Low Speed Airfoil Data V5 SELIG

Summary of Low-Speed Airfoil DataGregory A. Williamson, Bryan D. McGranahan, Benjamin A. Broughton, Robert W. Deters,John B. Brandt, and Michael S. Selig

Volume 5

Summary of Low-Speed Airfoil Data

Volume 5

Gregory A. WilliamsonBryan D. McGranahanBenjamin A. BroughtonRobert W. DetersJohn B. BrandtMichael S. Selig

Department of Aerospace EngineeringUniversity of Illinois at Urbana-Champaign


Volume 5

Copyright c© 2012 byGregory A. Williamson, Bryan D. McGranahan, Benjamin A. Broughton, Robert W. Deters,John B. Brandt, and Michael S. SeligAll rights reserved.

On the cover: Supra 134-in (3.4 m) span sailplane design by Prof. Mark Drela (MIT)with his AG40d airfoil used at the root section and tested in this volume.

Williamson, Gregory AlanSummary of Low-Speed Airfoil Data – Volume 5 / by Gregory A. Williamson, Bryan D. McGranahan,Benjamin A. Broughton, Robert W. Deters, John B. Brandt, and Michael S. Selig.Includes bibliographical references.1. Aerofoils (Airfoils). 2. Aerodynamics. 3. Airplanes—Models.

I. Model Aviation. II. Title

Contents

Preface and Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . iii

List of Figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . v

List of Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xi

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiii

Chapter 1 The Airfoils Tested . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

Chapter 2 Experimental Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.1 Experimental Techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.2 Data Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Chapter 3 Summary of Airfoil Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.1 General Comments for Wind Tunnel Tests and Results . . . . . . . . . . . . . . . . . . . . 13

3.2 Discussion of Selected Airfoil Models and Test Conditions . . . . . . . . . . . . . . . . . . 16

Chapter 4 Airfoil Profiles and Performance Plots . . . . . . . . . . . . . . . . . . . . . . . . . . 21

References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 295

Appendix A Tabulated Airfoil Coordinates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 297

Appendix B Tabulated Drag Polar Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 319

ii Summary of Low-Speed Airfoil Data

Preface and Acknowledgments

Summary of Low-Speed Airfoil Data – Volume 5 represents the fifth installment in a series of books doc-umenting the ongoing work of the University of Illinois at Urbana-Champaign Low-Speed Airfoil Tests(UIUC LSATs). The project’s purpose remains unchanged since the original work was performed and doc-umented in Airfoils at Low Speeds (SoarTech 8)1 by Michael Selig, John Donovan, and David Fraser atPrinceton — to develop and test airfoils for low-speed aircraft, in particular, RC model airplanes. However,the application of the results have spread beyond the realm of RC model airplanes to unmanned aerial vehi-cles (UAVs), low-speed propellers, wind turbines, and more since the time SoarTech 8 was published. Thiscurrent project is an example of the broadening scope of the LSATs program with the examination of flapson sailplane airfoils, various leading edges on a flat plate model, and others.

The UIUC LSATs team during the acquisition of the data presented in Volume 5 consisted of MichaelSelig (Assoc. Prof.) as project advisor, Bryan McGranahan (M.S.) as project coordinator from 2002 to2003, John Brandt (M.S.) as project coordinator from 2003 to 2005, and Robert Deters (Ph.D. candidate) asproject coordinator from 2005 to 2010. Testing was conducted during many test campaigns from 2002 to2010.

It is our intention for Summary of Low-Speed Airfoil Data – Volume 5 to follow closely the format ofthe previous volumes making the current installment easy to navigate for those familiar with the series. Asbefore, Chapter 1 discusses the scope of the current tests and briefly describes the airfoils and configurationstested. Chapter 2 gives an overview of the testing facility, LSATs measurement hardware, and flow qual-ity of the UIUC low-speed subsonic wind tunnel. Additionally, Chapter 2 provides a comparison betweenUIUC LSATs data and data obtained by NASA Langley in the Low-Turbulence Pressure Tunnel (LTPT).Chapter 3 discusses the airfoils tested, and Chapter 4 contains the corresponding performance plots, in-cluding pitching-moment data. Appendices A and B list tabulated airfoil coordinates and drag polar datarespectively.

Our research into the aerodynamics of airfoils at low Reynolds number would not have been possiblewithout the generous contributions of many individuals to which are indebted. We thank all of our pastcontributors who are mentioned in our prior volumes. The foundation provided by those supporters pavedthe way to producing the results reported here. For the time period that spans this volume, the generoussupporters who provided funding include Neal Brutsche, Pete Carr, Les Garber, John Hunter, Dave Jones,Mr. Kraus, Ing. Jaroslav Lnenicka, Eric Loos, Larry McNay, Gilbert Morris, Pete Peterson, John Rimmer,Jerry Robertson, Phil Rockwell, Allan Scidmore, Martin Simons, Arthur Slagle, Herk Stokely, Dr. WilfriedStoll, Craig Sutter, and Jose Tellez.

Also greatly appreciated is the time and effort spent by many people constructing the wind tunneltunnel models presented in this volume. These individuals include Thomas Akers (S9000), Mark Drela(hardware fabrication and installation, filling, shaping, and polishing of the AG12, AG16, AG35, AG40d,and AG455ct) and his team Laszlo Horvath (foam core cutting for all), Rob Glover (AG12, AG16 core prepand bagging), John Jenks (AG24, AG35 core prep and bagging), Oleg Golovidov (AG455, AG40 core prep

iv Summary of Low-Speed Airfoil Data

and bagging), Ralph Cooney (MA409), Camille Goudeseune and Michel Goudeseune (Flat Plate), ChrisGreaves (CAL2263m), Tim Lampe (CAL1215j, S8064), Mark Nankivil (NACA 43012A), Jim Thomas(CAL4014l), and Yvan Tinel (E387, S1223).

We are also indebted to our collaborators in the UIUC Aerodynamics Research Laboratory includingMichael Bragg, Greg Elliott, and Andy Broeren (now at NASA Glenn). We also want to thank formerstudents who participated in taking data for this volume: Tomo Sato, Kian Tehrani, and Paul Gush.

And finally, special thanks go to the sponsors of our research in the UIUC Applied AerodynamicsGroup. This ongoing research for our sponsors has been instrumental in maintaining the continuity of ourlow Reynolds number test program and overall laboratory activities and infrastructure. These sponsorsinclude Ford Motorsports, NASA Glenn, AeroVironment, DOE National Renewable Energy Laboratory,WindLite, Newman Haas Racing Naval Research Laboratory, Siemens Canada, Jaguar Racing, Oracle Rac-ing/Farr Yacht Design, Continuum Dynamics, Luna Rossa, Arcturus UAV, Spin Master, ICON Aircraft,FlexSys, Northern Power Systems, GE Energy, and 3M.

List of Figures

1.1 Airfoils tested from Fall 2002 through Summer 2010. . . . . . . . . . . . . . . . . . . . . . . 1

2.1 UIUC low-speed subsonic wind tunnel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2.2 Experimental Setup (Plexiglas R© splitter plates and traverse enclosure box not shown for clarity). 6

2.3 Turbulence intensity taken at tunnel center with the LSATs test apparatus installed with nomodel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2.4 Comparison between UIUC and LTPT E387 drag coefficient data for Re = 60,000, 100,000,200,000, 300,000, and 460,000. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.5 Comparison between UIUC and LTPT E387 lift coefficient data for Re = 60,000, 100,000,200,000, 300,000, and 460,000. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.1 Schematic of the Gurney flap and boundary-layer trip configurations used on the S1223. . . . . 16

3.2 Schematic of the AG35-r showing chord line and wing mounts (drawing by Drela15). . . . . . 17

3.3 AG455ct-02r airfoil with four flap positions −4, −2, 0, and 2 deg, shown with the verticalscale exaggerated 3X (drawing by Drela15). . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.4 Baseline flat plate and various leading-edge configurations drawn to scale (but not full span). . 18

4.1 Comparison between the true and actual AG12. . . . . . . . . . . . . . . . . . . . . . . . . . 30

4.2 Inviscid velocity distributions for the AG12. . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

4.3 Drag polar for the AG12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

4.4 Lift and moment characteristics for the AG12. . . . . . . . . . . . . . . . . . . . . . . . . . . 32








vi Summary of Low-Speed Airfoil Data


4.13 Comparison between the true and actual AG35-r. . . . . . . . . . . . . . . . . . . . . . . . . 48

4.14 Inviscid velocity distributions for the AG35-r. . . . . . . . . . . . . . . . . . . . . . . . . . . 48

4.15 Drag polar for the AG35-r. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

4.16 Lift and moment characteristics for the AG35-r. . . . . . . . . . . . . . . . . . . . . . . . . . 50

4.17 Comparison between the true and actual AG40d-02r. . . . . . . . . . . . . . . . . . . . . . . 54

4.18 Inviscid velocity distributions for the AG40d-02r. . . . . . . . . . . . . . . . . . . . . . . . . 54

4.19 Drag polar for the AG40d-02r. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

4.20 Lift and moment characteristics for the AG40d-02r. . . . . . . . . . . . . . . . . . . . . . . . 56

4.21 Inviscid velocity distributions for the AG40d-02r with a −2 deg flap. . . . . . . . . . . . . . . 60

4.22 Drag polar for the AG40d-02r with a −2 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . 61

4.23 Lift and moment characteristics for the AG40d-02r with a −2 deg flap. . . . . . . . . . . . . . 62

4.24 Inviscid velocity distributions for the AG40d-02r with a 2 deg flap. . . . . . . . . . . . . . . . 66

4.25 Drag polar for the AG40d-02r with a 2 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . . 67

4.26 Lift and moment characteristics for the AG40d-02r with a 2 deg flap. . . . . . . . . . . . . . . 68




4.30 Inviscid velocity distributions for the AG40d-02r with a −15 deg flap. . . . . . . . . . . . . . 78

4.31 Drag polar for the AG40d-02r with a −15 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 79

4.32 Lift and moment characteristics for the AG40d-02r with a −15 deg flap. . . . . . . . . . . . . 80

4.33 Inviscid velocity distributions for the AG40d-02r with a −10 deg flap. . . . . . . . . . . . . . 82

4.34 Drag polar for the AG40d-02r with a −10 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 83

4.35 Lift and moment characteristics for the AG40d-02r with a −10 deg flap. . . . . . . . . . . . . 84

4.36 Inviscid velocity distributions for the AG40d-02r with a −5 deg flap. . . . . . . . . . . . . . . 86

4.37 Drag polar for the AG40d-02r with a −5 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . 87

4.38 Lift and moment characteristics for the AG40d-02r with a −5 deg flap. . . . . . . . . . . . . . 88




4.42 Inviscid velocity distributions for the AG40d-02r with a 10 deg flap. . . . . . . . . . . . . . . 94

List of Figures vii

4.43 Drag polar for the AG40d-02r with a 10 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . 95

4.44 Lift and moment characteristics for the AG40d-02r with a 10 deg flap. . . . . . . . . . . . . . 96









4.53 Drag polar for the gap sealed AG40d-02r. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109

4.54 Lift and moment characteristics for the gap sealed AG40d-02r. . . . . . . . . . . . . . . . . . 110

4.55 Drag polar for the gap sealed AG40d-02r with a −2 deg flap. . . . . . . . . . . . . . . . . . . 113

4.56 Lift and moment characteristics for the gap sealed AG40d-02r with a −2 deg flap. . . . . . . . 114

4.57 Drag polar for the gap sealed AG40d-02r with a 4 deg flap. . . . . . . . . . . . . . . . . . . . 117

4.58 Lift and moment characteristics for the gap sealed AG40d-02r with a 4 deg flap. . . . . . . . . 118

4.59 Aileron Response for the AG40d-02r at Re = 100, 000. . . . . . . . . . . . . . . . . . . . . . 121

4.60 Comparison between the true and actual AG455ct-02r. . . . . . . . . . . . . . . . . . . . . . 122

4.61 Inviscid velocity distributions for the AG455ct-02r with a −0.4 deg flap. . . . . . . . . . . . . 122

4.62 Drag polar for the AG455ct-02r with a −0.4 deg flap. . . . . . . . . . . . . . . . . . . . . . . 123

4.63 Lift and moment characteristics for the AG455ct-02r with a −0.4 deg flap. . . . . . . . . . . . 124




4.67 Inviscid velocity distributions for the AG455ct-02r with a 1.6 deg flap. . . . . . . . . . . . . . 134

4.68 Drag polar for the AG455ct-02r with a 1.6 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 135

4.69 Lift and moment characteristics for the AG455ct-02r with a 1.6 deg flap. . . . . . . . . . . . . 136




4.73 Inviscid velocity distributions for the AG455ct-02r with a −15.4 deg flap. . . . . . . . . . . . 146

viii Summary of Low-Speed Airfoil Data

4.74 Drag polar for the AG455ct-02r with a −15.4 deg flap. . . . . . . . . . . . . . . . . . . . . . 147

4.75 Lift and moment characteristics for the AG455ct-02r with a −15.4 deg flap. . . . . . . . . . . 148

4.76 Inviscid velocity distributions for the AG455ct-02r with a −10.4 deg flap. . . . . . . . . . . . 150

4.77 Drag polar for the AG455ct-02r with a −10.4 deg flap. . . . . . . . . . . . . . . . . . . . . . 151

4.78 Lift and moment characteristics for the AG455ct-02r with a −10.4 deg flap. . . . . . . . . . . 152










4.88 Inviscid velocity distributions for the AG455ct-02r with a 14.6 deg flap. . . . . . . . . . . . . 166

4.89 Drag polar for the AG455ct-02r with a 14.6 deg flap. . . . . . . . . . . . . . . . . . . . . . . 167

4.90 Lift and moment characteristics for the AG455ct-02r with a 14.6 deg flap. . . . . . . . . . . . 168

4.91 Drag polar for the gap sealed AG455ct-02r with a −0.4 deg flap. . . . . . . . . . . . . . . . . 169

4.92 Lift and moment characteristics for the gap sealed AG455ct-02r with a −0.4 deg flap. . . . . . 170

4.93 Drag polar for the gap sealed AG455ct-02r with a −2.4 deg flap. . . . . . . . . . . . . . . . . 171

4.94 Lift and moment characteristics for the gap sealed AG455ct-02r with a −2.4 deg flap. . . . . . 172

4.95 Drag polar for the gap sealed AG455ct-02r with a 3.6 deg flap. . . . . . . . . . . . . . . . . . 173

4.96 Lift and moment characteristics for the gap sealed AG455ct-02r with a 3.6 deg flap. . . . . . . 174

4.97 Aileron Response for the AG455ct-02r at Re = 60, 000. . . . . . . . . . . . . . . . . . . . . 175

4.98 Aileron Response for the AG455ct-02r at Re = 100, 000. . . . . . . . . . . . . . . . . . . . . 176

4.99 Comparison between the true and actual CAL1215j. . . . . . . . . . . . . . . . . . . . . . . . 178

4.100 Inviscid velocity distributions for the CAL1215j. . . . . . . . . . . . . . . . . . . . . . . . . 178

4.101 Drag polar for the CAL1215j. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179

4.102 Lift and moment characteristics for the CAL1215j. . . . . . . . . . . . . . . . . . . . . . . . 180

4.103 Comparison between the true and actual CAL2263m. . . . . . . . . . . . . . . . . . . . . . . 184

4.104 Inviscid velocity distributions for the CAL2263m. . . . . . . . . . . . . . . . . . . . . . . . . 184

List of Figures ix

4.105 Drag polar for the CAL2263m. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185

4.106 Lift and moment characteristics for the CAL2263m. . . . . . . . . . . . . . . . . . . . . . . . 186

4.107 Comparison between the true and actual CAL4014l. . . . . . . . . . . . . . . . . . . . . . . . 190

4.108 Inviscid velocity distributions for the CAL4014l. . . . . . . . . . . . . . . . . . . . . . . . . 190

4.109 Drag polar for the CAL4014l. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 191

4.110 Lift and moment characteristics for the CAL4014l. . . . . . . . . . . . . . . . . . . . . . . . 192

4.111 Comparison between the true and actual E387 (E). . . . . . . . . . . . . . . . . . . . . . . . . 196

4.112 Inviscid velocity distributions for the E387 (E). . . . . . . . . . . . . . . . . . . . . . . . . . 196

4.113 Drag polar for the E387 (E). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 197

4.114 Lift and moment characteristics for the E387 (E). . . . . . . . . . . . . . . . . . . . . . . . . 198

4.115 Schematic of the baseline leading edge configuration. . . . . . . . . . . . . . . . . . . . . . . 202

4.116 Lift and moment characteristics for a flat plate with the baseline leading edge. . . . . . . . . . 203

4.117 Schematic of the leading edge serrations (Case A) configuration. . . . . . . . . . . . . . . . . 206

4.118 Lift and moment coefficient characteristics for a flat plate with leading edge serrations (Case A).207

4.119 Schematic of the leading edge serrations (Case B) configuration. . . . . . . . . . . . . . . . . 210

4.120 Lift and moment coefficient characteristics for a flat plate with leading edge serrations (Case B).211

4.121 Schematic of the leading edge serrations (Case C) configuration. . . . . . . . . . . . . . . . . 214

4.122 Lift and moment coefficient characteristics for a flat plate with leading edge serrations (Case C).215

4.123 Schematic of the leading edge serrations (Case D) configuration. . . . . . . . . . . . . . . . . 218

4.124 Lift and moment coefficient characteristics for a flat plate with leading edge serrations (Case D).219

4.125 Schematic of the leading edge square wave configuration. . . . . . . . . . . . . . . . . . . . . 222

4.126 Lift and moment characteristics for a flat plate with leading edge square waves. . . . . . . . . 223

4.127 Schematic of the leading edge configuration with small holes. . . . . . . . . . . . . . . . . . . 226

4.128 Lift and moment characteristics for a flat plate with small holes on the leading edge. . . . . . . 227

4.129 Schematic of the leading edge configuration with large holes. . . . . . . . . . . . . . . . . . . 230

4.130 Lift and moment characteristics for a flat plate with large holes on the leading edge. . . . . . . 231

4.131 Schematic of the leading edge configuration with small cubes. . . . . . . . . . . . . . . . . . 232

4.132 Lift and moment characteristics for a flat plate with small cubes on the leading edge. . . . . . 233

4.133 Schematic of the leading edge configuration with large cubes. . . . . . . . . . . . . . . . . . . 236

4.134 Lift and moment characteristics for a flat plate with large cubes on the leading edge. . . . . . . 237

4.135 Comparison between the true and actual MA409. . . . . . . . . . . . . . . . . . . . . . . . . 240

x Summary of Low-Speed Airfoil Data

4.136 Inviscid velocity distributions for the MA409. . . . . . . . . . . . . . . . . . . . . . . . . . . 240

4.137 Drag Polar for the MA409 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 241

4.138 Lift and moment characteristics for the MA409. . . . . . . . . . . . . . . . . . . . . . . . . . 242

4.139 Comparison between the true and actual NACA 43012A. . . . . . . . . . . . . . . . . . . . . 246

4.140 Inviscid velocity distributions for the NACA 43012A. . . . . . . . . . . . . . . . . . . . . . . 246

4.141 Drag polar for the NACA 43012A. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 247

4.142 Lift and moment characteristics for the NACA 43012A. . . . . . . . . . . . . . . . . . . . . . 248

4.143 Comparison between the true and actual S1223. . . . . . . . . . . . . . . . . . . . . . . . . . 252

4.144 Inviscid velocity distributions for the S1223. . . . . . . . . . . . . . . . . . . . . . . . . . . . 252

4.145 Lift and moment characteristics for the S1223. . . . . . . . . . . . . . . . . . . . . . . . . . . 253

4.146 Lift and moment characteristics for the S1223 with Gurney flap of h/c = 4.17%. . . . . . . . 257





4.151 Lift and moment characteristics for the S1223 with a boundary-layer trip of t/c = 0.11%. . . . 269

4.152 Lift and moment characteristics for the S1223 with a boundary-layer trip of t/c = 0.19%. . . . 271



4.155 Drag polar for the S8064. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273




4.159 Drag polar for the S9000. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 279


4.161 Inviscid velocity distributions for the S9000 with a 2.5 deg flap. . . . . . . . . . . . . . . . . . 284

4.162 Drag polar for the S9000 with a 2.5 deg flap . . . . . . . . . . . . . . . . . . . . . . . . . . . 285

4.163 Lift and moment characteristics for the S9000 with a 2.5 deg flap . . . . . . . . . . . . . . . . 286

4.164 Inviscid velocity distributions for the S9000 with a 5 deg flap. . . . . . . . . . . . . . . . . . . 290

4.165 Drag polar for the S9000 with a 5 deg flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . 291

4.166 Lift and moment characteristics for the S9000 with a 5 deg flap . . . . . . . . . . . . . . . . . 292

List of Tables

3.1 Airfoils Tested . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

3.2 Airfoil Model Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.3 Gurney Flap Thickness and Height Test Configurations and Maximum Lift CoefficientResults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

3.4 Boundary-Layer Trip Thickness and Width Test Configurations . . . . . . . . . . . . . . . 19

4.1 Test Matrix and Run Number Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

xii Summary of Low-Speed Airfoil Data

Nomenclature

Symbols

c airfoil chord

Cl airfoil lift coefficient

cf flap chord

cf/c flap-chord ratio

Cd airfoil drag coefficient

Cm airfoil moment coefficient about the quarter chord

h Gurney flap height

Re Reynolds number based on airfoil chord

t airfoil thickness, or trip height

t/c airfoil thickness ratio, or trip height ratio

w trip width

x distance from leading edge

α angle of attack

δf flap deflection

Abbreviations

LSATs Low-Speed Airfoil Tests

LTPT Low-Turbulence Pressure Tunnel at NASA Langley Research Center

RC Radio Controlled

UAV Unmanned Aerial Vehicle

UIUC University of Illinois at Urbana-Champaign

xiv Summary of Low-Speed Airfoil Data

Chapter 1

The Airfoils Tested

This volume of Summary of Low-Speed Airfoil Data documents the wind-tunnel test results of 16 airfoils(shown in Fig. 1.1) ranging from a flat plate model with various leading edge configurations to the highlift S1223 with boundary-layer trips and Gurney flaps. Previous LSATs volumes1–5 have documented theperformance of over 100 other airfoil wind tunnel models.

AG12 CAL1215j MA409

AG16 CAL2263m NACA 43012A

AG24 CAL4014l S1223

AG35−r E387 S8064

AG40d−02r Flat Plate S9000

AG455ct−02r

Fig. 1.1: Airfoils tested from Fall 2002 through Summer 2010.

The AG airfoils designed by Drela6 for sailplane applications can be seen in the left column of Fig. 1.1.Most of the tests conducted on these six airfoils were performed at the lower end of the Reynolds numberrange of the LSATs setup (Re ≤ 100,000). These airfoils were designed to provide good penetration at lowReynolds numbers. The AG40d-02r and AG455ct-02r were tested with numerous flap deflections rangingfrom −15 to 30 deg, but most of the tests were performed between −2 and 4 deg.

The first three airfoils in the middle column are the CAL airfoils designed by Christopher Lyon—formermember of the UIUC LSATs team and the author of Volume 3. The three airfoils tested for Volume 5 haddifferent design goals and therefore applications. The CAL1215j airfoil is a derivative of the Clark-Y airfoil

2 Summary of Low-Speed Airfoil Data

and is designed to operate at slightly lower Cl values compared with the Clark-Y. The Clark-Y airfoil waspreviously tested and documented in SoarTech 8 and Volume 3. The CAL2263m is also a “derivative” of theClark-Y airfoil, but it was designed to obtain lower drag with similar stall characteristics compared with theClark-Y. The CAL2263m sports a flat bottom aft of 30% of chord for ease of construction. The CAL4014lis a reflexed flying-wing airfoil similar to the MH45 airfoil. The MH45 airfoil was previously tested anddocumented in Volume 1.

The Eppler E387 airfoil has been the benchmark airfoil for data validation since the inception of theLSATs program at UIUC. For that reason, it is included in Volume 5. The E387 has been extensively studiedby many researchers for comparisons between wind tunnel facilities.

The flat plate model was tested with various leading edge configurations from leading edge serrations(shark tooth) to cubes placed on the upper surface near the leading edge. These tests examined the effectsof these various leading edge configurations on lift and moment characteristics. The motivation for thesetests was driven by Camille Goudeseune, and his interests in the prevalence of various leading-edge con-figurations that are found in nature, such as on the flippers of humpback whales.7 All of the leading edgeconfigurations were examined at the lower range of the LSATs setup (Re ≤ 120,000).

The MA409 airfoil was designed by Michael Achterberg—former US F1C Team Member. It wasdesigned for applications in F1C class freeflight and has proven itself in numerous unlimited flyoffs for fastclimb and good glide endurance. The MA409 has a low zero-lift drag coefficent, which aids in the overallperformance. It should be noted that the MA409 was tested in Volume 1, but the wind tunnel model hadwarped since those tests. Thus, the MA409 was redigitized when it was tested for this volume. Thus, theperformance results seen in Volume 5 do not reflect the true airfoil performance nor agree with results fromVolume 1.

The NACA 43012A is the airfoil used on the Schweizer SGS 1-26, single-seat, mid-wing glider andthe Schweizer SGS 2-33, two-seat, high-wing training glider. Because the airfoil proved successful on afull-scale glider, there was interest on the part of the model maker in applying this airfoil to RC sailplanes,which operate at significantly lower Reynolds numbers. Therefore, the LSATs group was asked to examinethe performance of the NACA 43012A at the Reynolds numbers experienced by RC gliders. The actualNACA 43012A wind tunnel model coordinates as tested do not agree well with the true NACA 43012Acoordinates. Thus, the performance results here in Volume 5 do not reflect the performance of the trueairfoil.

The high lift S1223 airfoil has been extensively tested in previous volumes of Summary of Low-SpeedAirfoil Data. The S1223 airfoil is able to obtain extremely high lift coefficients (Cl ≈ 2.2) for a singleelement (no slats or flaps) because the design philosophy combined the favorable effects of a concave pres-sure recovery and aft loading.2 At the design Reynolds number of 200,000, a Clmax of 2.23 is obtained asseen in Fig. 4.145, but it exhibits a large hysteresis loop at Re ≤ 200,000. Thus, boundary-layer trips withdifferent heights were examined in an attempt to remove the hysteresis loop. This examination proved tobe successful. Gurney flaps of varying heights were also examined in an attempt to incrementally increaseClmax , which also proved successful.

The S8064 airfoil was designed by Selig for application to Quickie 500 RC racing. It is currently usedon the Viper 500 RC airplane by Great Planes Model Manufacturing Company.8 In straight flight, the S8064was designed to operate at a lift coefficient of 0.0 to 0.05. Lift coefficients in a turn were taken to be 0.4 to0.6 during the design process.9

Chapter 1: The Airfoils Tested 3

The S9000 airfoil was designed by Selig for RC sailplane applications and used on the carbon Black-hawk 133.5-in span (open-class) RC sailplane designed around 1991. The S9000 coordinates were madepublic on December 6, 200210 and wind tunnel tested thereafter.

Additional airfoils not included here in Volume 5 have been tested since Volume 4 was published. Thesetests, found in Refs. 11 and 12, document the performance of the ND-LoFoil, S1052, and S1054 airfoils.The ND-LoFoil airfoil was tested with and with boundary-layer trips. The S1052 and S1054 airfoils wereflapped and intended to be used on flying-wing UAVs.

