wind Tunnel,” Aeronautical Research Council R & M No. 25