Chapter 2

Experimental Methods

All experiments were performed in the UIUC Department of Aerospace Engineering Aerodynamics Re-search Laboratory. The low-speed subsonic wind tunnel has been in service at the University of IllinoisUrbana-Champaign since the early 1990s and has been used for all of the Summary of Airfoil Data vol-umes.2–5 Summarized descriptions of the low-speed wind tunnel, lift and drag measurement techniques,and data validation are presented in this chapter. A more detailed discussion of the experimental techniquesused to collect the airfoil performance data can be found in Refs. 1–5.

2.1 Experimental Techniques

Testing was conducted in the UIUC low-turbulence subsonic wind tunnel seen in Fig. 2.1. The wind tunnel isan open-return type with a 7.5:1 contraction ratio. The rectangular test section is 2.8×4.0 ft in cross sectionand 8-ft long. Over the length of the test section, the width increases by approximately 0.5 in to accountfor boundary-layer growth along the tunnel side walls. Test-section speeds are variable up to 160 mph via a125-hp alternating current electric motor driving a five-bladed fan. The tunnel settling chamber contains a4-in thick honeycomb and four anti-turbulence screens to ensure adequate turbulence levels. All of the datapresented in this volume were taken with the LSATs rig where the models have a nominal 12-in chord and33 5/8-in span. The LSATs test apparatus has a Reynolds number range of 40,000 to 500,000. A Reynoldsnumber of 500,000 required a nominal test section speed of 80 ft/sec (54.5 mph).

The LSATs experimental setup, depicted in Fig. 2.2, has several unique features that make it distinctfrom all the other experimental setups used in the UIUC low-turbulence subsonic wind tunnel. The airfoilswere mounted horizontally and isolated from the tunnel side-wall boundary layers and the support hardwareby two 3/8-in thick and 6-ft long Plexiglas R© splitter plates. One side of the airfoil model (left) was freeto rotate, and it included a rotary potentiometer to measure the angle of attack. On the other side (right), amotor with a worm-drive system was used to set the angle of attack. The motor was mounted to a carriagethat was free to move vertically on a precision-ground shaft but was not free to rotate. This carriage wasconnected to a pushrod that transferred the lift force to a load cell via a fulcrum-supported beam outsideof the tunnel. Also connected to the carriage were two lever arms. Between these two arms, a load cellwas connected to measure the pitching moment of the airfoil. The mechanical arrangement for the momentmeasurement and the overall system calibration procedure is detailed in Volume 3.4

Drag was measured using a wake rake that included eight total pressure probes over a spanwise distanceof 10.5 in. These probes were placed horizontally in order to measure eight drag profiles over the midspanof the model. The wake rake was moved vertically to capture the wake profiles, and pressure measurements


Silencer

Fan

Diffuser

Test Section

Inlet

Fig. 2.1: UIUC low-speed subsonic wind tunnel.

were taken about every 0.085 in. The drag values calculated from each of the eight wakes were averaged.

An extensive study of the flow quality of the UIUC low-speed subsonic wind tunnel was conducted inVolume 4. The study included measurements of the freestream turbulence, variation in dynamic pressure(related to velocity) across the test section, and the angle of the flow relative to the centerline of the tunnel.The turbulence intensity results can be seen in Fig. 2.3. The effect of the LSATs test apparatus on theturbulence intensity was not constant with respect to Reynolds number. It can be seen from Fig. 2.3 that ata Reynolds number of 100,000 the turbulence intensity was relatively unchanged by the addition of the testapparatus, but there was an increase in turbulence intensity for Re ≥ 200,000. By adding a 3-Hz high-passfilter, the lower turbulence intensity indicates the LSATs test apparatus mainly affected by low frequencyrange by adding velocity fluctuations.5

Traverse

Lift Load Cell

Precision Bearing

Lift Balance Beam

Counter Weight

Wake Probes

Lift Carriage

Motor

Pushrod

Sensor

Sector Gear

Moment Load Cell

αα

Fig. 2.2: Experimental Setup (Plexiglas R© splitter plates and traverse enclosure box not shown for clarity).

Chapter 2: Experimental Methods 7

100,000 200,000 300,000 400,000 500,0000.05

0.1

0.15

0.2

0.25

Reynolds Number

Mea

n T

urbu

lenc

e In

tens

ity (

%)

Average Turbulence Intensity (Empty Test Section)

DC CoupledHPF = 0.1HzHPF = 3HzHPF = 10Hz

100,000 200,000 300,000 400,000 500,0000.05

0.1

0.15

0.2

0.25

Reynolds Number

Mea

n T

urbu

lenc

e In

tens

ity (

%)

Average Turbulence Intensity (Test Apparatus Installed)

DC CoupledHPF = 0.1HzHPF = 3HzHPF = 10Hz

Fig. 2.3: Turbulence intensity taken at tunnel center with the LSATs test apparatus installed with no model.

The importance of knowing the shape of the actual airfoil being tested is a major cornerstone of airfoiltesting because slight variations from the true airfoil shape can have large effects on the airfoil performance.Therefore, it is important to measure the actual shape of the airfoil because no model can be made withouterror. The actual coordinates of the airfoils tested were measured with a coordinate measuring machine andare given in Appendix A. The coordinates of the true airfoils (as designed) are given in Appendix A.

2.2 Data Validation

Data validation is an important aspect of instilling confidence in any experiment. Perhaps the simplestway to validate wind-tunnel data is through comparison with a known standard. Determining exactly whatthat standard is, however, proves to be difficult. Although no specific facility produces perfect data, thereare those considered better than others based on tunnel turbulence level, test-section geometry, and modelquality. The Low-Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center, with its lowturbulence and tall test section, produces high quality data.13 For this reason, data taken in the LTPT will beused as the standard.

The airfoil data reported here was taken over several years, and validation runs were performed at thebeginning of each wind tunnel entry. The results from each validation session showed results similar to theSpring 2003 validation data. Therefore, only the Spring 2003 lift, drag, and moment data are used in thischapter as seen in Figs. 2.4 and 2.5. Figure 2.4 shows a comparison between UIUC LSATs and LTPT dragpolars for the E387 airfoil. The LSATs drag data compares fairly well with that taken at NASA Langley. Thelift and moment data (see Fig. 2.5) also shows good agreement with LTPT data for all Reynolds numbers upto stall, after which point three-dimensional end effects are likely to be the cause for the slight discrepancies.More detailed information regarding data validation including oil flow visualization comparisons betweenthe UIUC LSATs and LTPT E387 can be found in Ref. 5.


0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl

UIUC model E: tested Spring 2003

LTPT model

E387

Re = 60,000

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl


LTPT model

E387

Re = 100,000

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl


LTPT model

E387

Re = 200,000

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl


LTPT model

E387

Re = 300,000

Fig. 2.4: Comparison between UIUC and LTPT E387 drag coefficient data for Re = 60,000, 100,000,200,000, 300,000, and 460,000.


0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl


LTPT model

E387

Re = 460,000

Figure 2.4: Continued.


−10 0 10 20−0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 Re = 60,000

LTPT modelLift Moment

UIUC model E: tested Spring 2003Lift Moment

Cm

0.1

0.0

−0.1

−0.2

−0.3

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 Re = 100,000



Cm

0.1

0.0

−0.1

−0.2

−0.3

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 Re = 200,000



Cm

0.1

0.0

−0.1

−0.2

−0.3

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 Re = 300,000



Cm

0.1

0.0

−0.1

−0.2

−0.3

Fig. 2.5: Comparison between UIUC and LTPT E387 lift coefficient data for Re = 60,000, 100,000,200,000, 300,000, and 460,000.


−10 0 10 20−0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 Re = 460,000



Cm

0.1

0.0

−0.1

−0.2

−0.3

Figure 2.5: Continued.

Chapter 3

Summary of Airfoil Data

Chapter 3 along with Chapter 4 are the heart of Summary of Low-Speed Airfoil Data. In these two chapters,the airfoil models are discussed and the collected data are presented. Table 3.1 lists each airfoil along withits thickness, camber, flap-chord (cf/c) ratio, flap hinge point, and representative quarter chord pitchingmoment (Cm,c/4). In the table, only airfoils that list a flap-chord ratio and hinge point were tested with atrailing edge flap. Table 3.2 lists the model construction method (surface finish), model accuracy, and modelbuilder(s). More detailed descriptions of the airfoil configurations are included in the profile plots providedin Chapter 4.

3.1 General Comments for Wind Tunnel Tests and Results

In this section, general comments about the wind tunnel models, test configurations, and data presented inChapter 4 are listed below. Therefore, the following comments are aimed to aid the reader in understandingthe experimental data presented here.

• The suffixes ‘(A)’, ‘(B)’, etc. refer to multiple versions of a particular airfoil. Owing to constructioninaccuracies, it is preferred to test multiple models of the same airfoil for two reasons. First, because eachmodel may vary from the true airfoil geometry in a slightly different way, a measure of the sensitivityof each airfoil to changes in geometry is attained, and second, by having several models constructed,the probability of receiving an extremely accurate model is increased. For this volume, the E387(E)represents the fifth E387 airfoil model tested. This “E” model is the most accurate E387 that has beentested in the series.

• The discussion of each airfoil is based on the actual shape of the model and not the designed true airfoilgeometry. When the average difference between the actual and true airfoil coordinates is large (greaterthan 0.010 in for a 12-in chord), the wind-tunnel data may not be an accurate representation of the trueairfoil performance. A better indication of the effects contour inaccuracies may have on performancecan be realized from the airfoil accuracy plots in Chapter 4. Differences at the trailing edge behave likecamber changes and therefore affect the useful lift range. Inaccuracies along the upper surface can havea large influence on bubble drag, while lower-surface errors are typically not as critical. Finally, if theairfoil is uniformly thicker, the performance will be more indicative of the true airfoil than if the surfaceerror is “wavy” or sloped.

• In some instances, the average difference listed in Table 3.2 may not be indicative of the actual modelaccuracy. Specifically, for the AG40d-02r, AG455ct-02r, and S9000, the 0-deg flap setting was offset by


Table 3.1: Airfoils Tested

Airfoil Thickness (%) Camber (%) cf/c (%) Hinge Point Cm,c/4

AG12 6.24 1.85 – – −0.040AG16 7.11 1.88 – – −0.050AG24 8.41 2.21 – – −0.060AG35-r 8.73 2.37 – – −0.050AG40d-02r 8.00 2.37 25 bottom −0.060AG455ct-02r 6.47 2.28 30 bottom −0.050CAL1215j 11.72 2.29 – – −0.060CAL2263m 11.72 3.54 – – −0.080CAL4014l 10.00 1.85 – – 0.005E387 (E) 9.06 3.79 – – −0.085Flat Plate 3.10 0.00 – – 0.000MA409 6.69 3.33 – – −0.050NACA 43012A 12.22 2.67 – – −0.010S1223 11.93 8.67 – – −0.290S8064 12.39 1.19 – – −0.024S9000 9.00 2.37 20 bottom −0.060

a slight increment. This offset, when digitized and compared with the true airfoil, appears as an error inthe shape when in fact it is slight flap offset and not an actual shape error.

• With some airfoils, the original number of coordinates used to define the geometry was not satisfac-tory to provide a smooth airfoil surface. Thus, the airfoil was mathematically smoothed to provide newcoordinates. In this volume, only the coordinates for the MA409 were smoothed.

• When characterizing surface finish qualities in Table 3.2, the term “smooth” refers to a fiberglass surfaceapplied via the vacuum bag method. This kind of smoothness is distinguished from a smooth mylarcovering that may show imperfections in the underlying layer.

• The models with trailing-edge flaps were hinged on the lower (bottom) surface as indicated in Table 3.1.The flaps were hinged with hinge tape and when indicated, sealed on the upper surface with hinge-gapsealing tape. The various sizes of the trailing edge flaps can be seen in Table 3.1

• The inviscid velocity distributions shown for the true airfoils in Chapter 4 were calculated using XFOIL.14

Flap effects were included when appropriate. For those interested in boundary layer effects on these dis-tributions, they are encouraged to read Volume 2, which contains predicted viscous velocity distributionsfor several airfoils as well as a brief discussion on their interpretation. With experience, much can begleaned from both the viscous and inviscid results to help interpret the airfoil drag polars and lift curves.

• The figures in Chapter 4 list the nominal Reynolds numbers. Actual Reynolds numbers can be found inAppendix B for each run. In most cases, the difference is typically no larger than ∆Re = 100 to 200.

• As stated previously, all drag coefficients were obtained by averaging spanwise drag coefficients fromeight wake surveys spaced 1.5 in apart approximately 1.25 chord lengths downstream of the model trailingedge. These spanwise coefficients are not documented here but are available upon request.

Chapter 3: Summary of Airfoil Data 15

Table 3.2: Airfoil Model Characteristics

Airfoil Surface Finish Avg. Diff. (in) Builder

AG12 smooth 0.0041 M. Drela, et al.AG16 smooth 0.0020 M. Drela, et al.AG24 smooth 0.0047 M. Drela, et al.AG35-r smooth 0.0042 M. Drela, et al.AG40d-02r smooth 0.0069 M. Drela, et al.AG455ct-02r smooth 0.0080 M. Drela, et al.CAL1215j mylar over balsa 0.0080 T. LampeCAL2263m varnish over balsa 0.0063 C. GreavesCAL4014l smooth 0.0094 J. ThomasE387 (E) smooth 0.0091 Y. TinelFlat Plate mylar over wood – C. Goudeseune & M. GoudeseuneMA409 smooth 0.0358* R. CooneyNACA 43012A mylar over balsa 0.0430 M. NankivilS1223 smooth 0.0099 Y. TinelS8064 mylar over balsa 0.0065 T. LampeS9000 smooth 0.0055 T. Akers

*Smoothed model coordinates were taken as true coordinates

• For the lift curves, increasing and decreasing angles of attack are denoted by solid triangles and opencircles, respectively.

• For the moment curves, increasing and decreasing angles of attack are denoted by inverted solid trianglesand open rectangles, respectively. As has become convention when plotting airfoil pitching-moment data,positive values of Cm,c/4 are at the bottom of the plot (i.e., the Cm axis is inverted).

• The pitching-moment data presented in Table 3.1 are representative values for the typical operating pointof the airfoil. As the results show, however, many airfoils exhibit large changes in Cm,c/4 as both angleof attack and Reynolds number change. Because of these large changes, the values listed in Table 3.1should only be used for comparative purposes. More accurate pitching-moment data than that providedin Table 3.1 is often necessary for detailed stability and control calculations. In such case, the variationsin Cm,c/4 should be accounted for by using the full set of pitching-moment data. This data, as plotted inChapter 4, can be obtained in tabulated form upon request.

• The use of boundary-layer trips on the upper and lower surfaces of airfoils is typically done to study theeffects of roughness on airfoil performance or to mitigate the effects of laminar separation bubbles. Avariety of trip geometries on various airfoils have been studied in past volumes of Summary of Low-SpeedAirfoil Data.1–5 With 2-D boundary-layer trips, the three most important parameters used to define thetrip are the airfoil surface (upper and/or lower) where the trip is located, the x-location of the trip (x/c),and the trip thickness (t/c). As is standard in most boundary-layer trip studies, “u.s.t” is used to signifythat the trip is placed on the upper surface of the airfoil while “l.s.t” is used for the lower surface of theairfoil. The x-location of the aft edge of the trip measured from the leading edge of the airfoil is given asa percent chord (x/c) to provide a nondimensionalized value. As with the x-location, the trip thickness is


Plain trip

Thickness: t

Width: w

th

Gurney

flap

Doublesidedtape

S1223

x/c (%)

Fig. 3.1: Schematic of the Gurney flap and boundary-layer trip configurations used on the S1223.

nondimensionalized by the airfoil chord (t/c) and given as a percentage. A schematic of the trip geometrycan be seen in Fig. 3.1. In this volume, two plain rectangular boundary-layer trips of different heightswere used on the S1223 and placed only on the upper surface.

• Gurney flaps are placed at the trailing edge of an airfoil and used to slightly increase the lift while min-imally affecting drag. Gurney flaps have been previously studied in past volumes of Summary of Low-Speed Airfoil Data2, 3 typically on high lift airfoils. The height of the Gurney flap, which drives theeffectiveness, is defined as the distance below the trailing edge normal to the airfoil lower surface at thetrailing edge. It is nondimensionalized by the airfoil chord (h/c) and given as a percentage. A schematicof a Gurney flap can be seen in Fig. 3.1. In this volume, five Gurney flaps of differing heights were usedon the S1223.

3.2 Discussion of Selected Airfoil Models and Test Conditions

AG35-r: The AG35 airfoil provided to the UIUC LSATs was rotated −1.563 deg to orient the chord linehorizontally. To avoid confusion, this rotated airfoil is defined as the AG35-r. Thus, the “-r” indicated thatthe airfoil has been rotated. The velocity distribution plot, performance plots, and tabulated coordinates areall presented with respect to the AG35-r airfoil. A schematic of the AG35-r airfoil with the chord line andwing mounts can be seen in Fig. 3.2.


AG35-r (AG35 rotated −1.563 deg)

Fig. 3.2: Schematic of the AG35-r showing chord line and wing mounts (drawing by Drela15).

-4 deg

-2

0

2

Fig. 3.3: AG455ct-02r airfoil with four flap positions −4, −2, 0, and 2 deg, shown with the vertical scaleexaggerated 3X (drawing by Drela15).

AG40d-02r & AG455ct-02r: Both the AG40d-02r and AG455ct-02r airfoils have a preset absolute flapdeflection of −2 deg incorporated into the airfoil coordinates. As seen in top view in Fig. 3.3, the chordline is referenced to this configuration. Thus, the angle of attack is also referenced with respect to thisconfiguration. The convention of defining the flap deflections used in Chapter 4 can be seen in Fig. 3.3.The AG455ct-02r airfoil set to an absolute flap deflection of −2 deg had a 2-deg inside corner on the uppersurface at the hinge line location while the lower surface was flat at the hinge point. For an absolute flapdeflection of 0 deg, the lower surface had a 2-deg inside corner while the upper surface was flat.15 The sameholds true for the AG40d-02r airfoil.

Flat Plate: The flat-plate model (t/c = 3.10%) was tested with the ten various leading-edge configurationsseen in Fig. 3.4 to study the effect on the lift and moment characteristics. Having a modified leading edgecan in some cases delay stall, increase maximum lift, and lower drag.7 The planform area of the windtunnel models was the same for each configuration. Therefore, the chord of each configuration was takento be 12 in. The baseline leading-edge model represents the standard flat-plate configuration used on manypopular light-weight flat-foam RC models, albeit with varying thicknesses.

Four serrated leading-edge configurations (second row in Fig. 3.4) were tested with varying amplitudesand wavelengths of the serrations. A slight variation of the serrated leading edge was the square-waveleading edge, which was tested with only one amplitude and wavelength.

The last four configurations (bottom row in Fig. 3.4) deviated from the others by leaving the leadingedge flat but adding features just rear of the leading edge. Of these last four configurations, there were twomain types. These two main types were holes near the leading edge (small holes and large holes) and cubeson the upper surface near the leading edge (small cubes and large cubes). In all four configurations (bothtypes), the holes/cubes were placed such that a 0.125-in gap existed between the hole/cube and the leadingedge. In Fig. 3.4, the drawings include both front and top views.


Baseline

Square Wave

Amplitude = 0.01575c

Wavelength = 0.063c

Serrations A

Amplitude = 0.013c

Wavelength = 0.052c

Serrations B

Amplitude = 0.049c

Wavelength = 0.052c

Serrations C

Amplitude = 0.013c

Wavelength = 0.125c

Serrations D

Amplitude = 0.052c

Wavelength = 0.125c

Small Holes

Hole diameter = 0.008c

LE to hole center = 0.014c

Spacing = 0.016c

Large Holes



Spacing = 0.052c

Small Cubes

Cube height = 0.005c

LE to cube front edge = 0.010c

Spacing = 0.010c

Large Cubes



Spacing = 0.042c

Fig. 3.4: Baseline flat plate and various leading-edge configurations drawn to scale (but not full span).


Table 3.3: Gurney Flap Thickness and Height Test Configurations and Maximum Lift Coefficient Results

Airfoil t/c (%) h/c (%) Clmax at Re = 250, 000 Figure/Page

S1223 – – 2.25 Fig. 4.145/p. 2530.13 4.17 2.52 Fig. 4.146/p. 2570.13 3.12 2.46 Fig. 4.147/p. 2590.13 2.08 2.43 Fig. 4.148/p. 2630.13 1.56 2.34 Fig. 4.149/p. 2650.13 1.04 2.36 Fig. 4.150/p. 267

Table 3.4: Boundary-Layer Trip Thickness and Width Test Configurations

Airfoil x/c (%) from aft edge t/c (%) w/c (%) Figure/Page

S1223 clean clean clean Fig. 4.145/p. 25315 0.11 2.60 Fig. 4.151/p. 26915 0.19 2.60 Fig. 4.152/p. 271

S1223: Three different configurations of the S1223 were tested. First, the S1223 was tested in a clean(baseline) configuration for comparison. The second configuration involved testing Gurney flaps of fivesizes ranging from h/c = 1.04% to 4.17%. As discussed previously, the height of the Gurney flap seen inFig. 3.1 is defined as the distance below the trailing edge normal to the airfoil lower surface at the trailingedge. The thickness of the Gurney flap was held constant at t/c = 0.13% (1/64 in). All of the test parameterscan be seen in Table 3.3. Along with the test parameters, the maximum lift coefficient for a Reynoldsnumber of 250,000 is also given in Table 3.3 to show the effects of the Gurney flap on maximum lift. Thethird configuration involved testing boundary-layer trips of two thicknesses (t/c = 0.11% and 0.19%) onthe upper surface at an x/c-location of 30% with respect to the aft edge of the trip. The width of the tripwas held constant at w/c = 2.60% (5/16 in). All of the test parameters can be seen in Table 3.4. For boththe Gurney flap and boundary-layer trip configurations, the figure and page numbers for each configurationare given in Table 3.3 and 3.4 respectively.

Chapter 4

Airfoil Profiles and Performance Plots

In this chapter, the airfoil profiles and performance plots are presented. For quick reference, the airfoilnames and important parameters are listed in the margins, e.g. parameters such as the flap-chord ratio(cf/c), flap deflection, boundary-layer trip size and location, Gurney flap size, and others. As a matter ofrecord, a table listing all the data sets, associated figures, figure page numbers, as well as run numbers isincluded at the beginning of this chapter. In the table, the two letters used in the run numbers correspondto the initials of the person who led the specific test. For some cases in the table, the drag data for a givenReynolds number was made from a composite of two individual runs. For these cases, two run numbers arelisted in the table.

The ‘Avg. Difference’ between the true and actual coordinates listed in the comparison plots is theaverage error in model construction. The accuracy of each model is graphically depicted before the data forthat airfoil is presented. The upper plot shows the actual digitized airfoil (dotted line) co-plotted with thetrue as designed airfoil (solid line). This plot allows the reader to see the full scale error of the actual model,but for accurate models, the discrepancy between the actual and true airfoil cannot be seen with this plot.Therefore, the lower plot was produced to show the discrepancies between the actual and true airfoil upper(solid line) and lower (dotted line) surfaces on a finer scale. The horizontal axis represents the true airfoiland the difference above or below the horizontal axis represents the error of the actual airfoil. For example,if the lower surface of the actual airfoil was thinner than the true airfoil, the dotted line would be above thehorizontal axis and vice versa. All models had a nominal chord of 12 in, and the Reynolds number wasbased on a 12-in chord.

22Sum

mary

ofLow

-SpeedA

irfoilData

Table 4.1: Test Matrix and Run Number Index

Model Configuration

(Builder)

Designer Fig. p. Fig. p. Re Run # Fig. p. Re Run #

AG12 Clean 4.1 30 4.3 31 40,000 BB05707 4.4 32 40,000 BB05706

(M. Drela, et al.) 4.2 60,000 BB05637 60,000 JB05636

Drela 80,000 TS05641 4.4 33 80,000 TS05640

100,000 BB05639 100,000 BB05638

150,000 JB05643 4.4 34 150,000 BM05642

200,000 JB05645 200,000 JB05644

300,000 JB05647 4.4 35 300,000 BB05646

AG16 Clean 4.5 36 4.7 37 40,000 BB05695 4.8 38 40,000 BB05694

(M. Drela, et al.) 4.6 60,000 TS05699 60,000 TS05698

Drela 80,000 JB05687 4.8 39 80,000 JB05686

100,000 TS05689 100,000 TS05688

150,000 BM05691 4.8 40 150,000 BM05690

200,000 BM05693 200,000 BM05692

300,000 JB05697 4.8 41 300,000 JB05696

AG24 Clean 4.9 42 4.11 43 60,000 BB05531 4.12 44 60,000 BB05530

(M. Drela, et al.) 4.10 80,000 BM05533 80,000 BM05532

Drela 100,000 BM05535 4.12 45 100,000 BM05534

150,000 BB05537 150,000 BB05536

200,000 BM05539 4.12 46 200,000 BM05538

300,000 BM05541 300,000 BM05540

400,000 BM05543 4.12 47 400,000 BM05542

AG35-r Clean 4.13 48 4.15 49 60,000 BM05388 4.16 50 60,000 BM05387

(M. Drela, et al.) 4.14 80,000 BB05396 80,000 BM05395

Drela 100,000 BB05391 4.16 51 100,000 BB05389

150,000 BM05393/BM05394 150,000 BM05392

200,000 BM05398/BM05399 4.16 52 200,000 BM05397

300,000 BB05402 300,000 BM05401

AG40d-02r Clean 4.17 54 4.19 55 60,000 TS05712 4.20 56 60,000 JB05711

(M. Drela, et al.) 0 deg flap 4.18 80,000 BM05714 80,000 TS05713

Drela 100,000 BM05716 4.20 57 100,000 BM05715

150,000 BB05718 150,000 BB05717

200,000 BB05720 4.20 58 200,000 BB05719

300,000 JB05722 300,000 JB05721

(continues) 500,000 JB05724 4.20 59 500,000 JB05723

Lift & Moment DataV-dist &

ProfileDrag Data

Chapter4:

AirfoilProfiles

andPerform

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23

Table 4.1: Continued

AG40d-02r Clean 4.21 60 4.22 61 60,000 JB05735 4.23 62 60,000 JB05734

(continued) -2 deg flap 80,000 JB05749 80,000 JB05748

100,000 BB05731 4.23 63 100,000 BB05730

150,000 JB05751 150,000 JB05750

200,000 BB05733 4.23 64 200,000 BB05732

300,000 JB05730 300,000 JB05729

450,000 JB05737 4.23 65 450,000 JB05736

Clean 4.24 66 4.25 67 60,000 TS05740 4.26 68 60,000 TS05741

2 deg flap 80,000 BB05823 80,000 BB05822

100,000 TS05743 4.26 69 100,000 TS05742

150,000 JB05821 150,000 JB05820

200,000 BB05745 4.26 70 200,000 BB05744

300,000 BB05747 300,000 BB05746

Clean 4.27 72 4.28 73 60,000 TS05753 4.29 74 60,000 TS05752

4 deg flap 80,000 BB05830 80,000 BB05829

100,000 BM05755 4.29 75 100,000 BM05754

150,000 BB05828 150,000 BB05827

200,000 BB05757 4.29 76 200,000 BB05756

300,000 JB05759 300,000 BB05758

Clean 4.30 78 4.31 79 100,000 JB05762 4.32 80 100,000 JB05761

-15 deg flap

Clean 4.33 82 4.34 83 100,000 BB05779 4.35 84 100,000 BB05778

-10 deg flap

Clean 4.36 86 4.37 87 100,000 BB05767 4.38 88 100,000 BB05766

-5 deg flap

Clean 4.39 90 4.40 91 100,000 BM05770 4.41 92 100,000 BM05769

5 deg flap

Clean 4.42 94 4.43 95 100,000 JB05773 4.44 96 100,000 JB05772

10 deg flap

Clean 4.45 98 4.46 99 100,000 TS05776 4.47 100 100,000 TS05775

15 deg flap

Clean 4.48 102 4.49 103 40,000 BB05825 4.50 104 40,000 BB05824

20 deg flap

Clean 4.51 106 4.52 107 40,000 BB05826

(continues) 30 deg flap

24Sum

mary

ofLow

-SpeedA

irfoilData


AG40d-02r Clean, gap sealed 4.53 109 100,000 BM05815 4.54 110 100,000 BM05814

(continued) 0 deg flap 200,000 TS05799 200,000 TS05798

300,000 BB05801 4.54 111 300,000 BB05800

Clean, gap sealed 4.55 113 100,000 JB05819 4.56 114 100,000 JB05818

-2 deg flap 200,000 JB05795 200,000 JB05794

300,000 BM05817 300,000 BM05816

500,000 TS05797 4.56 115 500,000 TS05796

Clean, gap sealed 4.57 117 60,000 BB05813 4.58 118 60,000 BB05812

4 deg flap 100,000 JB05809 100,000 JB05808

200,000 JB05811 4.58 119 200,000 JB05810

Aileron Response 4.59 121 100,000

AG455ct-02r Clean 4.60 122 4.62 123 60,000 BM05458 4.63 124 60,000 BM05456

(M. Drela, et al.) -0.4 deg flap 4.61 80,000 BB05461 80,000 BM05459

Drela 100,000 BB05462 4.63 125 100,000 BM05460

150,000 BM05464 150,000 BM05463

200,000 BM05466 4.63 126 200,000 BM05465

300,000 BB05468 300,000 BB05467

Clean 4.64 128 4.65 129 60,000 BM05445 4.66 130 60,000 BM05444

-2.4 deg flap 80,000 BM05447 80,000 BM05446

100,000 BM05449 4.66 131 100,000 BB05448

150,000 BM05451 150,000 BM05450

200,000 BB05453 4.66 132 200,000 BB05452

300,000 BM05455 300,000 BB05454

450,000 BM05529 4.66 133 450,000 BM05528

Clean 4.67 134 4.68 135 40,000 BM05471 4.69 136 40,000 BB05469

1.6 deg flap 60,000 BM05472 60,000 BB05470

80,000 BM05474 4.69 137 80,000 BM05473

100,000 BB05476 100,000 BB05475

150,000 BM05478 4.69 138 150,000 BM05477

Clean 4.70 140 4.71 141 40,000 BM05480 4.72 142 40,000 BM05479

3.6 deg flap 60,000 BB05482 60,000 BB05481

80,000 BB05484 4.72 143 80,000 BB05483

100,000 BM05486 100,000 BM05485

(continues) 150,000 BM05488 4.72 144 150,000 BM05487

Chapter4:

AirfoilProfiles

andPerform

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25


AG455ct-02r Clean 4.73 146 4.74 147 60,000 BB05502 4.75 148 60,000 BM05501

(continued) -15.4 deg flap 100,000 BM05504 100,000 BM05503

Clean 4.76 150 4.77 151 60,000 BM05506 4.78 152 60,000 BM05505

-10.4 deg flap 100,000 BB05508 100,000 BM05507

Clean 4.79 154 4.80 155 60,000 BB05511/BM05513 4.81 156 60,000 BB05510

-5.4 deg flap 100,000 BB05515 100,000 BB05514

Clean 4.82 158 4.83 159 60,000 BM05517 4.84 160 60,000 BM05516

4.6 deg flap 100,000 BM05519 100,000 BM05518

Clean 4.85 162 4.86 163 60,000 BM05521 4.87 164 60,000 BM05520

9.6 deg flap 100,000 BB05523 100,000 BB05522

Clean 4.88 166 4.89 167 60,000 BB05525 4.90 168 60,000 BB05524

14.6 deg flap 100,000 BM05527 100,000 BM05526

Clean, gap sealed 4.91 169 60,000 BM05494 4.92 170 60,000 BM05493

-0.4 deg flap 100,000 BM05496 100,000 BM05495

Clean, gap sealed 4.93 171 100,000 BB05498 4.94 172 100,000 BB05497

-2.4 deg flap 300,000 BM05500 300,000 BB05499

Clean, gap sealed 4.95 173 40,000 BB05490 4.96 174 40,000 BM05489

3.6 deg flap 60,000 BM05492 60,000 BB05491

Aileron Response 4.97 175 60,000

4.98 176 100,000

CAL1215j Clean 4.99 178 4.101 179 100,000 BM05557 4.102 180 100,000 BM05556

(T. Lampe) 4.100 200,000 BM05562 200,000 BM05558

Lyon 300,000 BB05564 4.102 181 300,000 BB05563

400,000 BM05566 400,000 BM05565

500,000 BM05568 4.102 182 500,000 BM05567

CAL2263m Clean 4.103 184 4.105 185 60,000 BB05416/BM05417 4.106 186 60,000 BB05415

(C. Greaves) 4.104 100,000 BM05439 100,000 BM05438

Lyon 200,000 BB05422 4.106 187 200,000 BB05421

300,000 BB05424 300,000 BB05423

400,000 BB05428 4.106 188 400,000 BB05427

500,000 BB05441 500,000 BB05440

CAL4014l Clean 4.107 190 4.109 191 100,000 BB05545 4.110 192 100,000 BB05544

(J. Thomas) 4.108 200,000 BM05547/BM05548 200,000 BB05546

Lyon 300,000 BM05550 4.110 193 300,000 BM05549

400,000 BB05555 400,000 BB05554

500,000 BB05553 4.110 194 500,000 BB05552

26Sum

mary

ofLow

-SpeedA

irfoilData


E387 (E) Clean 4.111 196 4.113 197 60,000 BB05385 4.114 198 100,000 BM05372

(Y. Tinel) 4.112 100,000 BM05374 200,000 BM05373

Eppler 200,000 BM05376 4.114 199 300,000 BM05378

300,000 BB05381 400,000 BM05379

460,000 BM05384 4.114 200 500,000 BM05383

Flat Plate Baseline 4.115 202 4.116 203 40,000 RD06131

(C. Goudeseune

&

M. Goudeseune)

60,000 JB06129

4.116 204 80,000 RD06132

100,000 RD06133

4.116 205 120,000 KT06135

Serrations A 4.117 206 4.118 207 40,000 JB06153

60,000 JB06152

4.118 208 80,000 JB06151

100,000 JB06154

4.118 209 120,000 JB06155

Serrations B 4.119 210 4.120 211 40,000 PG06168

60,000 RD06163

4.120 212 80,000 RD06164

100,000 PG06165

4.120 213 120,000 PG06166

Serrations C 4.121 214 4.122 215 40,000 RD06157

60,000 RD06158

4.122 216 80,000 RD06159

100,000 RD06160

4.122 217 120,000 RD06161

Serrations D 4.123 218 4.124 219 40,000 KT06142

60,000 KT06145

4.124 220 80,000 KT06144

100,000 KT06146

4.124 221 120,000 RD06147

Square Wave 4.125 222 4.126 223 40,000 PG06169

60,000 PG06170

4.126 224 80,000 PG06171

100,000 PG06172

(continues) 4.126 225 120,000 PG06173

Chapter4:

AirfoilProfiles

andPerform

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27


Flat Plate Small Holes 4.127 226 4.128 227 40,000 KT06182

(continued) 60,000 KT06183

4.128 228 100,000 KT06184

Large Holes 4.129 230 4.130 231 60,000 RD06175

100,000 RD06176

Small Cubes 4.131 232 4.132 233 40,000 KT06186

60,000 KT06187

4.132 234 100,000 KT06188

Large Cubes 4.133 236 4.134 237 40,000 PG06179

60,000 PG06178

4.134 238 100,000 PG06180

MA409 Clean 4.135 240 4.137 241 40,000 06269RD 4.138 242 40,000 06268RD

(R. Cooney) 4.136 60,000 06265RD 60,000 06264RD

Achterberg 100,000 06267RD 4.138 243 100,000 06266RD

200,000 06271RD 200,000 06270RD

300,000 06274RD 4.138 244 300,000 06273RD

NACA 43012A Clean 4.139 246 4.141 247 60,000 BB05887 4.142 248 60,000 BB05886

(M. Nankivil) 4.140 100,000 JB05890 100,000 JB05889

200,000 TS05892/BM05893 4.142 249 200,000 TS05891

300,000 BB05897 300,000 BB05896

400,000 JB05899 4.142 250 400,000 JB05898

500,000 JB05901 500,000 JB05900

S1223 Clean 4.143 252 4.145 253 80,000 RD06080

(Y. Tinel) 4.144 100,000 RD06081

Selig 4.145 254 120,000 RD06082

140,000 PG06083

4.145 255 160,000 PG06084

180,000 PG06085

4.145 256 200,000 PG06086

250,000 KT06087

Gurney flap 4.146 257 160,000 PG06121

h/c = 4.17% 180,000 PG06120

4.146 258 200,000 JB06118

(continues) 250,000 JB06119

28Sum

mary

ofLow

-SpeedA

irfoilData


S1223 Gurney flap 4.147 259 140,000 PG06113

(continued) h/c = 3.12% 160,000 JB06117

4.147 260 180,000 JB06116

200,000 JB06114

4.147 261 250,000 JB06115

Gurney flap 4.148 263 160,000 KT06099

h/c = 2.08% 180,000 KT06098

4.148 264 200,000 JB06097

250,000 JB06096

Gurney flap 4.149 265 160,000 RD06092

h/c = 1.56% 180,000 RD06093

4.149 266 200,000 RD06094

250,000 JB06095

Gurney flap 4.150 267 160,000 KT06091

h/c = 1.04% 180,000 KT06090

4.150 268 200,000 KT06089

250,000 PG06088

u.s.t. 4.151 269 160,000 06557RD

t/c = 0.11% 180,000 06558RD

u.s.t. 4.152 271 160,000 06555RD

t/c = 0.19% 180,000 06556RD

S8064 Clean 4.153 272 4.155 273 100,000 JB05614 4.156 274 100,000 BB05616

(T. Lampe) 4.154 200,000 BB05617 200,000 JB05615

Selig 300,000 TS05619 4.156 275 300,000 BB05618

400,000 BM05622 400,000 BM05621

500,000 BM05624 4.156 276 500,000 BM05623

S9000 Clean 4.157 278 4.159 279 60,000 BM05781 4.160 280 60,000 BM05780

(T. Akers) 0 deg flap 4.158 100,000 JB05784 100,000 JB05783

Selig 200,000 TS05786 4.160 281 200,000 TS05785

300,000 BB05788 300,000 BB05787

400,000 BM05790/BM05791 4.160 282 400,000 BB05789

(continues) 500,000 BM05793 500,000 BM05792

Chapter4:

AirfoilProfiles

andPerform

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29


S9000 Clean 4.161 284 4.162 285 60,000 JB05840 4.163 286 60,000 JB05839

(continued) 2.5 deg flap 100,000 TS05842/BB05843 100,000 TS05841

200,000 JB05838 4.163 287 200,000 BM05837

300,000 BB05846 300,000 BB05845

400,000 JB05848 4.163 288 400,000 BB05847

500,000 JB05850 500,000 JB05849

Clean 4.164 290 4.165 291 60,000 TS05853/TS05854 4.166 292 60,000 JB05851

5 deg flap 100,000 BM05856 100,000 BM05855

200,000 BB05858 4.166 293 200,000 BM05857

300,000 BB05860 300,000 BB05859

400,000 TS05881 4.166 294 400,000 BM05880

500,000 BM05884 500,000 BM05883

30Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG12

Avg. Difference = 0.0041 in.

Fig.4.1:Com

parisonbetw

eenthe

trueand

actualAG

12.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

−0.000.230.460.690.921.14

α0l

= −2.0

x/c

V/V∞

AG12

Fig.4.2:Inviscidvelocity

distributionsforthe

AG

12.

AG

12

Chapter4:

AirfoilProfiles

andPerform

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31

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000

AG12 (M. Drela, et al.)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.3:Drag

polarfortheA

G12.

AG

12

32Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.4:Liftand

mom

entcharacteristicsforthe

AG

12.

AG

12

Chapter4:

AirfoilProfiles

andPerform

ancePlots

33

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.4:Continued.

AG

12

34Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.4:Continued.

AG

12

Chapter4:

AirfoilProfiles

andPerform

ancePlots

35

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.4:Continued.

AG

12

36Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG16


Fig.4.5:Com

parisonbetw

eenthe

trueand

actualAG

16.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

0.010.240.470.710.941.16

α0l

= −2.1

x/c

V/V∞

AG16


distributionsforthe

AG

16.

AG

16

Chapter4:

AirfoilProfiles

andPerform

ancePlots

37

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.7:Drag

polarfortheA

G16.

AG

16

38Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.8:Liftand

mom


AG

16.

AG

16

Chapter4:

AirfoilProfiles

andPerform

ancePlots

39

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.8:Continued.

AG

16

40Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.8:Continued.

AG

16

Chapter4:

AirfoilProfiles

andPerform

ancePlots

41

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.8:Continued.

AG

16

42Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG24


Fig.4.9:Com

parisonbetw

eenthe

trueand

actualAG

24.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

0.070.310.540.771.001.24

α0l

= −2.6

x/c

V/V∞

AG24


distributionsforthe

AG

24.

AG

24

Chapter4:

AirfoilProfiles

andPerform

ancePlots

43

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000Re = 150,000

Re = 200,000Re = 300,000Re = 400,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.11:Drag

polarfortheA

G24.

AG

24

44Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.12:Liftand

mom


AG

24.

AG

24

Chapter4:

AirfoilProfiles

andPerform

ancePlots

45

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.12:Continued.

AG

24

46Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.12:Continued.

AG

24

Chapter4:

AirfoilProfiles

andPerform

ancePlots

47

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.12:Continued.

AG

24

48Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG35−r (AG35 rotated −1.563 deg)


Fig.4.13:Com

parisonbetw

eenthe

trueand

actualAG

35-r.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

0.010.250.490.720.951.19

α0l

= −2.1

x/c

V/V∞

AG35−r


distributionsforthe

AG

35-r.

AG

35-r

Chapter4:

AirfoilProfiles

andPerform

ancePlots

49

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000

AG35−r (M. Drela, et al.)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.15:Drag

polarfortheA

G35-r.

AG

35-r

50Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.16:Liftand

mom


AG

35-r.

AG

35-r

Chapter4:

AirfoilProfiles

andPerform

ancePlots

51

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.16:Continued.

AG

35-r

52Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.16:Continued.

AG

35-r

Chapter4:

AirfoilProfiles

andPerform

ancePlots

53

54Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG40d−02r


Fig.4.17:Com

parisonbetw

eenthe

trueand

actualAG

40d-02r.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

0.120.360.590.821.051.28

α0l

= −3.1

x/c

V/V∞

AG40d−02r Flap 0 deg


distributionsforthe

AG

40d-02r.

AG

40d-02rFlap

0deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

55

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000Re = 150,000

Re = 200,000Re = 300,000Re = 450,000

AG40d−02r (M. Drela, et al.)Flap = 0 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.19:Drag

polarfortheA

G40d-02r.

AG

40d-02rFlap

0deg

cf /c

=25%

56Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.20:Liftand

mom


AG

40d-02r.

AG

40d-02rFlap

0deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

57

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.20:Continued.

AG

40d-02rFlap

0deg

cf /c

=25%

58Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.20:Continued.

AG

40d-02rFlap

0deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

59

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 450,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.20:Continued.

AG

40d-02rFlap

0deg

cf /c

=25%

60Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

−0.020.210.450.680.911.14

α0l

= −1.8

x/c

V/V∞

AG40d−02r Flap −2 deg


distributionsforthe

AG

40d-02rwith

a−

2deg

flap.

AG

40d-02rFlap

−2

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

61

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000Re = 150,000

Re = 200,000Re = 300,000Re = 500,000

AG40d−02r (M. Drela, et al.)Flap = −2 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.22:Drag

polarfortheA

G40d-02rw

itha−

2deg

flap.

AG

40d-02rFlap

−2

degcf /c

=25%

62Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.23:Liftand

mom


AG

40d-02rwith

a−

2deg

flap.

AG

40d-02rFlap

−2

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

63

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.23:Continued.

AG

40d-02rFlap

−2

degcf /c

=25%

64Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.23:Continued.

AG

40d-02rFlap

−2

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

65

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.23:Continued.

AG

40d-02rFlap

−2

degcf /c

=25%

66Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

0.260.500.730.961.191.42

α0l

= −4.3

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a2

degflap.

AG

40d-02rFlap

2deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

67

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.25:Drag

polarfortheA

G40d-02rw

itha

2deg

flap.

AG

40d-02rFlap

2deg

cf /c

=25%

68Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.26:Liftand

mom


AG

40d-02rwith

a2

degflap.

AG

40d-02rFlap

2deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

69

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.26:Continued.

AG

40d-02rFlap

2deg

cf /c

=25%

70Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.26:Continued.

AG

40d-02rFlap

2deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

71

72Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0

0.410.640.871.101.33

α0l

= −5.5

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a4

degflap.

AG

40d-02rFlap

4deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

73

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.28:Drag

polarfortheA

G40d-02rw

itha

4deg

flap.

AG

40d-02rFlap

4deg

cf /c

=25%

74Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.29:Liftand

mom


AG

40d-02rwith

a4

degflap.

AG

40d-02rFlap

4deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

75

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.29:Continued.

AG

40d-02rFlap

4deg

cf /c

=25%

76Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.29:Continued.

AG

40d-02rFlap

4deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

77

78Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

2.0 4.0 6.0 8.010.0

−0.47−0.24−0.010.220.46

α0l

= 6.1

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a−

15deg

flap.

AG

40d-02rFlap

−15

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

79

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.31:Drag

polarfortheA

G40d-02rw

itha−

15deg

flap.

AG

40d-02rFlap

−15

degcf /c

=25%

80Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.32:Liftand

mom


AG

40d-02rwith

a−

15deg

flap.

AG

40d-02rFlap

−15

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

81

82Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

0.0 2.0 4.0 6.0 8.0

−0.35−0.120.110.340.58

α0l

= 3.0

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a−

10deg

flap.

AG

40d-02rFlap

−10

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

83

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.34:Drag

polarfortheA

G40d-02rw

itha−

10deg

flap.

AG

40d-02rFlap

−10

degcf /c

=25%

84Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.35:Liftand

mom


AG

40d-02rwith

a−

10deg

flap.

AG

40d-02rFlap

−10

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

85

86Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.0

−0.23−0.000.230.470.700.93

α0l

= 0.0

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a−

5deg

flap.

AG

40d-02rFlap

−5

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

87

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.37:Drag

polarfortheA

G40d-02rw

itha−

5deg

flap.

AG

40d-02rFlap

−5

degcf /c

=25%

88Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.38:Liftand

mom


AG

40d-02rwith

a−

5deg

flap.

AG

40d-02rFlap

−5

degcf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

89

90Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0

0.240.480.710.941.171.40

α0l

= −6.1

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a5

degflap.

AG

40d-02rFlap

5deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

91

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.40:Drag

polarfortheA

G40d-02rw

itha

5deg

flap.

AG

40d-02rFlap

5deg

cf /c

=25%

92Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.41:Liftand

mom


AG

40d-02rwith

a5

degflap.

AG

40d-02rFlap

5deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

93

94Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−6.0−4.0−2.0 0.0 2.0 4.0

0.370.600.831.061.291.52

α0l

= −9.2

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a10

degflap.

AG

40d-02rFlap

10deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

95

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.43:Drag

polarfortheA

G40d-02rw

itha

10deg

flap.

AG

40d-02rFlap

10deg

cf /c

=25%

96Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.44:Liftand

mom


AG

40d-02rwith

a10

degflap.

AG

40d-02rFlap

10deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

97

98Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−6.0−4.0−2.0 0.0 2.0 4.0

0.720.951.181.411.641.87

α0l

= −12.3

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a15

degflap.

AG

40d-02rFlap

15deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

99

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.46:Drag

polarfortheA

G40d-02rw

itha

15deg

flap.

AG

40d-02rFlap

15deg

cf /c

=25%

100Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.47:Liftand

mom


AG

40d-02rwith

a15

degflap.

AG

40d-02rFlap

15deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

101

102Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−8.0−6.0−4.0−2.0 0.0 2.0

0.841.071.301.531.761.98

α0l

= −15.4

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a20

degflap.

AG

40d-02rFlap

20deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

103

0.00 0.02 0.04 0.06 0.08 0.10 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.49:Drag

polarfortheA

G40d-02rw

itha

20deg

flap.

AG

40d-02rFlap

20deg

cf /c

=25%

104Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.50:Liftand

mom


AG

40d-02rwith

a20

degflap.

AG

40d-02rFlap

20deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

105

106Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−8.0−6.0−4.0−2.0 0.0 2.0

1.541.761.992.212.432.64

α0l

= −21.9

x/c

V/V∞



distributionsforthe

AG

40d-02rwith

a30

degflap.

AG

40d-02rFlap

30deg

cf /c

=25%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

107

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.52:Liftand

mom


AG

40d-02rwith

a30

degflap.

AG

40d-02rFlap

30deg

cf /c

=25%

108Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

109

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000

Re = 200,000

AG40d−02r (M. Drela, et al.)Flap = 0 deggap sealed

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.53:Drag

polarforthegap

sealedA

G40d-02r.

AG

40d-02rFlap

0deg

cf /c

=25%

gapsealed

110Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.54:Liftand

mom


gapsealed

AG

40d-02r.

AG

40d-02rFlap

0deg

cf /c

=25%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

111

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.54:Continued.

AG

40d-02rFlap

0deg

cf /c

=25%

gapsealed

112Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

113

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000Re = 200,000

Re = 300,000Re = 500,000

AG40d−02r (M. Drela, et al.)Flap = −2 deggap sealed

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.55:Drag

polarforthegap

sealedA

G40d-02rw

itha−

2deg

flap.

AG

40d-02rFlap

−2

degcf /c

=25%

gapsealed

114Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.56:Liftand

mom


gapsealed

AG

40d-02rwith

a−

2deg

flap.

AG

40d-02rFlap

−2

degcf /c

=25%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

115

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.56:Continued.

AG

40d-02rFlap

−2

degcf /c

=25%

gapsealed

116Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

117

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000

Re = 200,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.57:Drag

polarforthegap

sealedA

G40d-02rw

itha

4deg

flap.

AG

40d-02rFlap

4deg

cf /c

=25%

gapsealed

118Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.58:Liftand

mom


gapsealed

AG

40d-02rwith

a4

degflap.

AG

40d-02rFlap

4deg

cf /c

=25%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

119

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.58:Continued.

AG

40d-02rFlap

4deg

cf /c

=25%

gapsealed

120Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

121

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl

−15 deg flap

−10 deg flap

−5 deg flap

0 deg flap

5 deg flap

10 deg flap

15 deg flap

AG40d−02raileron responseRe = 100,000

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.59:Aileron

Response

fortheA

G40d-02rat

Re

=100,000.

AG

40d-02rA

ileronR

esponseRe

=100,000

cf /c

=25%

122Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

AG455ct−02r


Fig.4.60:Com

parisonbetw

eenthe

trueand

actualAG

455ct-02r.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0 8.0

−0.150.080.310.540.771.001.23

α0l

= −2.7

x/c

V/V∞

AG455ct−02r Flap −0.4 deg


distributionsforthe

AG

455ct-02rwith

a−

0.4deg

flap.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

123

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000

Re = 150,000Re = 200,000Re = 300,000

AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.62:Drag

polarfortheA

G455ct-02rw

itha−

0.4deg

flap.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

124Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.63:Liftand

mom


AG

455ct-02rwith

a−

0.4deg

flap.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

125

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.63:Continued.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

126Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.63:Continued.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

127

128Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0 8.0

−0.30−0.070.160.390.620.851.08

α0l

= −1.4

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a−

2.4deg

flap.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

129

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 80,000Re = 100,000Re = 150,000

Re = 200,000Re = 300,000Re = 450,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.65:Drag

polarfortheA

G455ct-02rw

itha−

2.4deg

flap.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

130Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.66:Liftand

mom


AG

455ct-02rwith

a−

2.4deg

flap.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

131

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.66:Continued.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

132Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.66:Continued.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

133

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 450,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.66:Continued.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

134Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0 8.0

0.000.230.460.690.921.151.38

α0l

= −4.0

x/c

V/V∞

AG455ct−02r Flap 1.6 deg


distributionsforthe

AG

455ct-02rwith

a1.6

degflap.

AG

455ct-02rFlap

1.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

135

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000Re = 60,000Re = 80,000

Re = 100,000Re = 150,000

AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.68:Drag

polarfortheA

G455ct-02rw

itha

1.6deg

flap.

AG

455ct-02rFlap

1.6deg

cf /c

=30%

136Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.69:Liftand

mom


AG

455ct-02rwith

a1.6

degflap.

AG

455ct-02rFlap

1.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

137

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.69:Continued.

AG

455ct-02rFlap

1.6deg

cf /c

=30%

138Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.69:Continued.

AG

455ct-02rFlap

1.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

139

140Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0 8.0

0.150.390.620.851.071.301.53

α0l

= −5.4

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a3.6

degflap.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

141

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000Re = 60,000Re = 80,000

Re = 100,000Re = 150,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.71:Drag

polarfortheA

G455ct-02rw

itha

3.6deg

flap.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

142Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.72:Liftand

mom


AG

455ct-02rwith

a3.6

degflap.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

143

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.72:Continued.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

144Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 150,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.72:Continued.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

145

146Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

2.0 4.0 6.0 8.010.0

−0.60−0.37−0.140.090.32

α0l

= 7.2

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a−

15.4deg

flap.

AG

455ct-02rFlap

−15.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

147

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.74:Drag

polarfortheA

G455ct-02rw

itha−

15.4deg

flap.

AG

455ct-02rFlap

−15.4

degcf /c

=30%

148Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.75:Liftand

mom


AG

455ct-02rwith

a−

15.4deg

flap.

AG

455ct-02rFlap

−15.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

149

150Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

0.0 2.0 4.0 6.0 8.0

−0.45−0.220.010.240.47

α0l

= 3.9

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a−

10.4deg

flap.

AG

455ct-02rFlap

−10.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

151

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.77:Drag

polarfortheA

G455ct-02rw

itha−

10.4deg

flap.

AG

455ct-02rFlap

−10.4

degcf /c

=30%

152Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.78:Liftand

mom


AG

455ct-02rwith

a−

10.4deg

flap.

AG

455ct-02rFlap

−10.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

153

154Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0

−0.30−0.070.160.390.62

α0l

= 0.6

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a−

5.4deg

flap.

AG

455ct-02rFlap

−5.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

155

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.80:Drag

polarfortheA

G455ct-02rw

itha−

5.4deg

flap.

AG

455ct-02rFlap

−5.4

degcf /c

=30%

156Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.81:Liftand

mom


AG

455ct-02rwith

a−

5.4deg

flap.

AG

455ct-02rFlap

−5.4

degcf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

157

158Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−4.0−2.0 0.0 2.0 4.0 6.0

0.230.460.690.921.151.38

α0l

= −6.0

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a4.6

degflap.

AG

455ct-02rFlap

4.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

159

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.83:Drag

polarfortheA

G455ct-02rw

itha

4.6deg

flap.

AG

455ct-02rFlap

4.6deg

cf /c

=30%

160Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.84:Liftand

mom


AG

455ct-02rwith

a4.6

degflap.

AG

455ct-02rFlap

4.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

161

162Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−6.0−4.0−2.0 0.0 2.0 4.0

0.380.610.841.071.301.52

α0l

= −9.3

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a9.6

degflap.

AG

455ct-02rFlap

9.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

163

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.86:Drag

polarfortheA

G455ct-02rw

itha

9.6deg

flap.

AG

455ct-02rFlap

9.6deg

cf /c

=30%

164Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.87:Liftand

mom


AG

455ct-02rwith

a9.6

degflap.

AG

455ct-02rFlap

9.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

165

166Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−6.0−4.0−2.0 0.0 2.0 4.0

0.760.991.221.441.671.89

α0l

= −12.7

x/c

V/V∞



distributionsforthe

AG

455ct-02rwith

a14.6

degflap.

AG

455ct-02rFlap

14.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

167

0.00 0.02 0.04 0.06 0.08 0.10 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.89:Drag

polarfortheA

G455ct-02rw

itha

14.6deg

flap.

AG

455ct-02rFlap

14.6deg

cf /c

=30%

168Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.90:Liftand

mom


AG

455ct-02rwith

a14.6

degflap.

AG

455ct-02rFlap

14.6deg

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

169

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000 Re = 100,000

AG455ct−02r (M. Drela, et al.)Flap = −0.4 deggap sealed

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.91:Drag

polarforthegap

sealedA

G455ct-02rw

itha−

0.4deg

flap.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

gapsealed

170Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.92:Liftand

mom


gapsealed

AG

455ct-02rwith

a−

0.4deg

flap.

AG

455ct-02rFlap

−0.4

degcf /c

=30%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

171

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000 Re = 300,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.93:Drag

polarforthegap

sealedA

G455ct-02rw

itha−

2.4deg

flap.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

gapsealed

172Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.94:Liftand

mom


gapsealed

AG

455ct-02rwith

a−

2.4deg

flap.

AG

455ct-02rFlap

−2.4

degcf /c

=30%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

173

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000 Re = 60,000

AG455ct−02r (M. Drela, et al.)Flap = 3.6 deggap sealed

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.95:Drag

polarforthegap

sealedA

G455ct-02rw

itha

3.6deg

flap.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

gapsealed

174Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.96:Liftand

mom


gapsealed

AG

455ct-02rwith

a3.6

degflap.

AG

455ct-02rFlap

3.6deg

cf /c

=30%

gapsealed

Chapter4:

AirfoilProfiles

andPerform

ancePlots

175

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl

−15 deg flap

−10 deg flap

−5 deg flap

0 deg flap

5 deg flap

10 deg flap

15 deg flap

AG455ct−02raileron responseRe = 60,000

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.97:Aileron

Response

fortheA

G455ct-02rat

Re

=60,000.

AG

455ct-02rA

ileronR

esponseRe

=60,000

cf /c

=30%

176Sum

mary

ofLow

-SpeedA

irfoilData

0.00 0.01 0.02 0.03 0.04 0.05−0.5

0.0

0.5

1.0

1.5

Cd

Cl

−15 deg flap

−10 deg flap

−5 deg flap

0 deg flap

5 deg flap

10 deg flap

15 deg flap

AG455ct−02raileron responseRe = 100,000

−10 0 10 20−0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.98:Aileron

Response

fortheA

G455ct-02rat

Re

=100,000.

AG

455ct-02rA

ileronR

esponseRe

=100,000

cf /c

=30%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

177

178Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

CAL1215j


Fig.4.99:Com

parisonbetw

eenthe

trueand

actualCA

L1215j.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.010.250.490.730.971.211.44

α0l

= −2.1

x/c

V/V∞

CAL1215j


distributionsforthe

CA

L1215j.

CA

L1215j

Chapter4:

AirfoilProfiles

andPerform

ancePlots

179

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000Re = 200,000Re = 300,000

Re = 400,000Re = 500,000

CAL1215j (T. Lampe)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.101:Drag

polarfortheC

AL

1215j.

CA

L1215j

180Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

CAL1215j (T. Lampe)

Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

CAL1215j (T. Lampe)

Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.102:Liftand

mom


CA

L1215j.

CA

L1215j

Chapter4:

AirfoilProfiles

andPerform

ancePlots

181

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

CAL1215j (T. Lampe)

Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

CAL1215j (T. Lampe)

Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.102:Continued.

CA

L1215j

182Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

CAL1215j (T. Lampe)

Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


CA

L1215j

Chapter4:

AirfoilProfiles

andPerform

ancePlots

183

184Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

CAL2263m


Fig.4.103:Com

parisonbetw

eenthe

trueand

actualCA

L2263m

.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.190.430.670.911.151.391.62

α0l

= −3.6

x/c

V/V∞

CAL2263m


distributionsforthe

CA

L2263m

.

CA

L2263m

Chapter4:

AirfoilProfiles

andPerform

ancePlots

185

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 400,000Re = 500,000

CAL2263m (C. Greaves)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.105:Drag

polarfortheC

AL

2263m.

CA

L2263m

186Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.106:Liftand

mom


CA

L2263m

.

CA

L2263m

Chapter4:

AirfoilProfiles

andPerform

ancePlots

187

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


CA

L2263m

188Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


CA

L2263m

Chapter4:

AirfoilProfiles

andPerform

ancePlots

189

190Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

CAL4014l


Fig.4.107:Com

parisonbetw

eenthe

trueand

actualCA

L4014l.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

0.0 2.0 4.0 6.0 8.010.012.0

0.010.250.480.720.951.181.42

α0l

= −0.1

x/c

V/V∞

CAL4014l


distributionsforthe

CA

L4014l.

CA

L4014l

Chapter4:

AirfoilProfiles

andPerform

ancePlots

191

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000Re = 200,000Re = 300,000

Re = 400,000Re = 500,000

CAL4014l (J. Thomas)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.109:Drag

polarfortheC

AL

4014l.

CA

L4014l

192Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.110:Liftand

mom


CA

L4014l.

CA

L4014l

Chapter4:

AirfoilProfiles

andPerform

ancePlots

193

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


CA

L4014l

194Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


CA

L4014l

Chapter4:

AirfoilProfiles

andPerform

ancePlots

195

196Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

E387 (E)


Fig.4.111:Com

parisonbetw

eenthe

trueand

actualE387

(E).

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.180.420.650.881.121.351.58

α0l

= −3.6

x/c

V/V∞

E387


distributionsforthe

E387

(E).

E387

(E)

Chapter4:

AirfoilProfiles

andPerform

ancePlots

197

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 460,000

E387 (E) (Y. Tinel)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.113:Drag

polarfortheE

387(E

).

E387

(E)

198Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 (E) (Y. Tinel)

Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 (E) (Y. Tinel)

Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.114:Liftand

mom


E387

(E).

E387

(E)

Chapter4:

AirfoilProfiles

andPerform

ancePlots

199

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 (E) (Y. Tinel)

Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 (E) (Y. Tinel)

Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


E387

(E)

200Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

E387 (E) (Y. Tinel)

Re = 460,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


E387

(E)

Chapter4:

AirfoilProfiles

andPerform

ancePlots

201

202Sum

mary

ofLow

-SpeedA

irfoilData

Baseline

Fig.4.115:Schematic

ofthebaseline

leadingedge

configuration.

FlatPlateB

aseline

Chapter4:

AirfoilProfiles

andPerform

ancePlots

203

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickBaseline

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.116:Liftand

mom

entcharacteristicsfora

flatplatew

iththe

baselineleading

edge.

FlatPlateB

aseline

204Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateB

aseline

Chapter4:

AirfoilProfiles

andPerform

ancePlots

205

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateB

aseline

206Sum

mary

ofLow

-SpeedA

irfoilData

Serrations A

Amplitude = 0.013c

Wavelength = 0.052c

Fig.4.117:Schematic

oftheleading

edgeserrations

(Case

A)configuration.

FlatPlateSerrations

A

Chapter4:

AirfoilProfiles

andPerform

ancePlots

207

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations A

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.118:Liftand

mom

entcoefficientcharacteristicsfora

flatplatew

ithleading

edgeserrations

(Case

A).

FlatPlateSerrations

A

208Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

A

Chapter4:

AirfoilProfiles

andPerform

ancePlots

209

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

A

210Sum

mary

ofLow

-SpeedA

irfoilData

Serrations B

Amplitude = 0.049c

Wavelength = 0.052c

Fig.4.119:Schematic

oftheleading

edgeserrations

(Case

B)configuration.

FlatPlateSerrations

B

Chapter4:

AirfoilProfiles

andPerform

ancePlots

211

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations B

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.120:Liftand

mom


flatplatew

ithleading

edgeserrations

(Case

B).

FlatPlateSerrations

B

212Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

B

Chapter4:

AirfoilProfiles

andPerform

ancePlots

213

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

B

214Sum

mary

ofLow

-SpeedA

irfoilData

Serrations C

Amplitude = 0.013c

Wavelength = 0.125c

Fig.4.121:Schematic

oftheleading

edgeserrations

(Case

C)configuration.

FlatPlateSerrations

C

Chapter4:

AirfoilProfiles

andPerform

ancePlots

215

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations C

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.122:Liftand

mom


flatplatew

ithleading

edgeserrations

(Case

C).

FlatPlateSerrations

C

216Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

C

Chapter4:

AirfoilProfiles

andPerform

ancePlots

217

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

C

218Sum

mary

ofLow

-SpeedA

irfoilData

Serrations D

Amplitude = 0.052c

Wavelength = 0.125c

Fig.4.123:Schematic

oftheleading

edgeserrations

(Case

D)configuration.

FlatPlateSerrations

D

Chapter4:

AirfoilProfiles

andPerform

ancePlots

219

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations D

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.124:Liftand

mom


flatplatew

ithleading

edgeserrations

(Case

D).

FlatPlateSerrations

D

220Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

D

Chapter4:

AirfoilProfiles

andPerform

ancePlots

221

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSerrations

D

222Sum

mary

ofLow

-SpeedA

irfoilData

Square Wave

Amplitude = 0.01575c

Wavelength = 0.063c

Fig.4.125:Schematic

oftheleading

edgesquare

wave

configuration.

FlatPlateSquare

Wave

Chapter4:

AirfoilProfiles

andPerform

ancePlots

223

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickSquare Wave

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.126:Liftand

mom


flatplatew

ithleading

edgesquare

waves.

FlatPlateSquare

Wave

224Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSquare

Wave

Chapter4:

AirfoilProfiles

andPerform

ancePlots

225

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSquare

Wave

226Sum

mary

ofLow

-SpeedA

irfoilData

Small Holes



Spacing = 0.016c

Fig.4.127:Schematic

oftheleading

edgeconfiguration

with

smallholes.

FlatPlateSm

allHoles

Chapter4:

AirfoilProfiles

andPerform

ancePlots

227

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickSmall Holes

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.128:Liftand

mom


flatplatew

ithsm

allholeson

theleading

edge.

FlatPlateSm

allHoles

228Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateSm

allHoles

Chapter4:

AirfoilProfiles

andPerform

ancePlots

229

230Sum

mary

ofLow

-SpeedA

irfoilData

Large Holes



Spacing = 0.052c

Fig.4.129:Schematic

oftheleading

edgeconfiguration

with

largeholes.

FlatPlateL

argeH

oles

Chapter4:

AirfoilProfiles

andPerform

ancePlots

231

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLarge Holes

Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLarge Holes

Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.130:Liftand

mom


flatplatew

ithlarge

holeson

theleading

edge.

FlatPlateL

argeH

oles

232Sum

mary

ofLow

-SpeedA

irfoilData

Small Cubes



Spacing = 0.010c

Fig.4.131:Schematic

oftheleading

edgeconfiguration

with

smallcubes.

FlatPlateSm

allCubes

Chapter4:

AirfoilProfiles

andPerform

ancePlots

233

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickSmall Cubes

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

Cm

Fig.4.132:Liftand

mom


flatplatew

ithsm

allcubeson

theleading

edge.

FlatPlateSm

allCubes

234Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

Cm


FlatPlateSm

allCubes

Chapter4:

AirfoilProfiles

andPerform

ancePlots

235

236Sum

mary

ofLow

-SpeedA

irfoilData

Large Cubes



Spacing = 0.042c

Fig.4.133:Schematic

oftheleading

edgeconfiguration

with

largecubes.

FlatPlateL

argeC

ubes

Chapter4:

AirfoilProfiles

andPerform

ancePlots

237

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

Flat Plate (C. Goudeseune)3.1% thickLarge Cubes

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.134:Liftand

mom


flatplatew

ithlarge

cubeson

theleading

edge.

FlatPlateL

argeC

ubes

238Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


FlatPlateL

argeC

ubes

Chapter4:

AirfoilProfiles

andPerform

ancePlots

239

240Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

MA409


Fig.4.135:Com

parisonbetw

eenthe

trueand

actualMA

409.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.012.0

0.240.470.700.931.161.391.621.84

α0l

= −4.1

x/c

V/V∞

MA409


distributionsforthe

MA

409.

MA

409

Chapter4:

AirfoilProfiles

andPerform

ancePlots

241

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 40,000Re = 60,000Re = 100,000

Re = 200,000Re = 300,000

MA409 (R. Cooney)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.137:Drag

PolarfortheM

A409

MA

409

242Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

MA409 (R. Cooney)

Re = 40,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

MA409 (R. Cooney)

Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.138:Liftand

mom


MA

409.

MA

409

Chapter4:

AirfoilProfiles

andPerform

ancePlots

243

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

MA409 (R. Cooney)

Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

MA409 (R. Cooney)

Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


MA

409

244Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

MA409 (R. Cooney)

Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


MA

409

Chapter4:

AirfoilProfiles

andPerform

ancePlots

245

246Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

NACA 43012A


Fig.4.139:Com

parisonbetw

eenthe

trueand

actualNA

CA

43012A.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

2.0 4.0 6.0 8.010.012.0

0.340.580.831.071.311.55

α0l

= −0.8

x/c

V/V∞

NACA 43012A


distributionsforthe

NA

CA

43012A.

NA

CA

43012A

Chapter4:

AirfoilProfiles

andPerform

ancePlots

247

0.00 0.02 0.04 0.06 0.08 0.10 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 400,000Re = 500,000

NACA 43012A (M. Nankivil)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.141:Drag

polarfortheN

AC

A43012A

.

NA

CA

43012A

248Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.142:Liftand

mom


NA

CA

43012A.

NA

CA

43012A

Chapter4:

AirfoilProfiles

andPerform

ancePlots

249

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


NA

CA

43012A

250Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


NA

CA

43012A

Chapter4:

AirfoilProfiles

andPerform

ancePlots

251

252Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

S1223


Fig.4.143:Com

parisonbetw

eenthe

trueand

actualS1223.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

2.0 4.0 6.0 8.010.012.014.0

1.822.062.292.512.742.963.18

α0l

= −14.2

x/c

V/V∞

S1223


distributionsforthe

S1223.

S1223clean

Chapter4:

AirfoilProfiles

andPerform

ancePlots

253

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)clean

Re = 80,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.145:Liftand

mom


S1223.

S1223clean

254Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 120,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 140,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223clean

Chapter4:

AirfoilProfiles

andPerform

ancePlots

255

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223clean

256Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223clean

Chapter4:

AirfoilProfiles

andPerform

ancePlots

257

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)Gurney Flap: h/c = 4.17%

Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.146:Liftand

mom


S1223w

ithG

urneyflap

ofh/c

=4.17%

.

S1223G

urneyFlap

h/c=

4.17%

258Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

4.17%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

259

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 140,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.147:Liftand

mom


S1223w

ithG

urneyflap

ofh/c

=3.12%

.

S1223G

urneyFlap

h/c=

3.12%

260Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

3.12%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

261

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

3.12%

262Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

263

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.148:Liftand

mom


S1223w

ithG

urneyflap

ofh/c

=2.08%

.

S1223G

urneyFlap

h/c=

2.08%

264Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

2.08%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

265

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.149:Liftand

mom


S1223w

ithG

urneyflap

ofh/c

=1.56%

.

S1223G

urneyFlap

h/c=

1.56%

266Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

1.56%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

267

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.150:Liftand

mom


S1223w

ithG

urneyflap

ofh/c

=1.04%

.

S1223G

urneyFlap

h/c=

1.04%

268Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl


Re = 250,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm


S1223G

urneyFlap

h/c=

1.04%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

269

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)u.s.t.: x/c = 15%, t/c = 0.11%

Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)u.s.t.: x/c = 15%, t/c = 0.11%

Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.151:Liftand

mom


S1223w

itha

boundary-layertripoft/c

=0.11%

.

S1223u.s.t.:x/c

=0.15%

t/c=

0.11%

270Sum

mary

ofLow

-SpeedA

irfoilData

Chapter4:

AirfoilProfiles

andPerform

ancePlots

271

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)u.s.t.: x/c = 15%, t/c = 0.19%

Re = 160,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

2.5

α (deg)

Cl

S1223 (Y. Tinel)u.s.t.: x/c = 15%, t/c = 0.19%

Re = 180,000

0.1

0.0

−0.1

−0.2

−0.3

−0.4

−0.5

Cm

Fig.4.152:Liftand

mom


S1223w

itha

boundary-layertripoft/c

=0.19%

.

S1223u.s.t.:x/c

=0.15%

t/c=

0.19%

272Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

S8064


Fig.4.153:Com

parisonbetw

eenthe

trueand

actualS8064.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

0.0 2.0 4.0 6.0 8.010.012.0

0.140.380.620.861.101.331.57

α0l

= −1.2

x/c

V/V∞

S8064


distributionsforthe

S8064.

S8064

Chapter4:

AirfoilProfiles

andPerform

ancePlots

273

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 100,000Re = 200,000Re = 300,000

Re = 400,000Re = 500,000

S8064 (T. Lampe)

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.155:Drag

polarfortheS8064.

S8064

274Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

S8064 (T. Lampe)

Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

S8064 (T. Lampe)

Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.156:Liftand

mom


S8064.

S8064

Chapter4:

AirfoilProfiles

andPerform

ancePlots

275

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

S8064 (T. Lampe)

Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

S8064 (T. Lampe)

Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S8064

276Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl

S8064 (T. Lampe)

Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S8064

Chapter4:

AirfoilProfiles

andPerform

ancePlots

277

278Sum

mary

ofLow

-SpeedA

irfoilData

Diff

eren

ce (

in)

−0.03

0.00

0.03

S9000


Fig.4.157:Com

parisonbetw

eenthe

trueand

actualS9000.

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.130.370.600.841.071.301.53

α0l

= −3.2

x/c

V/V∞

S9000 Flap 0 deg


distributionsforthe

S9000.

S9000Flap

0deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

279

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 400,000Re = 500,000

S9000 (T. Akers)Flap = 0 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.159:Drag

polarfortheS9000.

S9000Flap

0deg

cf /c

=20%

280Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.160:Liftand

mom


S9000.

S9000Flap

0deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

281

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

0deg

cf /c

=20%

282Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

0deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

283

284Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.300.530.761.001.231.461.69

α0l

= −4.6

x/c

V/V∞

S9000 Flap 2.5 deg


distributionsforthe

S9000w

itha

2.5deg

flap.

S9000Flap

2.5deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

285

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 400,000Re = 500,000

S9000 (T. Akers)Flap = 2.5 deg

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.162:Drag

polarfortheS9000

with

a2.5

degflap

S9000Flap

2.5deg

cf /c

=20%

286Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.163:Liftand

mom


S9000w

itha

2.5deg

flap

S9000Flap

2.5deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

287

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

2.5deg

cf /c

=20%

288Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

2.5deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

289

290Sum

mary

ofLow

-SpeedA

irfoilData

0 0.5 10

0.5

1

1.5

2

2.5 α Cl

−2.0 0.0 2.0 4.0 6.0 8.010.0

0.460.690.931.161.391.621.85

α0l

= −6.0

x/c

V/V∞

S9000 Flap 5 deg


distributionsforthe

S9000w

itha

5deg

flap.

S9000Flap

5deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

291

0.00 0.01 0.02 0.03 0.04 0.05 −0.5

0.0

0.5

1.0

1.5

Cd

Cl

Re = 60,000Re = 100,000Re = 200,000

Re = 300,000Re = 400,000Re = 500,000


−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

2.0

α (deg)

Cl

Fig.4.165:Drag

polarfortheS9000

with

a5

degflap

S9000Flap

5deg

cf /c

=20%

292Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 60,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 100,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

Fig.4.166:Liftand

mom


S9000w

itha

5deg

flap

S9000Flap

5deg

cf /c

=20%

Chapter4:

AirfoilProfiles

andPerform

ancePlots

293

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 200,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 300,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

5deg

cf /c

=20%

294Sum

mary

ofLow

-SpeedA

irfoilData

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 400,000

0.1

0.0

−0.1

−0.2

−0.3

Cm

−10 0 10 20 −0.5

0.0

0.5

1.0

1.5

α (deg)

Cl


Re = 500,000

0.1

0.0

−0.1

−0.2

−0.3

Cm


S9000Flap

5deg

cf /c

=20%

References

[1] Selig, M. S., Donovan, J. F., and Fraser, D. B., Airfoils at Low Speeds, SoarTech 8, SoarTech Publica-tions, Virginia Beach, VA, 1989.

[2] Selig, M. S., Guglielmo, J. J., Broeren, A. P., and Giguere, P., Summary of Low-Speed Airfoil Data -Vol. 1, SoarTech Publications, Virginia Beach, VA, 1995.

[3] Selig, M. S., Lyon, C. A., Giguere, P., Ninham, C. P., and Guglielmo, J. J., Summary of Low-SpeedAirfoil Data - Vol. 2, SoarTech Publications, Virginia Beach, VA, 1996.

[4] Lyon, C. A., Broeren, A. P., Giguere, P., Gopalarathnam, A., and Selig, M. S., Summary of Low-SpeedAirfoil Data - Vol. 3, SoarTech Publications, Virginia Beach, VA, 1997.

[5] Selig, M. S. and McGranahan, B. D., “Wind Tunnel Aerodynamic Tests of Six Airfoils for Use ofSmall Wind Turbines,” National Renewable Energy Laboratory, NREL/SR-500-34515, Golden, CO,2004.

[6] Anonymous, “Charles River Radio Controllers – Mark Drela’s AG and HT Air-foils,” http://www.charlesriverrc.org/articles/drela-airfoilshop/markdrela-ag-ht-airfoils.htm, Accessed September 2012.

[7] Selig, M. S., Deters, R. W., and Williamson, G. A., “Wind Tunnel Testing Airfoils at Low ReynoldsNumbers,” AIAA Paper 2011–875, January 2011.

[8] Anonymous, “Great Planes Model Manufacturing Company - Great Planes Viper 500 ARF,” http://www.greatplanes.com/airplanes/gpma1265.html, Accessed September 2012.

[9] Anonymous, “RC Universe Forum – Viper Airfoil for 428,” http://www.rcuniverse.com/forum/m_3624364/anchors_3628127/mpage_1/key_/anchor/tm.htm#3628127,Accessed September 2012.

[10] Selig, M. S., “S9000 Airfoil,” http://www.ae.illinois.edu/m-selig/ads/ref/misc_refs.html#5, Accessed September 2012.

[11] Broughton, B. A. and Selig, M. S., “Wind Tunnel Testing of the ND-LoFoil Airfoil with and withoutBoundary Layer Trips,” University of Illinois at Urbana-Champaign, Dept. of Aeronautical and Astro-nautical Engineering, AAE 01-05, UILU-ENG-01-0505, prepared for the T.J. Mueller, University ofNotre Dame, February 2001, 27 pages.

[12] Broughton, B. A., Carroll, C. A., and Selig, M. S., “Wind Tunnel Testing of the S1052 and S1054Flapped Airfoils for Unmanned Flying Wing Designs,” University of Illinois at Urbana-Champaign,Dept. of Aeronautical and Astronautical Engineering, AAE 01-02, UILU-ENG-01-0502, prepared forthe Naval Research Laboratory, Washington, DC, February 2001, 76 pages.


[13] Evangelista, R., McGhee, R. J., and Walker, B. S., “Correlation of Theory to Wind-Tunnel Data atReynolds Numbers below 500,000,” Lecture Notes in Engineering: Low Reynolds Number Aerody-namics, T.J. Mueller (ed.), Vol. 54, Springer-Verlag, New York, June 1989, pp. 131–145.

[14] Drela, M., “XFOIL: An Analysis and Design System for Low Reynolds Number Airfoils,” LectureNotes in Engineering: Low Reynolds Number Aerodynamics, T.J. Mueller (ed.), Vol. 54, Springer-Verlag, New York, June 1989, pp. 1–12.

[15] Drela, M., personal communication, 2003, 2011–2012.

Appendix A

Tabulated Airfoil Coordinates

Appendix A contains both the true (as designed) coordinates and actual (as built) coordinates for the airfoilstested during this study. For any given airfoil, the true coordinates are listed first.


AG12True

x/c y/c1.000000 0.0004710.994142 0.0010420.982204 0.0022220.968696 0.0035220.954910 0.0048230.941098 0.0060970.927274 0.0073450.913441 0.0085750.899593 0.0097870.885734 0.0109850.871861 0.0121720.857979 0.0133510.844090 0.0145220.830201 0.0156860.816312 0.0168450.802424 0.0179960.788539 0.0191410.774659 0.0202780.760778 0.0214070.746901 0.0225260.733027 0.0236340.719152 0.0247310.705280 0.0258170.691409 0.0268890.677538 0.0279480.663670 0.0289940.649803 0.0300240.635935 0.0310380.622068 0.0320360.608205 0.0330160.594342 0.0339780.580478 0.0349200.566619 0.0358410.552761 0.0367410.538905 0.0376170.525049 0.0384700.511197 0.0392960.497347 0.0400960.483500 0.0408660.469656 0.0416070.455816 0.0423130.441978 0.0429860.428143 0.0436220.414316 0.0442180.400492 0.0447740.386673 0.0452850.372861 0.0457510.359053 0.0461650.345250 0.046528

0.331461 0.0468360.317683 0.0470810.303912 0.0472620.290156 0.0473750.276418 0.0474130.262691 0.0473710.248985 0.0472430.235302 0.0470240.221638 0.0467040.208004 0.0462790.194402 0.0457370.180830 0.0450660.167304 0.0442600.153819 0.0433010.140392 0.0421800.127049 0.0408770.113793 0.0393710.100663 0.0376400.087677 0.0356530.074897 0.0333770.062374 0.0307700.050223 0.0277890.038618 0.0244000.027917 0.0206110.018756 0.0166010.011861 0.0128190.007244 0.0096420.004260 0.0070740.002309 0.0049470.001028 0.0031190.000261 0.0014850.000000 -0.0000110.000299 -0.0014810.001263 -0.0029040.002871 -0.0042290.005195 -0.0055650.008563 -0.0070320.013654 -0.0087220.021338 -0.0106190.031664 -0.0124660.043571 -0.0139820.056254 -0.0151470.069286 -0.0160030.082543 -0.0166090.095952 -0.0170160.109468 -0.0172630.123058 -0.0173750.136704 -0.0173780.150385 -0.0172830.164098 -0.0171050.177834 -0.0168510.191595 -0.0165310.205382 -0.016156

0.219178 -0.0157300.232992 -0.0152610.246823 -0.0147530.260663 -0.0142130.274519 -0.0136430.288396 -0.0130520.302276 -0.0124440.316162 -0.0118210.330060 -0.0111870.343964 -0.0105460.357874 -0.0098990.371793 -0.0092500.385719 -0.0086010.399644 -0.0079550.413564 -0.0073120.427475 -0.0066780.441384 -0.0060500.455289 -0.0054340.469185 -0.0048290.483074 -0.0042400.496967 -0.0036670.510856 -0.0031120.524739 -0.0025750.538618 -0.0020600.552501 -0.0015660.566379 -0.0010940.580251 -0.0006480.594125 -0.0002270.608000 0.0001670.621871 0.0005360.635737 0.0008760.649609 0.0011890.663478 0.0014740.677342 0.0017300.691207 0.0019560.705074 0.0021530.718938 0.0023210.732801 0.0024580.746669 0.0025650.760534 0.0026430.774396 0.0026890.788262 0.0027070.802128 0.0026950.815988 0.0026530.829854 0.0025810.843721 0.0024810.857585 0.0023510.871458 0.0021940.885322 0.0020110.899184 0.0017980.913057 0.0015570.926935 0.0012930.940803 0.001006

Appendix A: Tabulated Airfoil Coordinates 299

0.954669 0.0006900.968513 0.0003520.982113 -0.0000020.994133 -0.0003231.000000 -0.000471

AG12Actual

x/c y/c1.000000 0.0005050.998230 0.0006820.988216 0.0015200.980396 0.0020740.973312 0.0026810.963613 0.0034470.950447 0.0047760.939079 0.0058470.922094 0.0073860.903276 0.0091430.879817 0.0112690.857327 0.0132710.830060 0.0156530.803951 0.0178590.768071 0.0208050.737178 0.0233220.702424 0.0260530.668729 0.0286410.636410 0.0310620.600485 0.0336510.560393 0.0364080.526223 0.0385470.484597 0.0409050.445840 0.0428250.404493 0.0446210.364794 0.0460050.327188 0.0468830.287466 0.0472650.253841 0.0471710.218409 0.0464540.184435 0.0450270.155343 0.0431840.123600 0.0402500.098271 0.0370350.071520 0.0324590.056470 0.0292090.045328 0.0262660.032755 0.0222030.022522 0.0180900.017904 0.0158880.013463 0.0135070.010019 0.0113440.006113 0.008438

0.003102 0.0055310.001501 0.0033950.000639 0.0018510.000137 -0.0007060.001732 -0.0033560.005585 -0.0064090.010168 -0.0084510.016911 -0.0105450.023973 -0.0121500.040621 -0.0147500.059788 -0.0165750.075527 -0.0174670.105921 -0.0183720.133915 -0.0185280.163923 -0.0182090.194454 -0.0175000.228820 -0.0164090.263189 -0.0150760.300827 -0.0133940.335346 -0.0117480.372250 -0.0099870.413117 -0.0080170.450201 -0.0063150.486537 -0.0047120.537954 -0.0026100.580379 -0.0011430.618519 -0.0000350.652793 0.0007400.688386 0.0013670.720718 0.0017520.750856 0.0019850.782507 0.0020250.811917 0.0020030.837081 0.0018710.886480 0.0013820.905846 0.0010530.926689 0.0006430.945487 0.0002230.987583 -0.0004050.991692 -0.0004320.994807 -0.0004210.997832 -0.0003680.998641 -0.0002931.000000 -0.000147

AG16True

x/c y/c0.999982 0.0002590.994055 0.0008950.982059 0.0021850.968510 0.003609

0.954662 0.0050470.940774 0.0064840.926885 0.0079120.912993 0.0093310.899100 0.0107430.885210 0.0121460.871319 0.0135390.857433 0.0149210.843545 0.0162910.829661 0.0176470.815777 0.0189890.801890 0.0203180.788007 0.0216310.774122 0.0229280.760234 0.0242100.746349 0.0254760.732460 0.0267260.718568 0.0279600.704680 0.0291790.690792 0.0303790.676901 0.0315610.663014 0.0327250.649130 0.0338680.635243 0.0349900.621358 0.0360900.607477 0.0371660.593595 0.0382170.579713 0.0392430.565836 0.0402410.551961 0.0412100.538084 0.0421480.524210 0.0430550.510343 0.0439270.496475 0.0447620.482606 0.0455620.468743 0.0463230.454885 0.0470410.441027 0.0477160.427174 0.0483470.413328 0.0489280.399483 0.0494590.385642 0.0499370.371812 0.0503590.357986 0.0507200.344163 0.0510210.330353 0.0512560.316555 0.0514200.302760 0.0515090.288982 0.0515220.275220 0.0514500.261465 0.0512890.247732 0.0510360.234021 0.050681


0.220327 0.0502190.206665 0.0496430.193035 0.0489420.179439 0.0481080.165899 0.0471280.152408 0.0459860.138987 0.0446710.125653 0.0431580.112413 0.0414310.099311 0.0394630.086363 0.0372220.073635 0.0346780.061172 0.0317870.049093 0.0285120.037574 0.0248300.026987 0.0207700.018028 0.0165680.011373 0.0127010.006948 0.0094950.004080 0.0069190.002203 0.0047810.000968 0.0029470.000230 0.0013120.000000 -0.0001950.000341 -0.0016730.001328 -0.0031300.002935 -0.0045520.005261 -0.0060460.008653 -0.0077300.013784 -0.0097250.021492 -0.0120250.031716 -0.0143310.043529 -0.0162910.056132 -0.0178870.069116 -0.0191470.082329 -0.0201220.095697 -0.0208630.109171 -0.0214000.122724 -0.0217650.136334 -0.0219770.149994 -0.0220570.163688 -0.0220190.177415 -0.0218770.191166 -0.0216440.204936 -0.0213330.218725 -0.0209520.232530 -0.0205090.246346 -0.0200120.260179 -0.0194690.274022 -0.0188840.287877 -0.0182670.301747 -0.0176270.315629 -0.016960

0.329522 -0.0162710.343431 -0.0155630.357351 -0.0148410.371284 -0.0141060.385229 -0.0133640.399185 -0.0126150.413144 -0.0118650.427092 -0.0111160.441028 -0.0103730.454954 -0.0096350.468873 -0.0089060.482781 -0.0081880.496682 -0.0074840.510577 -0.0067940.524471 -0.0061210.538359 -0.0054650.552243 -0.0048310.566124 -0.0042180.580004 -0.0036280.593877 -0.0030610.607748 -0.0025210.621614 -0.0020070.635476 -0.0015210.649333 -0.0010640.663189 -0.0006370.677044 -0.0002380.690898 0.0001280.704753 0.0004610.718610 0.0007640.732471 0.0010320.746331 0.0012650.760195 0.0014690.774062 0.0016330.787929 0.0017650.801802 0.0018630.815681 0.0019210.829560 0.0019440.843441 0.0019320.857326 0.0018830.871211 0.0017990.885098 0.0016800.898988 0.0015210.912876 0.0013240.926767 0.0010970.940657 0.0008390.954550 0.0005380.968415 0.0002080.982033 -0.0001490.994056 -0.0004941.000001 -0.000670

AG16Actual

x/c y/c1.000000 0.0008150.998728 0.0009530.991688 0.0020280.984173 0.0027870.978396 0.0034070.963070 0.0050530.945040 0.0068520.927906 0.0086220.909848 0.0105750.892850 0.0122990.867679 0.0148790.844498 0.0172200.814715 0.0202630.786169 0.0229710.755036 0.0259440.726636 0.0286150.695687 0.0313500.671374 0.0334280.638131 0.0362240.601100 0.0392480.573694 0.0413620.531611 0.0443620.489188 0.0470840.398833 0.0516210.361817 0.0528760.323007 0.0537060.289307 0.0539920.252929 0.0537770.218211 0.0529060.180839 0.0510790.159297 0.0495820.141549 0.0480150.134797 0.0472910.114939 0.0448460.100592 0.0427430.077957 0.0387430.062804 0.0354210.053841 0.0330970.046415 0.0309570.038933 0.0285320.032101 0.0260420.027799 0.0243210.023528 0.0224790.021230 0.0214070.017249 0.0193800.013952 0.0175190.009929 0.0149090.006473 0.0122450.004048 0.009962


0.001877 0.0073330.000779 0.0054840.005057 -0.0036250.007805 -0.0050220.010263 -0.0060890.013732 -0.0073750.018302 -0.0087610.030018 -0.0115300.037964 -0.0129630.050149 -0.0146790.061599 -0.0159400.075006 -0.0170840.096442 -0.0183960.120894 -0.0193210.148281 -0.0197500.177107 -0.0198100.207568 -0.0192440.241409 -0.0181060.277127 -0.0166900.314872 -0.0149720.349176 -0.0133000.393866 -0.0110270.426083 -0.0094150.465362 -0.0074310.498947 -0.0057830.553830 -0.0032550.588967 -0.0017920.626002 -0.0004400.663262 0.0007290.694986 0.0015210.731267 0.0022120.762621 0.0026260.792839 0.0029220.817618 0.0029950.848593 0.0029680.867268 0.0029010.884831 0.0027180.903411 0.0024520.922802 0.0020840.931966 0.0018780.939497 0.0016730.948464 0.0015040.955668 0.0013440.960393 0.0011820.965343 0.0010920.972199 0.0009010.977309 0.0007960.982570 0.0005830.986560 0.0004970.990207 0.0003430.995900 0.0002130.997726 0.0001631.000000 0.000165

AG24True

x/c y/c1.000000 0.0003120.994048 0.0010430.982038 0.0026300.968488 0.0044860.954647 0.0064210.940769 0.0083780.926890 0.0103280.913006 0.0122580.899118 0.0141650.885230 0.0160480.871339 0.0179080.857451 0.0197440.843559 0.0215620.829670 0.0233570.815779 0.0251300.801883 0.0268790.787989 0.0286020.774091 0.0302970.760188 0.0319640.746285 0.0336000.732376 0.0352040.718461 0.0367760.704547 0.0383130.690631 0.0398130.676709 0.0412770.662788 0.0427020.648868 0.0440870.634943 0.0454320.621017 0.0467360.607093 0.0479970.593166 0.0492140.579236 0.0503890.565310 0.0515170.551385 0.0526000.537456 0.0536350.523528 0.0546220.509606 0.0555580.495682 0.0564430.481756 0.0572760.467835 0.0580530.453918 0.0587740.440000 0.0594370.426087 0.0600400.412180 0.0605790.398274 0.0610530.384372 0.0614580.370481 0.0617920.356594 0.0620500.342711 0.062231

0.328841 0.0623280.314984 0.0623380.301131 0.0622540.287296 0.0620740.273479 0.0617890.259672 0.0613940.245889 0.0608830.232132 0.0602480.218395 0.0594800.204696 0.0585720.191034 0.0575140.177412 0.0562950.163854 0.0549040.150354 0.0533250.136934 0.0515450.123613 0.0495420.110400 0.0472960.097342 0.0447840.084457 0.0419710.071816 0.0388320.059466 0.0353230.047533 0.0314130.036200 0.0270830.025838 0.0223930.017129 0.0176340.010712 0.0133410.006482 0.0098570.003763 0.0071070.001997 0.0048650.000855 0.0029630.000198 0.0012870.000001 -0.0002300.000293 -0.0017190.001184 -0.0032350.002714 -0.0047380.004994 -0.0063010.008364 -0.0080340.013488 -0.0100530.021193 -0.0123770.031412 -0.0147260.043212 -0.0167960.055804 -0.0185190.068780 -0.0199180.081987 -0.0210420.095351 -0.0219380.108824 -0.0226390.122377 -0.0231750.135989 -0.0235650.149652 -0.0238270.163351 -0.0239770.177085 -0.0240260.190844 -0.0239860.204622 -0.023867


0.218421 -0.0236750.232237 -0.0234200.246064 -0.0231060.259909 -0.0227400.273764 -0.0223260.287631 -0.0218710.301514 -0.0213810.315408 -0.0208560.329313 -0.0203010.343234 -0.0197180.357166 -0.0191110.371111 -0.0184820.385067 -0.0178350.399034 -0.0171730.413004 -0.0164980.426962 -0.0158130.440907 -0.0151210.454842 -0.0144240.468769 -0.0137230.482685 -0.0130210.496593 -0.0123190.510494 -0.0116170.524394 -0.0109190.538286 -0.0102240.552174 -0.0095350.566059 -0.0088520.579942 -0.0081770.593817 -0.0075110.607689 -0.0068570.621556 -0.0062140.635419 -0.0055850.649276 -0.0049710.663131 -0.0043720.676986 -0.0037880.690839 -0.0032210.704693 -0.0026700.718548 -0.0021360.732407 -0.0016190.746265 -0.0011220.760127 -0.0006450.773993 -0.0001940.787859 0.0002300.801731 0.0006220.815611 0.0009750.829491 0.0012860.843374 0.0015500.857263 0.0017590.871153 0.0019090.885046 0.0019910.898943 0.0019970.912839 0.0019200.926737 0.0017580.940634 0.001501

0.954532 0.0011490.968401 0.0007070.982019 0.0001850.994039 -0.0003641.000000 -0.000659

AG24Actual

x/c y/c1.000000 0.0009130.999208 0.0010370.991941 0.0020660.986634 0.0029440.979522 0.0039070.974384 0.0046690.960150 0.0067560.942990 0.0092090.924086 0.0118240.901426 0.0149230.877652 0.0181090.852626 0.0213770.823584 0.0250600.795223 0.0285800.767484 0.0319340.735246 0.0356360.703237 0.0391710.670010 0.0426370.630963 0.0464560.597032 0.0495220.560502 0.0525100.521653 0.0553660.479354 0.0579630.440799 0.0599140.403083 0.0613910.365784 0.0623100.334281 0.0626970.295212 0.0625870.258215 0.0617330.223211 0.0601350.191437 0.0578810.154136 0.0542030.127260 0.0505210.101749 0.0461350.077853 0.0408540.066922 0.0379900.056572 0.0349530.050186 0.0328890.044234 0.0307750.034297 0.0267750.028286 0.0240600.020580 0.0201440.018252 0.018855

0.015134 0.0169750.012560 0.0152850.010860 0.0140810.008487 0.0122500.006870 0.0109030.005436 0.0095820.004158 0.0083080.002767 0.0066640.001613 0.0049010.000956 0.0037220.000122 0.0015650.000079 0.0009890.000001 -0.0001260.000128 -0.0012600.000457 -0.0021310.001034 -0.0030670.004501 -0.0061810.005538 -0.0068080.011565 -0.0096640.018506 -0.0119920.022396 -0.0130790.030689 -0.0149660.041519 -0.0169250.049752 -0.0181120.067217 -0.0201870.080488 -0.0214010.108947 -0.0230780.137397 -0.0239780.169180 -0.0243090.199472 -0.0243460.232751 -0.0239210.267297 -0.0229110.300176 -0.0218590.337244 -0.0203820.378891 -0.0185170.414685 -0.0167900.452125 -0.0149490.495925 -0.0127110.540631 -0.0104010.576810 -0.0085840.617256 -0.0065560.652837 -0.0049200.686773 -0.0033860.720336 -0.0019720.752377 -0.0007450.782352 0.0002740.811684 0.0011160.837081 0.0016830.861622 0.0020780.884034 0.0022770.905993 0.0023590.925194 0.0021460.941065 0.001806


0.954473 0.0015360.967302 0.0012090.977787 0.0008670.984280 0.0005220.990342 0.0000990.994594 -0.0000900.997513 -0.0001400.999231 -0.0001391.000000 -0.000118

AG35-rTrue

x/c y/c1.000337 0.0012450.995077 0.0019570.985369 0.0032740.973811 0.0048420.961211 0.0065510.948180 0.0083180.934972 0.0101100.921629 0.0119190.908277 0.0137310.894978 0.0155350.881941 0.0173030.869446 0.0189980.857212 0.0206560.844986 0.0223150.832543 0.0240030.819963 0.0256630.806718 0.0272920.793426 0.0289270.780162 0.0305580.766892 0.0321910.753640 0.0338210.740403 0.0354490.727204 0.0370730.714059 0.0386900.701071 0.0402880.688370 0.0418500.676265 0.0433400.664894 0.0447380.652934 0.0461150.640385 0.0473470.628548 0.0483540.615519 0.0493580.602357 0.0503720.588989 0.0514020.575515 0.0524410.562044 0.0534790.548602 0.0545140.535187 0.0555480.521829 0.056577

0.508494 0.0576040.495235 0.0586260.482099 0.0596380.469305 0.0606230.456850 0.0615790.444357 0.0624290.431863 0.0632140.419284 0.0639340.406674 0.0645800.393976 0.0651570.381170 0.0656540.368220 0.0660700.355157 0.0664010.342058 0.0666340.328962 0.0667730.315909 0.0668130.302921 0.0667510.289991 0.0665880.277118 0.0663210.264273 0.0659480.251441 0.0654650.238611 0.0648700.225809 0.0641580.213049 0.0633260.200367 0.0623690.187792 0.0612840.175347 0.0600810.163055 0.0587440.150914 0.0572610.138907 0.0556160.127070 0.0538020.115424 0.0518080.103979 0.0496160.092738 0.0472070.081633 0.0445440.070749 0.0416230.060274 0.0384730.050365 0.0351160.041196 0.0316020.032999 0.0280330.025959 0.0245430.020144 0.0212700.015474 0.0182920.011772 0.0156250.008836 0.0132430.006499 0.0111030.004648 0.0091810.003174 0.0074190.002004 0.0057710.001111 0.0042140.000481 0.0027280.000110 0.0012950.000000 -0.000071

0.000137 -0.0014060.000556 -0.0027570.001319 -0.0041180.002416 -0.0054580.003842 -0.0067300.005643 -0.0079410.007850 -0.0091050.010523 -0.0102450.013787 -0.0113890.017841 -0.0125600.022954 -0.0137760.029430 -0.0150440.037491 -0.0163280.047073 -0.0175600.057842 -0.0186640.069391 -0.0196040.081411 -0.0203780.093721 -0.0210020.106220 -0.0214930.118844 -0.0218690.131560 -0.0221420.144350 -0.0223260.157195 -0.0224320.170086 -0.0224650.183018 -0.0224370.195990 -0.0223520.208993 -0.0222180.222026 -0.0220410.235083 -0.0218300.248159 -0.0215820.261256 -0.0213060.274375 -0.0210040.287512 -0.0206800.300665 -0.0203400.313836 -0.0199850.327032 -0.0196240.340240 -0.0192630.353448 -0.0189030.366656 -0.0185420.379871 -0.0181820.393091 -0.0178210.406313 -0.0174600.419530 -0.0170990.432743 -0.0167380.445957 -0.0163780.459177 -0.0160170.472398 -0.0156560.485612 -0.0152960.498826 -0.0149360.512043 -0.0145740.525264 -0.0142130.538484 -0.0138520.551699 -0.013493


0.564907 -0.0131310.578117 -0.0127720.591332 -0.0124100.604548 -0.0120490.617761 -0.0116890.630980 -0.0113280.644201 -0.0109670.657420 -0.0106070.670636 -0.0102450.683849 -0.0098850.697066 -0.0095250.710286 -0.0091640.723506 -0.0088020.736720 -0.0084430.749933 -0.0080810.763151 -0.0077200.776373 -0.0073590.789591 -0.0069990.802806 -0.0066390.816013 -0.0062780.829224 -0.0059180.842438 -0.0055560.855650 -0.0051950.868855 -0.0048360.882060 -0.0044740.895271 -0.0041150.908487 -0.0037540.921703 -0.0033930.934907 -0.0030330.948066 -0.0026740.961081 -0.0023190.973699 -0.0019730.985310 -0.0016570.995133 -0.0013891.000408 -0.001244

AG35-rActual

x/c y/c1.000000 0.0010910.997432 0.0015230.995894 0.0017350.991424 0.0023680.987464 0.0030570.983225 0.0036470.973355 0.0050890.958941 0.0071520.944288 0.0091840.925458 0.0117800.906066 0.0145000.878721 0.0183070.859359 0.020904

0.831023 0.0246350.803034 0.0282440.787107 0.0302430.778725 0.0312640.757512 0.0338600.722589 0.0381630.690150 0.0420090.620846 0.0494830.583881 0.0524520.543548 0.0555810.509604 0.0581830.473268 0.0609250.433347 0.0636020.395245 0.0655830.356010 0.0668500.317912 0.0672110.279489 0.0666030.248140 0.0654960.208234 0.0631570.183573 0.0611020.152062 0.0575530.124497 0.0535610.099375 0.0487070.078567 0.0436100.066361 0.0401280.056908 0.0370900.046268 0.0332810.040537 0.0310120.035126 0.0286520.029980 0.0262200.025664 0.0240440.023653 0.0229620.020719 0.0212630.018941 0.0201730.015173 0.0176660.012122 0.0154930.011068 0.0147180.008651 0.0128010.007015 0.0113830.005367 0.0097720.003431 0.0075830.001900 0.0054030.000878 0.0033860.000117 0.0014570.000001 0.0001050.000189 -0.0018520.000875 -0.0039230.001552 -0.0051210.004144 -0.0077840.005865 -0.0089760.007829 -0.0100380.010349 -0.0111660.013137 -0.012254

0.017718 -0.0136770.021507 -0.0147030.033446 -0.0170440.051223 -0.0192280.071310 -0.0208660.093365 -0.0221460.119325 -0.0231390.146112 -0.0235980.174589 -0.0236320.205193 -0.0233110.238143 -0.0225580.275599 -0.0215210.314083 -0.0204890.349016 -0.0195000.387418 -0.0183930.427894 -0.0172310.462334 -0.0162230.502851 -0.0150300.551370 -0.0135910.586165 -0.0125880.616656 -0.0117130.650112 -0.0107340.683456 -0.0097530.710481 -0.0089600.739217 -0.0080960.773629 -0.0070180.800531 -0.0061860.820882 -0.0055680.843980 -0.0049340.865258 -0.0042720.879368 -0.0039050.894557 -0.0034470.905311 -0.0030960.912955 -0.0028830.920631 -0.0026780.931923 -0.0023510.943772 -0.0020290.950670 -0.0018360.961138 -0.0015580.971357 -0.0012750.980229 -0.0010640.986103 -0.0008620.987723 -0.0008470.991512 -0.0007270.995581 -0.0006230.998177 -0.0005360.999641 -0.0005201.000000 -0.000501


AG40d-02rTrue

x/c y/c1.000000 0.0004780.994054 0.0010770.982165 0.0024470.968707 0.0040000.955028 0.0055140.941303 0.0070400.927596 0.0085610.913882 0.0101230.900153 0.0116890.886424 0.0132470.872689 0.0147870.858978 0.0163070.845256 0.0178130.831558 0.0193020.817862 0.0207690.804186 0.0222120.790513 0.0236290.776850 0.0250170.763195 0.0263710.758006 0.0268760.755235 0.0271440.751672 0.0275630.747315 0.0281330.736722 0.0295100.723013 0.0312470.712660 0.0325320.695454 0.0346040.681588 0.0362120.667686 0.0377730.656489 0.0389940.650920 0.0395890.639796 0.0407510.625817 0.0421690.613984 0.0433300.597880 0.0448550.583944 0.0461210.570012 0.0473390.558942 0.0482720.549985 0.0490030.541959 0.0496400.528043 0.0507040.514128 0.0517160.500179 0.0526770.486200 0.0535850.472209 0.0544360.458206 0.0552290.444178 0.0559620.430126 0.0566350.416049 0.057244

0.401961 0.0577860.387896 0.0582560.373869 0.0586540.359878 0.0589760.345919 0.0592180.331997 0.0593770.318109 0.0594510.304252 0.0594360.290430 0.0593250.276625 0.0591170.262828 0.0588030.249051 0.0583790.235294 0.0578380.221551 0.0571690.207837 0.0563640.194155 0.0554240.180501 0.0543310.166900 0.0530740.153345 0.0516400.139858 0.0500120.126454 0.0481710.113141 0.0460940.099962 0.0437590.086936 0.0411340.074119 0.0381970.061549 0.0349070.049346 0.0312310.037694 0.0271220.026971 0.0226190.017869 0.0179820.011096 0.0137170.006584 0.0101800.003679 0.0073160.001804 0.0049300.000652 0.0028750.000091 0.0010850.000013 -0.0004280.000284 -0.0018300.001126 -0.0032690.002626 -0.0047430.004894 -0.0063150.008268 -0.0080840.013250 -0.0100820.020143 -0.0121830.028974 -0.0142260.039695 -0.0161130.051627 -0.0176980.064190 -0.0189570.077232 -0.0199430.090870 -0.0207150.105062 -0.0213040.119601 -0.0217320.134234 -0.022023

0.148850 -0.0221920.163428 -0.0222660.177968 -0.0222620.192444 -0.0221900.206857 -0.0220640.221090 -0.0218950.235287 -0.0216920.249436 -0.0214560.263570 -0.0211920.277689 -0.0209000.291816 -0.0205830.305952 -0.0202430.320107 -0.0198800.334290 -0.0194990.348500 -0.0191030.362730 -0.0186900.377003 -0.0182660.391304 -0.0178310.405592 -0.0173860.419896 -0.0169340.434181 -0.0164740.448442 -0.0160110.462684 -0.0155410.476907 -0.0150660.491095 -0.0145880.505262 -0.0141080.519422 -0.0136240.533574 -0.0131410.547715 -0.0126580.555357 -0.0123960.565055 -0.0120630.575458 -0.0117090.589006 -0.0112470.602553 -0.0107880.608953 -0.0105710.616233 -0.0103260.629888 -0.0098710.643526 -0.0094200.653333 -0.0090980.662365 -0.0088010.670765 -0.0085290.684400 -0.0080910.698020 -0.0076560.712999 -0.0071850.724857 -0.0068180.738189 -0.0064090.743882 -0.0062360.747950 -0.0061140.753502 -0.0059480.758125 -0.0058100.765054 -0.0056050.778506 -0.0052110.791927 -0.004828


0.805336 -0.0044520.818895 -0.0040830.832453 -0.0037240.846006 -0.0033770.859583 -0.0030450.873153 -0.0027240.886712 -0.0024200.900308 -0.0021330.913918 -0.0018520.927552 -0.0015970.941205 -0.0013540.954907 -0.0011280.968612 -0.0009090.982101 -0.0007160.994042 -0.0005571.000000 -0.000478

AG40d-02rActual

x/c y/c1.000000 0.0002830.990989 0.0013110.973043 0.0034380.954935 0.0054940.939409 0.0071540.918306 0.0095480.898215 0.0117890.876763 0.0142020.825857 0.0198510.793702 0.0233400.766521 0.0260940.738633 0.0294080.702342 0.0339720.667117 0.0380490.636811 0.0412790.598439 0.0450100.563137 0.0481460.522877 0.0512870.477931 0.0542900.439679 0.0563210.397393 0.0582130.362714 0.0591650.327708 0.0597080.290868 0.0596370.252756 0.0588630.217607 0.0572380.180313 0.0545750.151338 0.0516730.123842 0.0480380.094467 0.0429260.073206 0.0381160.052963 0.032479

0.040003 0.0280940.034163 0.0258100.026938 0.0226950.021327 0.0199500.014666 0.0161410.010001 0.0130100.007804 0.0112680.005521 0.0092440.004708 0.0084020.002975 0.0063540.001462 0.0041710.000457 0.0021060.000105 0.0008220.000004 -0.0001610.000147 -0.0009730.001486 -0.0035200.003356 -0.0054050.007541 -0.0081000.010916 -0.0096170.015749 -0.0113650.023172 -0.0134490.030185 -0.0149660.045148 -0.0174230.067756 -0.0197990.090376 -0.0212260.114505 -0.0222170.140179 -0.0228370.173693 -0.0230850.205520 -0.0229920.241251 -0.0225570.274734 -0.0219740.307351 -0.0212370.345101 -0.0203370.382510 -0.0192870.420936 -0.0180830.459624 -0.0169040.498837 -0.0156180.552033 -0.0138180.590354 -0.0124870.623461 -0.0113270.661500 -0.0100780.694860 -0.0089380.725654 -0.0079340.758741 -0.0068920.788048 -0.0060340.815384 -0.0052490.840282 -0.0046130.867870 -0.0038410.886589 -0.0033680.908515 -0.0028870.925994 -0.0023850.945888 -0.0019520.960939 -0.001512

0.976995 -0.0012290.989888 -0.0007510.995443 -0.0005510.997540 -0.0003971.000000 -0.000237

AG455ct-02rTrue

x/c y/c1.000000 0.0004220.994170 0.0009540.982316 0.0020350.968892 0.0032210.955202 0.0044240.941444 0.0056240.927663 0.0067910.913863 0.0079450.900051 0.0090790.886239 0.0101960.872425 0.0112900.858612 0.0123600.844813 0.0134160.831008 0.0144480.817215 0.0154630.803422 0.0164590.789639 0.0174320.775854 0.0183930.762077 0.0193320.748302 0.0202520.734533 0.0211560.720760 0.0220410.710075 0.0227170.704818 0.0230440.702188 0.0232070.699579 0.0234580.694361 0.0239610.680231 0.0253100.666108 0.0266350.659173 0.0272780.648776 0.0282350.637922 0.0292210.623855 0.0304810.607516 0.0319150.595808 0.0329230.581851 0.0341030.567919 0.0352570.556858 0.0361530.547916 0.0368650.539911 0.0374940.526047 0.0385560.512218 0.0395850.498385 0.040583


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AG455ct-02rActual

x/c y/c1.000000 0.0004880.999334 0.0005850.992577 0.0014500.984324 0.0023240.977756 0.0030820.970349 0.0035030.963038 0.0041110.947330 0.0056460.926492 0.0075040.909954 0.0089830.888459 0.0108160.865706 0.0126400.840791 0.0145570.814702 0.0164270.787289 0.0182770.757929 0.0201410.727475 0.0219090.687524 0.0248030.658313 0.0275620.625135 0.0305900.587918 0.0337950.515952 0.0394730.470748 0.0426420.431370 0.0450080.393190 0.0470090.354360 0.0486050.318301 0.049610


0.280791 0.0501820.246697 0.0502180.211778 0.0495930.179151 0.0482630.144672 0.0458680.115528 0.0428160.092142 0.0394840.069824 0.0353070.049584 0.0302890.025383 0.0215920.017818 0.0177720.013341 0.0151300.010813 0.0134680.007720 0.0111530.006007 0.0096440.004424 0.0080900.003229 0.0067620.002505 0.0058350.001340 0.0039450.000677 0.0025250.000104 0.0008510.000005 -0.0001890.000084 -0.0007640.000052 -0.0006810.000489 -0.0018120.000913 -0.0024990.001852 -0.0034950.004680 -0.0052680.010442 -0.0075170.014862 -0.0087250.021208 -0.0100370.026771 -0.0109510.035643 -0.0121560.044735 -0.0131330.056534 -0.0140550.077638 -0.0151090.099559 -0.0157610.124306 -0.0161450.153575 -0.0162650.182524 -0.0160320.214094 -0.0155520.245662 -0.0150230.282032 -0.0143140.321311 -0.0135710.352669 -0.0129550.394479 -0.0120700.423808 -0.0114610.462308 -0.0107430.498729 -0.0099840.550621 -0.0089990.598723 -0.0080620.636002 -0.0073670.669935 -0.006742

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CAL1215jTrue

x/c y/c1.000000 0.0000000.997947 0.0002740.992015 0.0011960.982503 0.0027400.969609 0.0049710.953602 0.0078680.934702 0.0113660.913107 0.0154350.889084 0.0200040.862869 0.0249170.834627 0.0300610.804570 0.0353780.772903 0.0407270.739795 0.0460210.705496 0.0511920.670182 0.0561260.634036 0.0607770.597301 0.0650440.560147 0.0688550.522781 0.0722030.485421 0.0750070.448249 0.0772330.411505 0.0788610.375370 0.0798400.340020 0.0801670.305676 0.0798060.272492 0.0787380.240645 0.0769940.210301 0.0745380.181574 0.0713900.154644 0.067593

0.129602 0.0631440.106522 0.0581150.085549 0.0525460.066731 0.0464850.050126 0.0400690.035821 0.0333550.023827 0.0264490.014227 0.0195380.007044 0.0127240.002256 0.0062270.000015 0.0005040.001169 -0.0044900.005826 -0.0093830.013289 -0.0142970.023461 -0.0189580.036322 -0.0232260.051851 -0.0270210.070016 -0.0302870.090774 -0.0330210.114029 -0.0352180.139683 -0.0368570.167622 -0.0379710.197692 -0.0385790.229753 -0.0386920.263623 -0.0383570.299103 -0.0376040.336001 -0.0364820.374089 -0.0350430.413130 -0.0333240.452886 -0.0313840.493091 -0.0292710.533494 -0.0270310.573821 -0.0247250.613791 -0.0223900.653140 -0.0200710.691580 -0.0178140.728837 -0.0156440.764646 -0.0135920.798739 -0.0116770.830868 -0.0099120.860790 -0.0083140.888269 -0.0068770.913096 -0.0055960.935069 -0.0044560.953994 -0.0034140.969786 -0.0023700.982464 -0.0013530.991927 -0.0005560.997915 -0.0000991.000000 0.000000


CAL1215jActual

x/c y/c1.000000 0.0024210.998617 0.0029940.993728 0.0039970.989628 0.0048970.979387 0.0067310.973380 0.0078660.967294 0.0089870.958765 0.0104740.948584 0.0122250.940596 0.0137490.935436 0.0145500.895041 0.0214730.873672 0.0250510.849180 0.0290430.822684 0.0335000.793092 0.0384150.764984 0.0429150.734554 0.0479500.702531 0.0527940.668808 0.0573280.631039 0.0619330.594449 0.0663810.552225 0.0704670.515387 0.0732760.475710 0.0758770.437987 0.0779280.400132 0.0793210.358882 0.0799360.323280 0.0799040.284693 0.0789130.251616 0.0771860.214991 0.0744090.181675 0.0708370.148243 0.0659570.121711 0.0611570.101363 0.0565830.086699 0.0526900.074803 0.0489250.063495 0.0449910.053870 0.0412520.044154 0.0371220.035554 0.0327850.029905 0.0296240.021744 0.0246590.014471 0.0198330.009480 0.0155430.005649 0.0111980.003326 0.0079680.001399 0.004726

0.000291 0.0021910.000064 -0.0010240.000266 -0.0018450.000300 -0.0021440.001795 -0.0042580.004219 -0.0070330.007388 -0.0097150.010298 -0.0115260.014627 -0.0140650.022280 -0.0177330.030405 -0.0209260.038430 -0.0232800.050772 -0.0263380.067018 -0.0294280.080657 -0.0313210.103873 -0.0337380.131247 -0.0357390.166796 -0.0372830.200942 -0.0380490.233344 -0.0382100.269202 -0.0378480.342634 -0.0360100.377854 -0.0347350.416746 -0.0330480.455290 -0.0313500.500280 -0.0293640.543453 -0.0270510.581345 -0.0248540.618769 -0.0227500.654484 -0.0205030.687806 -0.0184830.722604 -0.0163880.753843 -0.0144640.786170 -0.0125110.811910 -0.0108690.838829 -0.0090300.861488 -0.0076450.884704 -0.0065970.905029 -0.0056930.923185 -0.0051040.940540 -0.0044340.955330 -0.0040170.967308 -0.0037290.975756 -0.0036300.982891 -0.0036560.988356 -0.0035470.992079 -0.0034630.996097 -0.0032070.997845 -0.0030811.000000 -0.002605

CAL2263mTrue

x/c y/c1.000000 0.0000000.997910 0.0003880.991905 0.0016800.982356 0.0038730.969567 0.0070250.953912 0.0110520.935718 0.0156940.915024 0.0207530.891896 0.0261870.866583 0.0319090.839252 0.0378170.810111 0.0438460.779346 0.0498480.747113 0.0557410.713650 0.0614560.679127 0.0668870.643723 0.0719860.607677 0.0766590.571168 0.0808340.534408 0.0844850.497600 0.0875170.460914 0.0898910.424585 0.0915830.388774 0.0925360.353639 0.0927550.319400 0.0922110.286202 0.0908900.254215 0.0888360.223608 0.0860240.194503 0.0824790.167078 0.0782470.141427 0.0733280.117620 0.0678000.095804 0.0617120.076027 0.0551240.058336 0.0481850.042848 0.0410210.029664 0.0337520.018960 0.0264600.010696 0.0191220.004733 0.0119100.001203 0.0050880.000060 -0.0011260.001739 -0.0060880.006849 -0.0101130.015228 -0.0138460.026702 -0.0171710.041137 -0.0200840.058410 -0.022515


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CAL2263mActual

x/c y/c1.000000 0.0012960.994630 0.0024220.985146 0.0048940.973912 0.0075960.958256 0.0107750.943282 0.0140670.927547 0.0178540.906763 0.0227790.886441 0.0275670.864121 0.0326240.839365 0.0380560.812579 0.0436280.785601 0.0487490.756512 0.0538870.723630 0.0595020.691715 0.064553

0.656336 0.0697210.619807 0.0746230.583816 0.0787500.546289 0.0825690.509890 0.0857080.468529 0.0884600.430722 0.0901240.389649 0.0910920.351970 0.0914480.317628 0.0903670.279125 0.0886200.242057 0.0863630.204761 0.0826890.175002 0.0785280.145093 0.0730290.113740 0.0656150.086148 0.0574010.064442 0.0495030.054604 0.0453940.042180 0.0394720.033986 0.0350510.022685 0.0277770.017442 0.0239690.012155 0.0195990.008230 0.0159830.006640 0.0143420.004789 0.0123450.003525 0.0107360.002493 0.0089820.002014 0.0081030.000975 0.0058310.000283 0.0035180.000031 -0.0011790.000142 -0.0021550.000528 -0.0044890.001006 -0.0060310.002672 -0.0087520.008222 -0.0130690.014915 -0.0164410.021401 -0.0184300.027763 -0.0202820.042398 -0.0231590.060824 -0.0250460.083627 -0.0265100.110004 -0.0276990.138018 -0.0285500.167135 -0.0290300.199455 -0.0291620.232209 -0.0289490.269210 -0.0282810.306838 -0.0270350.342818 -0.0258120.380887 -0.023951

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CAL4014lTrue

x/c y/c1.000000 0.0000000.998201 -0.0001090.992738 -0.0004140.983486 -0.0008170.970390 -0.0010610.953644 -0.0008630.933525 -0.0001370.910244 0.0011660.884054 0.0031170.855217 0.0057440.824066 0.0090980.790972 0.0131440.756257 0.0178110.720292 0.0230110.683445 0.0285830.646013 0.0343320.608272 0.0400770.570534 0.0455880.532839 0.0505550.495176 0.0549410.457797 0.0587760.420864 0.0619950.384557 0.0646090.349092 0.066528


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0.984447 -0.0014170.992950 -0.0006060.998210 -0.0001371.000000 -0.000000

CAL4014lActual

x/c y/c1.000000 0.0004260.996887 0.0002760.992416 0.0000870.985943 -0.0001040.976097 -0.0004800.964889 -0.0007820.951172 -0.0009930.936634 -0.0008020.918927 -0.0000040.897579 0.0014040.878131 0.0029940.855003 0.0053890.831823 0.0080950.800774 0.0121290.772710 0.0160490.743372 0.0203370.711941 0.0250420.675659 0.0307270.646856 0.0350040.612683 0.0401450.578234 0.0451500.541021 0.0503830.501649 0.0551790.464273 0.0589170.426141 0.0620830.388574 0.0646800.349332 0.0667300.310652 0.0680730.269851 0.0684820.236071 0.0679130.202309 0.0662510.172313 0.0639200.138541 0.0600800.108009 0.0548600.085705 0.0499240.060191 0.0428960.051739 0.0401120.041159 0.0359190.034063 0.0326570.025916 0.0283540.020025 0.0248480.015628 0.0218870.013620 0.0203400.011483 0.018493

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1.000000 -0.000683

E387True

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E387 (E)Actual

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MA409True

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MA409Actual

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NACA 43012ATrue

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NACA 43012AActual

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S1223True

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S1223Actual

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S8064True

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S8064Actual

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S9000True

x/c y/c1.000000 0.0000000.999170 0.0001100.996710 0.0005000.992700 0.0012100.987230 0.0022500.980400 0.0036100.972290 0.0052300.962940 0.0070500.952400 0.0090000.940650 0.0110300.927680 0.0131400.913520 0.0153600.898220 0.0176800.881840 0.0201000.864430 0.0226100.846060 0.0252000.826770 0.027850

0.806640 0.0305600.785740 0.0333100.764120 0.0360800.741850 0.0388500.719020 0.0416000.695660 0.0442800.671850 0.0469100.647650 0.0494600.623130 0.0519100.598360 0.0542200.573380 0.0564100.548280 0.0584500.523120 0.0603300.497960 0.0620200.472860 0.0635100.447880 0.0648100.423110 0.0659000.398590 0.0667400.374360 0.0673500.350500 0.0677200.327090 0.0678500.304150 0.0677100.281730 0.0673000.259890 0.0666500.238710 0.0657300.218200 0.0645400.198410 0.0630800.179390 0.0613800.161200 0.0594200.143860 0.0571900.127380 0.0547200.111800 0.0520400.097190 0.0491300.083550 0.0459900.070870 0.0426500.059190 0.0391600.048560 0.0355300.038960 0.0317400.030390 0.0278500.022870 0.0239200.016430 0.0199700.011080 0.0159700.006750 0.0120200.003450 0.0082000.001230 0.0046300.000130 0.0013600.000200 -0.0013900.002030 -0.0036800.005370 -0.0060000.010060 -0.0082700.016060 -0.0104200.023360 -0.0124300.031940 -0.014310


0.041770 -0.0160300.052840 -0.0175800.065120 -0.0189700.078590 -0.0201800.093220 -0.0212200.108960 -0.0220800.125800 -0.0227600.143700 -0.0232600.162600 -0.0236000.182460 -0.0237800.203250 -0.0237900.224890 -0.0236700.247350 -0.0234000.270550 -0.0230100.294450 -0.0224900.318960 -0.0218800.344030 -0.0211700.369580 -0.0203700.395530 -0.0195100.421830 -0.0185800.448380 -0.0175900.475100 -0.0165700.501930 -0.0155100.528770 -0.0144200.555530 -0.0133100.582150 -0.0121700.608530 -0.0109900.634590 -0.0097700.660300 -0.0084400.685670 -0.0070700.710620 -0.0057000.735070 -0.0043800.758930 -0.0031300.782120 -0.0019700.804570 -0.0009200.826180 0.0000000.846870 0.0007900.866560 0.0014200.885180 0.0019100.902650 0.0022400.918890 0.0024200.933820 0.0024500.947380 0.0023400.959500 0.0021200.970110 0.0018100.979170 0.0014300.986630 0.0010400.992470 0.0006500.996650 0.0003200.999160 0.0000901.000000 0.000000

S9000Actual

x/c y/c1.000000 0.0019970.996908 0.0022570.992925 0.0027510.987128 0.0037020.981697 0.0046370.968688 0.0068170.953986 0.0091730.936326 0.0119480.918382 0.0147140.897465 0.0177840.872825 0.0213140.848942 0.0245960.821115 0.0283510.793419 0.0319490.763059 0.0357230.731859 0.0394710.699832 0.0432360.664494 0.0470810.629255 0.0506930.592803 0.0540970.556117 0.0572030.516344 0.0601590.476361 0.0626930.438549 0.0645840.399237 0.0661080.359858 0.0671080.323605 0.0673750.286039 0.0669720.248737 0.0657790.212779 0.0637500.181019 0.0610670.152976 0.0579360.123012 0.0535920.095021 0.0483270.072438 0.0428600.051133 0.0363090.033022 0.0289680.016622 0.0200570.010298 0.0153400.007029 0.0123250.005081 0.0102640.003655 0.0084850.002556 0.0069100.001825 0.0057140.001209 0.0045240.000552 0.0030260.000197 -0.0014430.000945 -0.0029350.001978 -0.004083

0.003126 -0.0050510.006861 -0.0073900.012715 -0.0099110.017810 -0.0116190.023587 -0.0132480.031222 -0.0149830.037716 -0.0161890.044828 -0.0173160.065449 -0.0198380.089032 -0.0217760.136463 -0.0238750.170576 -0.0245870.201230 -0.0247630.236176 -0.0245590.269418 -0.0241130.306761 -0.0233820.344938 -0.0223900.382486 -0.0212800.419831 -0.0200640.460024 -0.0186030.496468 -0.0171380.541057 -0.0151100.584925 -0.0131420.615615 -0.0118560.655208 -0.0100310.688406 -0.0084570.722878 -0.0067810.756278 -0.0050200.785123 -0.0035060.814652 -0.0020070.840403 -0.0010090.871503 -0.0000630.891031 0.0005070.910600 0.0008140.929921 0.0007890.945259 0.0006530.960046 0.0004170.975675 0.0000490.987131 -0.0000090.991945 -0.0000611.000000 -0.000141

Appendix B

Tabulated Drag Polar Data

Appendix B contains all of the polar data seen in Chapter 4. The data presented in this appendix is identifiedby airfoil name, figure number, and run number. The same data along with all eight spanwise Cd valuesused to calculate the average Cd is available upon request. As a note, the flap deflections are defined withthe following notation: “p” is positive and “n” is negative. For example, the AG40d-02r with a −10 deg flapwould have identified as “AG40d-02r fn10”.


AG12Fig. 4.3

Run: bb05707 interpRe = 39617.3

α Cl Cd

-3.05 -0.090 0.0200-1.98 0.010 0.0123-1.49 0.059 0.0131-0.96 0.117 0.0136-0.45 0.185 0.01280.02 0.227 0.01321.07 0.335 0.01582.10 0.431 0.01163.15 0.596 0.01754.24 0.733 0.02485.25 0.803 0.02716.31 0.889 0.03957.24 0.960 0.06068.29 1.024 0.06429.28 1.058 0.1083

10.31 1.027 0.119311.26 0.988 0.1743


α Cl Cd

-3.01 -0.105 0.0138-1.97 -0.028 0.0127-1.54 0.012 0.0112-1.03 0.059 0.0105-0.46 0.116 0.01220.04 0.159 0.01161.12 0.262 0.01192.11 0.393 0.01483.21 0.541 0.01594.16 0.612 0.01845.24 0.705 0.02046.23 0.795 0.02487.24 0.884 0.02758.23 0.947 0.04029.30 0.997 0.0676

10.27 1.055 0.134011.32 0.993 0.1693

Run: ts05641 interpRe = 79922.2

α Cl Cd

-3.05 -0.118 0.0139-2.03 -0.038 0.0116-1.48 0.004 0.0096-0.91 0.057 0.0095

-0.46 0.100 0.01010.06 0.145 0.01031.09 0.251 0.01102.17 0.389 0.01253.16 0.500 0.01434.22 0.592 0.01605.29 0.683 0.01716.20 0.758 0.02037.25 0.847 0.02668.22 0.915 0.03259.23 0.980 0.0496

10.24 1.018 0.134411.27 0.975 0.1670


α Cl Cd

-3.04 -0.124 0.0131-2.00 -0.031 0.0095-1.42 0.016 0.0086-0.92 0.059 0.0093-0.43 0.110 0.00980.06 0.159 0.00831.07 0.271 0.01102.21 0.405 0.01173.17 0.490 0.01314.11 0.569 0.01515.17 0.670 0.01806.20 0.758 0.02117.30 0.850 0.02598.25 0.924 0.03679.24 0.989 0.0590

10.25 1.027 0.135511.23 0.964 0.1641

Run: jb05643 interpRe = 149943.1

α Cl Cd

-3.02 -0.137 0.0116-1.99 -0.040 0.0091-1.45 0.009 0.0081-0.90 0.061 0.0074-0.43 0.109 0.00840.06 0.156 0.00881.08 0.264 0.01012.11 0.382 0.00883.14 0.484 0.01024.12 0.586 0.01185.20 0.688 0.01386.21 0.771 0.01697.29 0.866 0.02178.27 0.942 0.02889.21 1.000 0.0422

10.19 1.059 0.133211.17 1.031 0.1639


α Cl Cd

-3.04 -0.131 0.0107-2.02 -0.032 0.0086-1.44 0.020 0.0076-0.96 0.063 0.0068-0.36 0.114 0.00810.05 0.147 0.00801.06 0.274 0.00822.12 0.385 0.00823.11 0.486 0.00924.14 0.595 0.01115.14 0.702 0.01306.19 0.795 0.01607.25 0.889 0.01968.27 0.962 0.02799.26 1.027 0.0416

10.24 1.068 0.132011.24 1.087 0.1609


α Cl Cd

-3.07 -0.129 0.0092-2.03 -0.021 0.0077-1.38 0.036 0.0070-0.97 0.062 0.0067-0.45 0.122 0.00740.06 0.182 0.00581.07 0.309 0.00652.11 0.413 0.00753.17 0.522 0.00874.13 0.622 0.00995.18 0.722 0.01176.18 0.819 0.01427.25 0.919 0.01878.24 1.001 0.02609.29 1.058 0.0399

10.27 1.101 0.1331

AG16Fig. 4.7


α Cl Cd

-3.10 -0.111 0.0161-2.05 0.007 0.0136

Appendix B: Tabulated Drag Polar Data 321

-1.45 0.058 0.0138-1.02 0.106 0.0134-0.48 0.173 0.01510.03 0.234 0.01431.11 0.365 0.01842.07 0.519 0.01583.18 0.661 0.01974.25 0.801 0.03085.17 0.873 0.03396.28 0.968 0.04317.27 1.035 0.06228.32 1.091 0.05129.31 1.119 0.0710

10.32 1.143 0.146111.32 1.115 0.1852


α Cl Cd

-3.03 -0.059 0.0143-1.95 0.037 0.0115-1.48 0.086 0.0120-0.96 0.133 0.0126-0.49 0.176 0.01260.05 0.221 0.01521.12 0.328 0.01552.12 0.493 0.01783.23 0.604 0.01704.19 0.693 0.02085.25 0.767 0.02686.23 0.842 0.02887.27 0.918 0.03088.26 0.987 0.03879.23 1.036 0.0600

10.27 1.016 0.114111.22 1.009 0.1644


α Cl Cd

-3.03 -0.103 0.0136-2.04 -0.017 0.0088-1.42 0.036 0.0104-0.96 0.091 0.0100-0.47 0.142 0.01160.05 0.195 0.01181.18 0.319 0.01452.13 0.480 0.01373.15 0.570 0.01534.13 0.650 0.01765.17 0.728 0.01866.20 0.813 0.02167.23 0.896 0.0284

8.23 0.953 0.03649.24 1.027 0.0499

10.24 1.005 0.132011.24 1.047 0.1684


α Cl Cd

-3.02 -0.091 0.0120-1.96 -0.007 0.0095-1.44 0.035 0.0085-1.00 0.085 0.0108-0.41 0.144 0.01150.10 0.195 0.00911.08 0.345 0.01412.15 0.482 0.01163.17 0.570 0.01304.17 0.658 0.01475.17 0.753 0.01646.22 0.846 0.02047.25 0.924 0.02578.28 0.990 0.03309.29 1.039 0.0475

10.22 1.057 0.130811.26 1.003 0.1645

Run: bm05691 interpRe = 150266.5

α Cl Cd

-3.07 -0.100 0.0104-1.92 0.008 0.0087-1.50 0.043 0.0082-0.85 0.102 0.0077-0.40 0.143 0.00850.06 0.204 0.00961.08 0.338 0.00912.19 0.446 0.00963.16 0.532 0.01064.16 0.634 0.01265.23 0.737 0.01466.22 0.831 0.01827.21 0.937 0.02498.23 1.009 0.03259.30 1.065 0.0447

10.16 1.065 0.143911.28 1.040 0.1734


α Cl Cd

-3.01 -0.085 0.0093-1.93 0.019 0.0082-1.41 0.061 0.0076

-0.99 0.104 0.0079-0.46 0.177 0.00900.01 0.229 0.00851.09 0.348 0.00772.12 0.451 0.00853.15 0.585 0.01024.17 0.691 0.01205.20 0.796 0.01416.21 0.884 0.01747.28 0.973 0.02238.25 1.043 0.02809.32 1.096 0.0404

10.26 1.107 0.145311.16 1.085 0.1720


α Cl Cd

-3.03 -0.078 0.0084-1.84 0.038 0.0073-1.47 0.080 0.0073-0.99 0.136 0.0067-0.46 0.199 0.00560.02 0.278 0.00621.08 0.383 0.00712.09 0.494 0.00803.10 0.598 0.00934.15 0.703 0.01075.21 0.809 0.01276.22 0.904 0.01577.30 0.998 0.01998.23 1.068 0.02459.29 1.128 0.0340

10.14 1.140 0.132011.22 1.121 0.1643

AG24Fig. 4.11

Run: bb05531 editRe = 60000.9

α Cl Cd

-5.90 -0.355 0.0339-4.87 -0.271 0.0237-3.82 -0.200 0.0191-2.83 -0.121 0.0156-1.83 -0.049 0.0127-1.30 -0.018 0.0148-0.76 0.041 0.0154-0.33 0.085 0.01560.22 0.147 0.01560.73 0.188 0.0181


1.26 0.242 0.01831.77 0.334 0.01952.34 0.440 0.01952.86 0.522 0.02053.30 0.566 0.02124.36 0.640 0.02395.40 0.725 0.02476.40 0.803 0.02657.42 0.889 0.02648.42 0.959 0.03039.44 1.005 0.0394

10.50 1.045 0.050511.46 1.073 0.0742

Run: bm05533 editRe = 79747.8

α Cl Cd

-5.86 -0.367 0.0342-4.84 -0.292 0.0233-3.86 -0.214 0.0179-2.86 -0.126 0.0138-1.84 -0.045 0.0122-1.26 0.004 0.0121-0.78 0.058 0.0106-0.27 0.110 0.01210.20 0.162 0.01210.70 0.217 0.01501.26 0.320 0.01611.82 0.403 0.01682.39 0.482 0.01732.83 0.525 0.01783.33 0.567 0.01914.36 0.649 0.01915.41 0.741 0.01986.46 0.825 0.02107.42 0.912 0.02318.43 0.986 0.02889.47 1.025 0.0367

10.44 1.068 0.048611.49 1.071 0.0763

Run: bm05535 editRe = 99958.8

α Cl Cd

-5.88 -0.380 0.0301-4.82 -0.289 0.0213-3.85 -0.203 0.0157-2.88 -0.114 0.0130-1.75 -0.027 0.0116-1.35 0.013 0.0105-0.74 0.079 0.0103-0.28 0.127 0.01320.21 0.212 0.0134

0.77 0.302 0.01341.28 0.379 0.01401.84 0.446 0.01392.37 0.496 0.01462.86 0.536 0.01463.31 0.571 0.01454.38 0.665 0.01645.41 0.757 0.01656.44 0.848 0.01947.41 0.934 0.02238.44 0.995 0.02799.50 1.049 0.0372

10.46 1.086 0.051111.49 1.106 0.0872

Run: bb05537 edit1Re = 150183.0

α Cl Cd

-5.98 -0.382 0.0288-4.88 -0.283 0.0184-3.88 -0.193 0.0138-2.90 -0.103 0.0113-1.80 -0.005 0.0106-1.28 0.061 0.0105-0.74 0.133 0.0099-0.26 0.210 0.00980.28 0.279 0.00930.81 0.334 0.00971.29 0.385 0.01001.83 0.439 0.01012.32 0.484 0.01062.82 0.536 0.01123.36 0.589 0.01134.41 0.694 0.01275.43 0.795 0.01436.44 0.888 0.01607.45 0.983 0.01958.47 1.047 0.02479.52 1.096 0.0322

10.44 1.131 0.042611.49 1.144 0.0629

Run: bm05539 editRe = 199977.1

α Cl Cd

-5.84 -0.378 0.0244-4.88 -0.295 0.0166-3.85 -0.201 0.0131-2.76 -0.103 0.0105-1.85 0.032 0.0093-1.31 0.113 0.0093-0.74 0.181 0.0082-0.22 0.238 0.0073

0.29 0.290 0.00780.71 0.335 0.00781.30 0.399 0.00861.80 0.443 0.00902.27 0.492 0.00922.79 0.545 0.00953.35 0.597 0.01024.39 0.706 0.01125.35 0.798 0.01266.42 0.900 0.01487.41 0.983 0.01808.47 1.055 0.02339.50 1.102 0.0303

10.51 1.130 0.040711.45 1.126 0.0604

Run: bm05541 editRe = 299824.5

α Cl Cd

-5.87 -0.372 0.0215-4.88 -0.287 0.0146-3.79 -0.180 0.0109-2.84 -0.056 0.0090-1.79 0.078 0.0075-1.32 0.129 0.0069-0.75 0.185 0.0067-0.30 0.232 0.00640.25 0.286 0.00640.79 0.342 0.00661.21 0.386 0.00691.88 0.456 0.00732.31 0.500 0.00782.83 0.553 0.00813.35 0.601 0.00864.40 0.708 0.00985.45 0.813 0.01146.40 0.897 0.01337.44 0.991 0.01618.45 1.066 0.02059.48 1.121 0.0268

10.49 1.163 0.035911.42 1.164 0.0575

Run: bm05543 editRe = 400101.5

α Cl Cd

-5.82 -0.376 0.0179-4.86 -0.286 0.0128-3.85 -0.165 0.0091-2.79 -0.033 0.0075-1.73 0.082 0.0069-1.28 0.131 0.0065-0.74 0.184 0.0063


-0.24 0.230 0.00610.26 0.280 0.00600.73 0.330 0.00611.24 0.382 0.00661.80 0.441 0.00682.34 0.492 0.00722.86 0.548 0.00753.38 0.601 0.00804.36 0.703 0.00915.38 0.798 0.01056.43 0.901 0.01247.45 0.987 0.01498.40 1.062 0.01869.46 1.121 0.0240

10.45 1.169 0.031311.49 1.179 0.0543

AG35-rFig. 4.15

Run: bm05388 editRe = 59903.7

α Cl Cd

-5.90 -0.375 0.0441-4.90 -0.298 0.0284-3.87 -0.225 0.0222-2.89 -0.153 0.0187-1.90 -0.082 0.0121-1.37 -0.055 0.0140-0.79 0.003 0.0149-0.36 0.051 0.01430.20 0.096 0.01440.60 0.146 0.01721.21 0.251 0.02291.78 0.371 0.02062.30 0.428 0.02242.79 0.471 0.02293.33 0.518 0.02604.32 0.594 0.02635.36 0.675 0.02636.38 0.760 0.02787.38 0.836 0.02998.40 0.901 0.03029.42 0.962 0.0313

10.44 1.002 0.039411.43 1.035 0.0518

Run: bb05396 editRe = 80133.6

α Cl Cd

-5.91 -0.392 0.0539-4.87 -0.300 0.0311

-3.86 -0.223 0.0225-2.91 -0.159 0.0181-1.91 -0.087 0.0134-1.32 -0.037 0.0124-0.84 0.009 0.0132-0.33 0.061 0.01310.24 0.145 0.01470.75 0.237 0.01511.27 0.322 0.01541.80 0.386 0.01552.33 0.430 0.01672.76 0.464 0.01793.35 0.509 0.01864.30 0.586 0.02005.31 0.664 0.02006.42 0.755 0.02107.41 0.836 0.02168.39 0.910 0.02469.44 0.975 0.0291

10.42 1.013 0.036411.46 1.044 0.0465

Run: bb05391 edit1Re = 100014.3

α Cl Cd

-5.95 -0.369 0.0444-4.95 -0.286 0.0258-3.89 -0.214 0.0204-2.88 -0.143 0.0165-1.87 -0.066 0.0135-1.36 -0.025 0.0124-0.88 0.024 0.0130-0.35 0.113 0.01430.29 0.222 0.01380.77 0.297 0.01371.24 0.353 0.01371.76 0.402 0.01472.28 0.445 0.01482.79 0.490 0.01503.35 0.545 0.01564.33 0.637 0.01615.33 0.730 0.01766.42 0.830 0.01967.40 0.914 0.02088.47 0.997 0.02489.44 1.054 0.0286

10.47 1.096 0.038711.45 1.124 0.0515

Run: bm05393 fullRe = 149798.4

α Cl Cd

-5.94 -0.390 0.0429-4.95 -0.307 0.0242-3.89 -0.226 0.0176-2.84 -0.141 0.0142-1.79 -0.008 0.0114-1.31 0.068 0.0100-0.81 0.142 0.0099-0.23 0.216 0.01010.26 0.275 0.01040.73 0.315 0.01041.23 0.352 0.01061.80 0.405 0.01072.25 0.443 0.01112.74 0.488 0.01153.34 0.547 0.01184.32 0.646 0.01295.33 0.747 0.01436.36 0.842 0.01557.39 0.927 0.01818.44 1.020 0.02169.45 1.080 0.0269

10.44 1.116 0.035011.42 1.138 0.0443

Run: bm05398 fullRe = 199881.3

α Cl Cd

-5.94 -0.420 0.0480-4.96 -0.326 0.0225-3.88 -0.243 0.0160-2.86 -0.125 0.0127-1.80 0.020 0.0100-1.31 0.076 0.0086-0.78 0.137 0.0085-0.28 0.191 0.00850.24 0.246 0.00870.72 0.289 0.00871.26 0.340 0.00891.73 0.383 0.00912.24 0.431 0.00942.80 0.489 0.00983.33 0.543 0.01034.29 0.641 0.01155.30 0.737 0.01256.38 0.842 0.01417.41 0.939 0.01648.41 1.021 0.01949.42 1.081 0.0243

10.43 1.127 0.031811.48 1.143 0.0426


Run: bb05402 editRe = 299680.8

α Cl Cd

-5.90 -0.413 0.0444-4.84 -0.316 0.0183-3.84 -0.211 0.0127-2.87 -0.079 0.0108-1.80 0.038 0.0090-1.34 0.081 0.0078-0.80 0.125 0.0069-0.29 0.196 0.00690.25 0.249 0.00710.71 0.295 0.00731.20 0.347 0.00771.73 0.402 0.00802.26 0.453 0.00842.77 0.504 0.00863.36 0.568 0.00904.37 0.668 0.00965.34 0.759 0.01076.38 0.863 0.01277.43 0.955 0.01488.43 1.038 0.01769.41 1.102 0.0224

10.50 1.152 0.029411.47 1.171 0.0401

AG40d-02r fp0Fig. 4.19


α Cl Cd

-3.02 -0.032 0.0130-1.99 0.059 0.0137-1.50 0.101 0.0163-1.00 0.140 0.0166-0.39 0.187 0.01910.07 0.227 0.02051.07 0.402 0.02282.11 0.539 0.02223.20 0.619 0.02444.17 0.699 0.02335.23 0.782 0.02676.20 0.854 0.02807.21 0.915 0.03408.25 0.960 0.04099.31 0.995 0.0609

10.26 1.020 0.1281


α Cl Cd

-3.04 -0.039 0.0128-2.04 0.045 0.0137-1.46 0.105 0.0145-1.01 0.150 0.0143-0.50 0.235 0.01540.08 0.355 0.01621.09 0.493 0.01552.10 0.558 0.01473.10 0.653 0.01664.17 0.766 0.01925.24 0.852 0.01986.19 0.926 0.02507.23 0.975 0.03058.22 1.028 0.03739.25 1.059 0.0527

10.21 1.039 0.1049


α Cl Cd

-3.06 -0.038 0.0118-1.98 0.092 0.0115-1.46 0.169 0.0123-0.97 0.245 0.0129-0.45 0.322 0.01370.11 0.401 0.01371.13 0.502 0.01332.13 0.590 0.01393.18 0.690 0.01454.22 0.779 0.01595.20 0.866 0.01836.26 0.948 0.02247.23 1.005 0.02848.23 1.053 0.03779.25 1.084 0.0510

10.24 1.077 0.1159


α Cl Cd

-3.01 0.018 0.0108-1.96 0.155 0.0104-1.40 0.240 0.0103-0.91 0.307 0.0102-0.49 0.351 0.01030.07 0.406 0.01061.14 0.503 0.01112.10 0.594 0.01123.21 0.690 0.01214.18 0.783 0.0137

5.25 0.875 0.01626.25 0.955 0.02027.21 1.011 0.02558.26 1.063 0.03409.32 1.095 0.0452

10.27 1.118 0.0755


α Cl Cd

-2.97 0.062 0.0109-1.94 0.172 0.0084-1.44 0.233 0.0084-0.93 0.287 0.0087-0.41 0.335 0.00880.03 0.382 0.00901.09 0.488 0.00912.14 0.586 0.00973.17 0.676 0.01084.17 0.778 0.01255.21 0.872 0.01446.23 0.957 0.01847.20 1.021 0.02308.29 1.084 0.03019.29 1.119 0.0394

10.21 1.131 0.0600


α Cl Cd

-2.97 0.073 0.0080-2.00 0.168 0.0072-1.45 0.231 0.0069-0.99 0.275 0.0072-0.43 0.338 0.00740.04 0.388 0.00741.10 0.496 0.00782.13 0.598 0.00863.18 0.703 0.00984.15 0.799 0.01125.15 0.893 0.01336.20 0.980 0.01667.22 1.057 0.01988.31 1.122 0.02679.24 1.158 0.0350

10.28 1.172 0.0559


α Cl Cd

-2.99 0.073 0.0075-1.97 0.177 0.0065-1.47 0.229 0.0062


-0.93 0.286 0.0059-0.41 0.345 0.00590.08 0.394 0.00641.13 0.501 0.00712.12 0.599 0.00793.18 0.708 0.00904.12 0.801 0.01045.25 0.903 0.01256.30 0.995 0.01527.26 1.074 0.01868.29 1.149 0.02319.35 1.195 0.0294

10.27 1.229 0.0430

AG40d-02r fn2Fig. 4.22


α Cl Cd

-3.08 -0.085 0.0131-2.06 -0.010 0.0125-1.52 0.037 0.0140-1.01 0.076 0.0147-0.49 0.121 0.01570.00 0.171 0.01850.51 0.241 0.01781.07 0.331 0.01992.07 0.487 0.02043.11 0.568 0.02254.19 0.656 0.02225.14 0.736 0.02316.15 0.815 0.02477.21 0.890 0.02708.22 0.953 0.03429.21 0.996 0.0449

10.26 1.034 0.0629


α Cl Cd

-3.03 -0.097 0.0119-1.97 -0.016 0.0109-1.50 0.025 0.0122-0.98 0.085 0.0135-0.47 0.173 0.01460.04 0.264 0.01550.60 0.352 0.01461.10 0.402 0.01452.18 0.495 0.01633.18 0.580 0.01674.18 0.671 0.0181

5.17 0.761 0.01956.24 0.846 0.02217.21 0.917 0.02828.24 0.975 0.03679.28 1.022 0.0473

10.26 1.047 0.0693


α Cl Cd

-3.07 -0.111 0.0121-2.00 -0.022 0.0110-1.50 0.055 0.0114-0.97 0.130 0.0114-0.49 0.199 0.01210.07 0.273 0.01290.56 0.319 0.01291.06 0.364 0.01342.09 0.450 0.01383.13 0.538 0.01444.21 0.638 0.01535.21 0.735 0.01766.22 0.834 0.01937.21 0.905 0.02438.24 0.972 0.03239.31 1.027 0.0418

10.28 1.061 0.0627


α Cl Cd

-3.00 -0.087 0.0103-2.00 0.048 0.0109-1.46 0.129 0.0093-0.94 0.195 0.0090-0.44 0.242 0.00970.03 0.287 0.01020.61 0.337 0.01071.08 0.377 0.01092.21 0.485 0.01133.16 0.574 0.01184.24 0.677 0.01305.20 0.770 0.01466.19 0.858 0.01757.25 0.929 0.02248.23 0.998 0.02939.23 1.043 0.0390

10.26 1.072 0.0545


α Cl Cd

-3.02 -0.064 0.0096-1.98 0.057 0.0091-1.45 0.127 0.0077-0.98 0.175 0.0081-0.48 0.221 0.00820.08 0.277 0.00880.56 0.326 0.00891.14 0.380 0.00932.15 0.480 0.00983.15 0.578 0.01044.16 0.680 0.01175.23 0.782 0.01356.25 0.874 0.01627.24 0.951 0.02038.27 1.020 0.02649.26 1.071 0.0347

10.27 1.094 0.0503


α Cl Cd

-3.00 -0.049 0.0089-1.99 0.047 0.0073-1.45 0.101 0.0066-0.99 0.156 0.0069-0.43 0.216 0.00710.03 0.261 0.00720.53 0.314 0.00751.07 0.370 0.00802.11 0.475 0.00823.16 0.582 0.00914.18 0.687 0.01025.21 0.786 0.01186.23 0.881 0.01447.23 0.964 0.01788.25 1.039 0.02269.31 1.096 0.0304

10.30 1.118 0.0434


α Cl Cd

-3.03 -0.054 0.0082-2.06 0.039 0.0072-1.48 0.097 0.0066-0.98 0.152 0.0059-0.47 0.212 0.00590.03 0.267 0.00610.52 0.317 0.00631.07 0.377 0.0065


2.15 0.487 0.00723.15 0.591 0.00824.18 0.688 0.00925.15 0.783 0.01076.24 0.891 0.01317.24 0.977 0.01578.31 1.073 0.01959.34 1.130 0.0248

10.26 1.173 0.0317



α Cl Cd

-5.07 -0.127 0.0201-4.00 -0.044 0.0169-3.53 -0.012 0.0149-2.99 0.034 0.0123-2.45 0.090 0.0160-2.04 0.132 0.0152-1.45 0.190 0.0168-1.01 0.231 0.0179-0.43 0.284 0.02190.06 0.335 0.02171.11 0.530 0.02252.19 0.693 0.02443.19 0.775 0.02214.20 0.840 0.02105.24 0.909 0.02536.23 0.977 0.03027.27 1.037 0.03348.27 1.077 0.04249.24 1.101 0.0643


α Cl Cd

-5.10 -0.101 0.0178-4.06 -0.013 0.0150-3.58 0.024 0.0140-3.00 0.091 0.0120-2.51 0.155 0.0151-1.98 0.203 0.0156-1.48 0.255 0.0151-0.93 0.341 0.0163-0.46 0.419 0.01730.06 0.508 0.01761.08 0.632 0.01682.14 0.699 0.01383.16 0.805 0.0170

4.22 0.895 0.02085.23 0.968 0.02246.15 1.009 0.02827.24 1.077 0.03588.26 1.116 0.04829.24 1.129 0.0648


α Cl Cd

-5.05 -0.110 0.0177-4.08 -0.004 0.0140-3.52 0.052 0.0138-3.04 0.101 0.0135-2.45 0.167 0.0130-1.97 0.220 0.0128-1.37 0.286 0.0141-0.90 0.340 0.0148-0.44 0.408 0.01600.10 0.484 0.01581.08 0.590 0.01482.15 0.677 0.01343.19 0.767 0.01534.16 0.850 0.01725.22 0.936 0.02086.24 1.000 0.02497.29 1.054 0.03148.26 1.097 0.04269.26 1.124 0.0559


α Cl Cd

-5.03 -0.042 0.0136-4.01 0.074 0.0115-3.51 0.131 0.0114-2.97 0.186 0.0112-2.46 0.239 0.0124-1.90 0.307 0.0124-1.46 0.362 0.0118-0.89 0.434 0.0120-0.44 0.482 0.01220.14 0.539 0.01171.11 0.633 0.01112.16 0.733 0.01153.21 0.823 0.01324.24 0.914 0.01585.21 0.989 0.01916.25 1.056 0.02467.21 1.105 0.03198.27 1.142 0.04149.26 1.167 0.0579


α Cl Cd

-5.06 -0.025 0.0121-4.02 0.090 0.0106-3.51 0.144 0.0100-2.97 0.198 0.0100-2.46 0.254 0.0097-1.97 0.300 0.0096-1.44 0.364 0.0097-0.98 0.410 0.0095-0.39 0.467 0.00940.06 0.512 0.00931.15 0.618 0.00922.12 0.709 0.01023.14 0.806 0.01204.17 0.894 0.01415.22 0.985 0.01706.24 1.058 0.02187.26 1.119 0.02778.25 1.158 0.03579.26 1.181 0.0484


α Cl Cd

-5.03 -0.006 0.0103-4.05 0.100 0.0089-3.56 0.147 0.0085-3.00 0.205 0.0083-2.49 0.259 0.0080-1.96 0.314 0.0078-1.45 0.368 0.0076-0.95 0.422 0.0074-0.44 0.476 0.00710.11 0.530 0.00741.12 0.627 0.00822.22 0.731 0.00953.17 0.822 0.01124.24 0.913 0.01315.21 1.000 0.01616.24 1.083 0.01957.26 1.148 0.02508.34 1.188 0.03309.26 1.211 0.0456




α Cl Cd

-6.05 -0.138 0.0232-5.02 -0.038 0.0196-4.51 0.006 0.0174-4.00 0.044 0.0169-3.50 0.104 0.0163-2.96 0.170 0.0174-2.47 0.223 0.0188-1.97 0.277 0.0187-1.38 0.338 0.0214-0.88 0.392 0.02130.09 0.510 0.02421.13 0.702 0.02572.19 0.849 0.02343.13 0.935 0.02064.27 1.015 0.02425.23 1.075 0.02716.24 1.128 0.03467.25 1.177 0.04208.27 1.205 0.0567


α Cl Cd

-7.09 -0.179 0.0263-6.09 -0.078 0.0193-5.04 0.016 0.0165-4.54 0.068 0.0158-4.04 0.109 0.0155-3.58 0.146 0.0157-3.05 0.194 0.0159-2.51 0.250 0.0156-1.94 0.310 0.0166-0.93 0.422 0.01950.06 0.615 0.01971.15 0.750 0.01882.18 0.837 0.01683.20 0.942 0.01864.20 1.015 0.02195.25 1.092 0.02726.22 1.144 0.03377.25 1.184 0.04338.27 1.213 0.0527


α Cl Cd

-6.09 -0.100 0.0200-5.08 0.022 0.0153-4.54 0.078 0.0148-4.03 0.125 0.0144-3.50 0.179 0.0143-3.03 0.230 0.0142-2.53 0.279 0.0149-1.96 0.340 0.0151-1.42 0.397 0.0159-0.91 0.448 0.01720.10 0.590 0.01841.10 0.732 0.01572.16 0.828 0.01533.09 0.920 0.01604.20 1.010 0.01845.22 1.086 0.02286.22 1.132 0.02997.20 1.172 0.03768.21 1.209 0.0485


α Cl Cd

-7.02 -0.069 0.0179-6.02 0.018 0.0147-5.52 0.070 0.0134-5.00 0.123 0.0129-4.50 0.179 0.0127-3.98 0.236 0.0127-3.45 0.287 0.0127-2.94 0.340 0.0125-1.92 0.445 0.0139-0.89 0.563 0.01340.14 0.671 0.01281.13 0.769 0.01192.15 0.856 0.01273.20 0.947 0.01514.21 1.027 0.01825.20 1.091 0.02336.23 1.142 0.03027.24 1.186 0.03868.27 1.213 0.0515


α Cl Cd

-6.08 0.009 0.0130-5.05 0.117 0.0119-4.49 0.174 0.0112-4.02 0.222 0.0110

-3.46 0.280 0.0110-2.99 0.326 0.0111-2.49 0.370 0.0112-1.97 0.422 0.0119-1.44 0.480 0.0117-0.91 0.538 0.01140.06 0.629 0.01011.11 0.739 0.00982.13 0.825 0.01113.17 0.914 0.01354.19 1.003 0.01615.20 1.077 0.02046.20 1.133 0.02537.21 1.180 0.03228.26 1.216 0.0418


α Cl Cd

-6.07 0.008 0.0111-4.99 0.119 0.0097-4.48 0.182 0.0093-4.05 0.226 0.0090-3.51 0.280 0.0089-2.93 0.345 0.0091-2.41 0.403 0.0092-1.92 0.454 0.0089-1.44 0.500 0.0083-0.92 0.555 0.00800.13 0.665 0.00781.14 0.757 0.00902.18 0.847 0.01083.19 0.945 0.01274.23 1.030 0.01525.23 1.103 0.01906.25 1.162 0.02327.28 1.222 0.03048.26 1.247 0.0395

AG40d-02r fn15Fig. 4.31


α Cl Cd

0.01 -0.461 0.01851.04 -0.278 0.02132.02 -0.198 0.02283.09 -0.107 0.02494.07 -0.045 0.02505.10 0.028 0.02275.62 0.067 0.0215


6.11 0.102 0.02026.66 0.151 0.01947.13 0.197 0.01917.68 0.248 0.01938.19 0.295 0.01998.66 0.341 0.01999.23 0.398 0.0211

10.23 0.496 0.024411.32 0.591 0.029612.23 0.658 0.0406

AG40d-02r fn10Fig. 4.34


α Cl Cd

-2.02 -0.356 0.0144-1.04 -0.277 0.01270.01 -0.120 0.01371.10 -0.012 0.01471.57 0.023 0.01582.09 0.069 0.01612.62 0.117 0.01743.14 0.163 0.01803.63 0.210 0.01804.15 0.259 0.01774.60 0.298 0.01595.18 0.357 0.01576.15 0.455 0.01647.21 0.551 0.01788.23 0.632 0.02049.28 0.713 0.0256

10.23 0.784 0.032711.34 0.841 0.0463



α Cl Cd

-3.02 -0.237 0.0132-2.04 -0.124 0.0126-1.03 0.016 0.0108-0.46 0.093 0.01220.03 0.150 0.01180.55 0.205 0.01241.04 0.248 0.01321.62 0.299 0.01372.12 0.344 0.0143

2.63 0.389 0.01623.13 0.434 0.01594.21 0.530 0.01525.21 0.621 0.01646.27 0.723 0.01817.22 0.800 0.02088.24 0.871 0.02619.27 0.935 0.0352

10.25 0.978 0.0472



α Cl Cd

-7.14 -0.125 0.0226-6.08 -0.004 0.0183-5.53 0.058 0.0166-5.04 0.111 0.0151-4.54 0.162 0.0155-4.01 0.218 0.0159-3.52 0.262 0.0156-2.95 0.315 0.0147-1.96 0.399 0.0165-0.94 0.478 0.01910.11 0.651 0.01921.17 0.782 0.01662.13 0.875 0.01493.18 0.953 0.01754.22 1.051 0.02065.20 1.106 0.02576.27 1.151 0.03277.27 1.196 0.04278.22 1.224 0.0557

AG40d-02r fp10Fig. 4.43


α Cl Cd

-9.07 0.012 0.0297-8.09 0.111 0.0239-7.59 0.159 0.0226-7.07 0.207 0.0221-6.59 0.250 0.0216-6.06 0.294 0.0206-5.52 0.334 0.0213-5.01 0.380 0.0218-4.53 0.402 0.0233

-3.96 0.440 0.0239-2.96 0.511 0.0240-1.94 0.578 0.0251-0.96 0.667 0.02830.05 0.825 0.02921.14 1.009 0.02022.18 1.077 0.01963.21 1.146 0.02394.25 1.185 0.02925.21 1.207 0.03756.20 1.240 0.04747.24 1.264 0.0639

AG40d-02r fp15Fig. 4.46


α Cl Cd

-10.18 -0.303 0.1046-9.10 0.064 0.0371-8.57 0.127 0.0344-8.13 0.170 0.0329-7.58 0.230 0.0318-7.09 0.293 0.0315-6.55 0.358 0.0333-5.99 0.392 0.0341-4.98 0.473 0.0356-4.00 0.551 0.0352-2.97 0.636 0.0365-1.94 0.724 0.0368-0.85 0.798 0.03860.15 0.979 0.04181.17 1.214 0.02312.23 1.302 0.02753.22 1.316 0.03514.22 1.321 0.04345.26 1.350 0.04996.18 1.389 0.0627

AG40d-02r fp20Fig. 4.49


α Cl Cd

-10.43 -0.427 0.1143-9.32 -0.086 0.0629-8.23 0.178 0.0525-7.23 0.309 0.0513-6.23 0.406 0.0508


-5.18 0.508 0.0462-4.21 0.610 0.0515-3.12 0.713 0.0577-2.11 0.803 0.0656-1.10 0.874 0.0681-0.09 0.939 0.06781.00 0.966 0.08411.96 1.249 0.06193.05 1.340 0.06314.04 1.351 0.08165.04 1.392 0.0774



α Cl Cd

-3.07 -0.028 0.0149-2.00 0.081 0.0166-1.54 0.144 0.0171-0.98 0.201 0.0184-0.50 0.252 0.01840.03 0.342 0.01970.60 0.449 0.02211.09 0.539 0.02462.09 0.664 0.02073.11 0.753 0.02244.17 0.837 0.02435.18 0.922 0.02706.18 0.986 0.02797.22 1.053 0.03338.23 1.098 0.04559.25 1.136 0.0658

10.20 1.161 0.1237


α Cl Cd

-3.05 -0.015 0.0116-1.99 0.093 0.0120-1.47 0.162 0.0125-0.93 0.254 0.0131-0.44 0.327 0.01500.11 0.412 0.01350.60 0.461 0.01311.11 0.505 0.01292.18 0.601 0.01333.18 0.687 0.01514.16 0.779 0.01585.20 0.873 0.01846.21 0.946 0.0221

7.23 1.016 0.02908.24 1.061 0.03959.29 1.099 0.0521

10.26 1.133 0.133311.25 0.948 0.1873


α Cl Cd

-3.96 -0.043 0.0108-3.00 0.077 0.0093-2.48 0.138 0.0086-1.96 0.198 0.0084-1.48 0.252 0.0085-0.92 0.309 0.0087-0.40 0.358 0.00880.12 0.406 0.00901.12 0.508 0.00922.14 0.607 0.00993.18 0.711 0.01124.17 0.811 0.01285.27 0.904 0.01556.25 0.979 0.01977.27 1.046 0.02458.23 1.094 0.03179.30 1.133 0.0442

10.26 1.141 0.0715



α Cl Cd

-2.97 -0.085 0.0105-1.94 0.009 0.0115-1.52 0.075 0.0128-0.94 0.166 0.0125-0.43 0.242 0.01210.04 0.298 0.01240.60 0.348 0.01281.08 0.390 0.01382.15 0.480 0.01413.16 0.561 0.01484.17 0.646 0.01585.22 0.742 0.01756.20 0.826 0.02097.24 0.893 0.02648.27 0.953 0.03469.28 0.992 0.0462

10.28 1.024 0.1131


α Cl Cd

-3.00 -0.030 0.0105-1.94 0.087 0.0092-1.43 0.151 0.0080-0.97 0.200 0.0082-0.43 0.249 0.00850.08 0.295 0.00870.61 0.349 0.00911.11 0.398 0.00952.12 0.500 0.00983.17 0.607 0.01054.17 0.702 0.01195.23 0.801 0.01386.20 0.886 0.01667.26 0.967 0.02198.27 1.029 0.02799.30 1.069 0.0375

10.27 1.089 0.056011.20 1.055 0.1634


α Cl Cd

-3.04 -0.014 0.0082-1.93 0.083 0.0067-1.47 0.141 0.0067-0.90 0.195 0.0071-0.50 0.238 0.00720.09 0.302 0.00750.58 0.353 0.00771.15 0.412 0.00792.14 0.514 0.00843.16 0.616 0.00924.17 0.720 0.01065.22 0.820 0.01236.22 0.917 0.01547.23 0.996 0.01908.31 1.069 0.02529.29 1.112 0.0336

10.30 1.125 0.0529


α Cl Cd

-3.04 -0.029 0.0079-1.98 0.075 0.0068-1.47 0.127 0.0061-0.97 0.189 0.0057-0.41 0.248 0.00590.17 0.307 0.00610.58 0.352 0.0064


1.13 0.413 0.00672.15 0.517 0.00743.22 0.628 0.00844.27 0.733 0.00975.25 0.824 0.01156.27 0.920 0.01397.31 1.012 0.01708.33 1.087 0.02099.33 1.149 0.0266

10.34 1.183 0.039511.19 1.171 0.1150



α Cl Cd

-6.12 -0.138 0.0225-5.07 -0.028 0.0183-4.54 0.014 0.0170-4.03 0.057 0.0160-3.55 0.116 0.0171-2.99 0.175 0.0163-2.46 0.224 0.0193-1.97 0.279 0.0174-1.46 0.325 0.0200-0.95 0.374 0.02030.05 0.482 0.02711.11 0.686 0.02882.17 0.823 0.02333.15 0.907 0.02144.19 0.979 0.02745.20 1.037 0.03106.22 1.094 0.03657.19 1.132 0.04448.22 1.172 0.0649


α Cl Cd

-6.07 -0.087 0.0191-5.03 0.001 0.0163-4.54 0.045 0.0151-4.02 0.084 0.0147-3.50 0.129 0.0146-3.01 0.184 0.0148-2.52 0.224 0.0149-1.99 0.289 0.0157-1.46 0.352 0.0162-0.95 0.424 0.01750.11 0.608 0.0176

1.15 0.718 0.01542.18 0.814 0.01423.22 0.898 0.01684.22 0.979 0.01985.25 1.037 0.02486.19 1.081 0.03017.20 1.116 0.03958.23 1.147 0.0528


α Cl Cd

-6.06 0.022 0.0138-5.01 0.125 0.0126-4.53 0.167 0.0123-4.00 0.220 0.0122-3.48 0.265 0.0126-3.01 0.309 0.0125-2.46 0.362 0.0128-1.95 0.428 0.0124-1.43 0.490 0.0121-0.91 0.553 0.01140.13 0.648 0.01021.15 0.751 0.00982.14 0.841 0.01123.21 0.936 0.01364.18 1.017 0.01615.21 1.083 0.02076.22 1.143 0.02557.23 1.190 0.03298.25 1.214 0.0434

AG455ct-02r fn0.4Fig. 4.62

Run: bm05458 editRe = 60157.0

α Cl Cd

-3.06 -0.002 0.0121-2.02 0.096 0.0125-1.45 0.162 0.0126-1.01 0.205 0.0109-0.47 0.253 0.01210.02 0.301 0.01580.61 0.346 0.01261.09 0.390 0.01622.14 0.560 0.01403.20 0.652 0.01884.13 0.721 0.02185.24 0.807 0.02756.17 0.881 0.02807.17 0.954 0.0324

8.25 1.024 0.04059.22 1.061 0.0547

10.26 1.030 0.1461

Run: bb05461 editRe = 80065.0

α Cl Cd

-3.09 -0.046 0.0148-1.97 0.062 0.0118-1.53 0.111 0.0113-0.95 0.172 0.0089-0.44 0.218 0.01030.02 0.268 0.01200.58 0.330 0.01291.06 0.404 0.01182.15 0.521 0.01673.18 0.628 0.01704.19 0.717 0.01995.20 0.808 0.02106.16 0.894 0.02427.25 0.980 0.03288.20 1.048 0.04059.29 1.098 0.0514

10.26 1.059 0.1465

Run: bb05462 editRe = 99969.7

α Cl Cd

-3.02 -0.062 0.0144-2.05 0.019 0.0089-1.49 0.072 0.0099-1.00 0.121 0.0089-0.43 0.176 0.00850.00 0.226 0.01090.57 0.294 0.01151.08 0.388 0.01132.08 0.498 0.01193.18 0.604 0.01374.15 0.699 0.01675.18 0.803 0.01716.22 0.890 0.02127.19 0.965 0.02618.23 1.032 0.03289.23 1.074 0.0456

10.22 1.033 0.1412

Run: bm05464 editRe = 150008.1

α Cl Cd

-3.01 -0.040 0.0117-1.98 0.052 0.0088-1.53 0.101 0.0075-0.98 0.159 0.0093


-0.40 0.212 0.00930.10 0.290 0.00930.56 0.359 0.00931.12 0.419 0.00882.11 0.510 0.01063.17 0.610 0.01194.15 0.704 0.01415.22 0.807 0.01636.21 0.899 0.01917.21 0.984 0.02388.29 1.053 0.03169.26 1.097 0.0420

10.21 1.057 0.1511

Run: bm05466 editRe = 199918.4

α Cl Cd

-2.97 -0.035 0.0110-2.00 0.057 0.0085-1.49 0.103 0.0073-0.96 0.168 0.0080-0.46 0.232 0.00780.06 0.304 0.00740.54 0.357 0.00781.08 0.415 0.00812.10 0.511 0.00893.08 0.607 0.01034.16 0.711 0.01225.12 0.809 0.01416.18 0.905 0.01697.16 0.986 0.02058.24 1.059 0.02589.21 1.111 0.0342

10.29 1.134 0.0508

Run: bb05468 editRe = 300031.8

α Cl Cd

-3.06 -0.050 0.0102-1.91 0.068 0.0081-1.54 0.120 0.0073-0.92 0.193 0.0065-0.48 0.249 0.00640.05 0.313 0.00660.52 0.359 0.00681.06 0.419 0.00732.09 0.519 0.00833.15 0.624 0.00954.16 0.723 0.01085.23 0.824 0.01256.17 0.913 0.01467.24 1.003 0.01758.21 1.078 0.0224

9.30 1.137 0.0298

AG455ct-02r fn2.4Fig. 4.65

Run: bm05445 edit1Re = 59882.5

α Cl Cd

-3.05 -0.184 0.0167-2.04 -0.094 0.0102-1.49 -0.050 0.0100-0.99 -0.005 0.0115-0.46 0.043 0.01040.02 0.095 0.01180.55 0.155 0.01301.05 0.215 0.01332.10 0.378 0.01523.18 0.480 0.01824.16 0.558 0.02075.15 0.641 0.02506.16 0.719 0.02847.20 0.798 0.02858.20 0.859 0.03339.26 0.921 0.0485

10.23 0.942 0.097711.15 0.900 0.1473

Run: bm05447 edit1Re = 79923.9

α Cl Cd

-3.01 -0.154 0.0188-2.02 -0.072 0.0123-1.49 -0.032 0.0109-1.02 0.004 0.0101-0.46 0.054 0.0104-0.03 0.091 0.00960.58 0.155 0.01121.03 0.213 0.01162.09 0.338 0.01243.13 0.432 0.01464.11 0.514 0.01785.15 0.611 0.01866.20 0.697 0.02057.23 0.783 0.02428.22 0.861 0.03089.25 0.925 0.0410

10.26 0.969 0.058411.22 0.894 0.1554

Run: bm05449 edit1Re = 100144.0

α Cl Cd

-3.01 -0.161 0.0180-2.00 -0.074 0.0121-1.50 -0.037 0.0088-1.00 0.003 0.0092-0.47 0.048 0.00940.11 0.102 0.00980.55 0.144 0.00951.08 0.225 0.01042.13 0.343 0.01143.11 0.432 0.01364.13 0.520 0.01535.23 0.632 0.01656.25 0.731 0.01987.20 0.818 0.02318.25 0.904 0.02959.25 0.976 0.0385

10.28 1.025 0.060111.27 0.989 0.1629

Run: bm05451 edit1Re = 150344.7

α Cl Cd

-3.03 -0.173 0.0149-2.03 -0.078 0.0101-1.50 -0.031 0.0079-1.02 0.011 0.0075-0.46 0.069 0.00740.01 0.117 0.00820.53 0.196 0.00901.06 0.265 0.00892.15 0.375 0.01033.10 0.465 0.01134.13 0.562 0.01265.17 0.661 0.01456.18 0.753 0.01697.24 0.847 0.02028.20 0.922 0.02579.27 0.997 0.0342

10.30 1.043 0.050311.17 1.034 0.1548

Run: bb05453 edit1Re = 199907.3

α Cl Cd

-3.04 -0.159 0.0134-2.07 -0.065 0.0099-1.48 -0.009 0.0075-1.04 0.027 0.0067-0.46 0.089 0.00730.02 0.155 0.0077


0.56 0.223 0.00751.11 0.284 0.00802.06 0.372 0.00913.19 0.488 0.00994.07 0.577 0.01115.17 0.688 0.01306.19 0.791 0.01527.22 0.879 0.01858.21 0.961 0.02259.27 1.035 0.0296

10.28 1.075 0.041211.28 1.033 0.1508

Run: bm05455 edit1Re = 300459.9

α Cl Cd

-3.13 -0.172 0.0120-2.06 -0.063 0.0090-1.47 -0.006 0.0079-1.03 0.039 0.0067-0.49 0.109 0.00620.04 0.168 0.00660.59 0.221 0.00681.02 0.267 0.00732.10 0.378 0.00803.08 0.477 0.00884.19 0.589 0.00995.19 0.689 0.01136.12 0.782 0.01297.26 0.884 0.01588.26 0.970 0.01909.28 1.044 0.0249

10.26 1.100 0.0335

Run: bm05529 editRe = 448918.9

α Cl Cd

-3.01 -0.150 0.0101-2.06 -0.045 0.0081-1.52 0.015 0.0074-1.02 0.046 0.0069-0.46 0.096 0.0058-0.03 0.147 0.00571.05 0.260 0.00632.08 0.361 0.00703.06 0.459 0.00784.11 0.579 0.00885.11 0.679 0.00996.21 0.775 0.0115

AG455ct-02r fp1.6Fig. 4.68

Run: bm05471 edit1Re = 39744.4

α Cl Cd

-4.02 -0.090 0.0198-3.02 -0.002 0.0146-2.01 0.077 0.0158-1.44 0.140 0.0130-0.98 0.206 0.0155-0.44 0.236 0.01590.09 0.288 0.02020.55 0.345 0.01871.07 0.363 0.02142.13 0.548 0.02143.18 0.702 0.02614.17 0.771 0.03015.23 0.825 0.03526.21 0.888 0.03697.24 0.951 0.05458.23 0.978 0.07659.25 0.956 0.0920

Run: bm05472 editRe = 60127.5

α Cl Cd

-4.04 -0.072 0.0175-3.01 0.010 0.0132-2.01 0.114 0.0153-1.46 0.173 0.0120-1.04 0.218 0.0126-0.42 0.279 0.01660.05 0.322 0.01430.57 0.367 0.02141.15 0.434 0.01862.18 0.625 0.01613.15 0.704 0.02074.19 0.769 0.02765.20 0.836 0.02926.19 0.911 0.03007.23 0.974 0.03688.28 1.029 0.05079.31 1.065 0.0836

Run: bm05474 editRe = 79874.3

α Cl Cd

-4.03 -0.048 0.0182-2.97 0.047 0.0129-2.10 0.134 0.0130-1.51 0.188 0.0127

-0.98 0.240 0.0121-0.50 0.282 0.01330.08 0.345 0.01540.55 0.419 0.01441.10 0.510 0.01282.19 0.613 0.01553.21 0.706 0.01554.14 0.785 0.01935.34 0.877 0.02236.20 0.943 0.02627.25 1.011 0.03288.26 1.072 0.04159.28 1.102 0.0604

Run: bb05476 edit1Re = 99825.9

α Cl Cd

-4.03 -0.046 0.0168-2.99 0.047 0.0130-1.98 0.141 0.0093-1.49 0.189 0.0101-0.98 0.244 0.0111-0.43 0.297 0.01270.11 0.353 0.01310.61 0.441 0.01561.13 0.522 0.01262.19 0.610 0.01243.18 0.696 0.01464.15 0.785 0.01745.24 0.882 0.02006.23 0.950 0.02407.27 1.014 0.03078.21 1.081 0.03929.25 1.117 0.0563

Run: bm05478 edit1Re = 150129.3

α Cl Cd

-4.00 -0.046 0.0148-2.99 0.053 0.0118-1.94 0.164 0.0099-1.43 0.226 0.0096-0.97 0.271 0.0097-0.44 0.340 0.01050.07 0.410 0.01060.58 0.476 0.01011.11 0.535 0.00992.13 0.629 0.01033.14 0.722 0.01224.16 0.811 0.01475.22 0.908 0.01776.22 0.991 0.02147.26 1.066 0.0269


8.30 1.129 0.03589.25 1.166 0.0473

AG455ct-02r fp3.6Fig. 4.71

Run: bm05480 edit1Re = 39715.2

α Cl Cd

-5.05 -0.058 0.0273-4.04 0.030 0.0199-3.01 0.108 0.0188-1.96 0.206 0.0230-1.49 0.257 0.0177-0.98 0.322 0.0189-0.43 0.357 0.02140.08 0.379 0.02260.52 0.415 0.02261.05 0.486 0.02322.15 0.675 0.02373.11 0.794 0.02744.16 0.893 0.02965.18 0.921 0.03816.25 1.011 0.04777.22 1.049 0.06178.24 1.069 0.0608

Run: bb05482 editRe = 59692.5

α Cl Cd

-5.09 -0.055 0.0223-4.01 0.046 0.0178-3.04 0.121 0.0150-1.99 0.237 0.0162-1.51 0.284 0.0150-0.99 0.329 0.0174-0.44 0.391 0.01820.10 0.435 0.02000.56 0.475 0.01931.15 0.586 0.02032.13 0.765 0.01833.20 0.833 0.01924.22 0.910 0.02515.21 0.974 0.02736.21 1.038 0.03347.26 1.102 0.04008.26 1.147 0.0559

Run: bb05484 editRe = 79707.0

α Cl Cd

-5.06 -0.058 0.0225-4.02 0.038 0.0171-2.99 0.125 0.0142-1.97 0.240 0.0133-1.47 0.283 0.0140-0.96 0.332 0.0152-0.42 0.388 0.01640.03 0.462 0.01720.60 0.550 0.01621.16 0.631 0.01592.21 0.722 0.01443.24 0.820 0.01804.25 0.888 0.02145.29 0.962 0.02516.26 1.032 0.03217.28 1.089 0.04068.26 1.132 0.0529

Run: bm05486 editRe = 99863.2

α Cl Cd

-5.07 -0.030 0.0189-4.02 0.072 0.0150-3.01 0.168 0.0129-2.01 0.259 0.0130-1.47 0.323 0.0137-0.96 0.379 0.0140-0.38 0.449 0.01560.15 0.523 0.01540.68 0.595 0.01561.11 0.656 0.01452.14 0.741 0.01353.17 0.827 0.01644.20 0.914 0.02005.22 0.986 0.02326.21 1.057 0.02837.21 1.110 0.03788.31 1.161 0.0522

Run: bm05488 editRe = 149668.0

α Cl Cd

-5.06 0.008 0.0162-3.99 0.136 0.0124-2.89 0.248 0.0113-1.93 0.345 0.0106-1.44 0.393 0.0106-0.93 0.450 0.0107-0.39 0.512 0.01170.08 0.567 0.0114

0.57 0.611 0.01121.10 0.664 0.01082.19 0.762 0.01103.17 0.843 0.01394.18 0.940 0.01685.22 1.033 0.02026.28 1.114 0.02527.24 1.170 0.03238.28 1.214 0.0417

AG455ct-02r fn15.4Fig. 4.74

Run: bb05502 editRe = 60016.5

α Cl Cd

-0.06 -0.611 0.02490.96 -0.552 0.02001.43 -0.518 0.01862.01 -0.401 0.02422.53 -0.338 0.02383.08 -0.293 0.02203.61 -0.257 0.02464.03 -0.210 0.02375.12 -0.115 0.02646.14 -0.032 0.02817.11 0.032 0.02538.20 0.133 0.02449.20 0.239 0.0253

10.20 0.338 0.033811.20 0.454 0.031712.23 0.553 0.0342

Run: bm05504 editRe = 99941.4

α Cl Cd

-0.03 -0.628 0.02490.94 -0.548 0.01911.50 -0.505 0.01592.07 -0.488 0.01982.50 -0.447 0.02173.03 -0.395 0.02133.49 -0.352 0.02154.06 -0.304 0.02245.13 -0.184 0.02186.03 -0.091 0.02167.11 0.017 0.01998.16 0.120 0.01989.21 0.222 0.0202

10.23 0.326 0.022911.22 0.422 0.025812.26 0.522 0.0312


AG455ct-02r fn10.4Fig. 4.77

Run: bm05506 editRe = 59833.4

α Cl Cd

-1.00 -0.406 0.0206-0.02 -0.325 0.01530.52 -0.271 0.01471.02 -0.186 0.01561.57 -0.104 0.01452.06 -0.066 0.01192.56 -0.015 0.01823.11 0.042 0.01584.13 0.110 0.01765.15 0.189 0.02676.13 0.245 0.02517.17 0.342 0.02528.20 0.436 0.03059.21 0.521 0.0291

10.24 0.599 0.030011.24 0.670 0.0394

Run: bb05508 editRe = 100143.9

α Cl Cd

-1.05 -0.427 0.0205-0.03 -0.340 0.01460.48 -0.312 0.01221.01 -0.272 0.01151.56 -0.227 0.01301.99 -0.181 0.01332.57 -0.127 0.01383.05 -0.077 0.01504.05 0.013 0.01585.08 0.127 0.01536.10 0.228 0.01537.10 0.329 0.01608.20 0.432 0.01879.23 0.532 0.0205

10.31 0.629 0.025911.23 0.706 0.0322

AG455ct-02r fn5.4Fig. 4.80

Run: bb05511 editRe = 59572.4

α Cl Cd

-2.03 -0.224 0.0193-1.01 -0.127 0.0119

-0.50 -0.067 0.01100.02 -0.027 0.01120.60 0.020 0.01011.06 0.056 0.01061.57 0.086 0.01192.08 0.158 0.01283.12 0.284 0.01424.10 0.365 0.02075.11 0.453 0.02246.18 0.541 0.03407.25 0.614 0.02968.17 0.687 0.02849.25 0.778 0.0390

10.28 0.834 0.0502

Run: bb05515 editRe = 99977.7

α Cl Cd

-2.02 -0.234 0.0158-1.01 -0.143 0.0104-0.54 -0.075 0.00910.03 -0.021 0.00900.54 0.028 0.00961.03 0.080 0.01001.60 0.152 0.01092.03 0.192 0.01163.12 0.281 0.01394.13 0.372 0.01475.11 0.452 0.01536.17 0.547 0.01697.25 0.638 0.01948.21 0.727 0.02329.30 0.816 0.0285

10.29 0.878 0.0396

AG455ct-02r fp4.6Fig. 4.83

Run: bm05517 editRe = 59747.2

α Cl Cd

-4.02 0.103 0.0171-3.01 0.191 0.0198-2.50 0.238 0.0152-1.99 0.295 0.0176-1.46 0.334 0.0191-0.91 0.378 0.0192-0.45 0.435 0.02040.03 0.473 0.02091.13 0.620 0.02202.16 0.793 0.01983.21 0.877 0.0224

4.16 0.944 0.02855.29 1.001 0.03026.18 1.043 0.03657.21 1.134 0.04738.29 1.177 0.0620

Run: bm05519 editRe = 100056.5

α Cl Cd

-4.02 0.145 0.0151-3.00 0.216 0.0146-2.50 0.267 0.0141-1.98 0.329 0.0137-1.40 0.391 0.0150-0.97 0.431 0.0151-0.41 0.496 0.01600.12 0.566 0.01671.12 0.717 0.01562.15 0.792 0.01493.19 0.882 0.01674.19 0.963 0.02025.18 1.035 0.02386.26 1.105 0.03037.22 1.153 0.03798.25 1.191 0.0511

AG455ct-02r fp9.6Fig. 4.86

Run: bm05521 editRe = 59852.6

α Cl Cd

-5.00 0.289 0.0245-3.97 0.372 0.0222-3.50 0.417 0.0225-2.99 0.464 0.0239-2.52 0.507 0.0242-1.96 0.556 0.0268-1.48 0.588 0.0273-0.98 0.642 0.0287-0.36 0.707 0.02990.11 0.755 0.03231.08 0.884 0.03122.16 1.072 0.02253.19 1.123 0.02874.22 1.167 0.03305.25 1.194 0.04246.27 1.259 0.0530


Run: bb05523 editRe = 100062.5

α Cl Cd

-5.00 0.409 0.0189-4.02 0.467 0.0220-3.46 0.530 0.0222-2.95 0.557 0.0216-2.42 0.590 0.0235-1.97 0.623 0.0231-1.37 0.658 0.0257-0.84 0.703 0.0265-0.39 0.753 0.02730.11 0.810 0.02831.16 0.990 0.01962.20 1.058 0.01773.22 1.129 0.02204.18 1.182 0.02845.22 1.229 0.03476.20 1.254 0.0435

AG455ct-02r fp14.6Fig. 4.89

Run: bb05525 editRe = 60037.2

α Cl Cd

-6.03 0.373 0.0381-5.07 0.462 0.0398-4.46 0.489 0.0424-4.01 0.519 0.0396-3.47 0.561 0.0399-2.96 0.610 0.0438-2.47 0.644 0.0445-1.98 0.682 0.0458-0.92 0.769 0.04610.04 0.814 0.04591.13 1.035 0.04712.23 1.235 0.03093.23 1.230 0.04104.17 1.239 0.04845.18 1.282 0.05676.26 1.314 0.0706

Run: bm05527 editRe = 99966.4

α Cl Cd

-6.04 0.390 0.0337-5.01 0.457 0.0355-3.98 0.533 0.0378-2.98 0.616 0.0392-1.90 0.712 0.0423-1.47 0.746 0.0431

-0.95 0.788 0.04360.03 0.854 0.04170.58 1.036 0.04250.61 1.044 0.04251.18 1.210 0.02202.16 1.273 0.02683.20 1.289 0.03464.22 1.301 0.04305.22 1.331 0.05146.19 1.348 0.0631

AG455ct-02r fn0.4Fig. 4.91

Run: bm05494 editRe = 59634.7

α Cl Cd

-3.02 -0.074 0.0163-2.03 0.002 0.0116-1.48 0.062 0.0122-0.94 0.109 0.0112-0.51 0.149 0.01360.00 0.191 0.01490.57 0.251 0.01280.99 0.287 0.01802.08 0.482 0.01513.11 0.570 0.01654.12 0.639 0.02065.19 0.728 0.0246

Run: bm05496 edit1Re = 100024.0

α Cl Cd

-3.02 -0.059 0.0145-2.04 0.014 0.0093-1.50 0.072 0.0095-0.97 0.125 0.0083-0.49 0.169 0.01030.04 0.223 0.01130.57 0.283 0.01131.13 0.372 0.01162.19 0.471 0.01193.25 0.574 0.01404.18 0.651 0.01615.21 0.749 0.0182

AG455ct-02r fn2.4Fig. 4.93

Run: bb05498 editRe = 99971.8

α Cl Cd

-3.10 -0.184 0.0173-2.03 -0.081 0.0118-1.51 -0.036 0.0094-0.98 0.007 0.0089-0.56 0.046 0.00930.02 0.106 0.00960.51 0.154 0.01091.12 0.245 0.01092.11 0.368 0.01153.19 0.469 0.01364.22 0.559 0.01455.24 0.654 0.0167

Run: bm05500 editRe = 299689.4

α Cl Cd

-3.08 -0.182 0.0121-1.97 -0.064 0.0090-1.49 -0.018 0.0080-0.99 0.029 0.0070-0.40 0.088 0.00600.03 0.141 0.00610.50 0.186 0.00661.03 0.243 0.00702.13 0.355 0.00793.06 0.454 0.00844.13 0.562 0.00965.17 0.664 0.0109

AG455ct-02r fp3.6Fig. 4.95

Run: bb05490 editRe = 39830.0

α Cl Cd

-5.06 -0.035 0.0317-4.00 0.043 0.0223-2.97 0.135 0.0149-1.91 0.247 0.0186-1.45 0.278 0.0145-0.93 0.342 0.0200-0.43 0.381 0.02170.11 0.409 0.02050.55 0.461 0.02291.11 0.510 0.0233


2.18 0.680 0.02633.20 0.796 0.02204.20 0.860 0.0311

Run: bm05492 editRe = 59794.0

α Cl Cd

-5.03 -0.072 0.0219-4.04 0.026 0.0183-2.95 0.127 0.0156-1.99 0.218 0.0176-1.48 0.261 0.0164-0.93 0.309 0.0184-0.44 0.352 0.01730.08 0.395 0.01730.60 0.450 0.02141.08 0.530 0.02032.18 0.721 0.01913.13 0.800 0.01754.19 0.880 0.0242

CAL1215jFig. 4.101

Run: bm05557 editRe = 99875.6

α Cl Cd

-6.12 -0.444 0.0211-5.14 -0.381 0.0197-4.13 -0.288 0.0171-3.11 -0.179 0.0168-2.03 -0.053 0.0161-1.08 0.067 0.0164-0.01 0.209 0.01721.06 0.333 0.01782.03 0.415 0.01903.08 0.503 0.01834.09 0.592 0.01765.11 0.676 0.01816.18 0.775 0.01887.17 0.874 0.02088.24 0.967 0.02389.21 1.039 0.0274

10.23 1.100 0.031411.24 1.146 0.039312.26 1.155 0.0562

Run: bm05562 editRe = 199938.5

α Cl Cd

-6.81 -0.492 0.0192-5.83 -0.396 0.0168

-4.80 -0.263 0.0149-3.74 -0.132 0.0125-2.70 -0.025 0.0111-1.74 0.062 0.0102-0.63 0.161 0.01000.42 0.275 0.01011.42 0.384 0.01032.41 0.483 0.01053.47 0.587 0.01104.50 0.687 0.01195.51 0.781 0.01316.54 0.877 0.01467.52 0.966 0.01678.52 1.047 0.01929.56 1.122 0.0227

10.52 1.174 0.028411.61 1.196 0.0373

Run: bb05564 editRe = 300452.3

α Cl Cd

-6.80 -0.489 0.0168-5.81 -0.366 0.0137-4.74 -0.237 0.0118-3.80 -0.138 0.0108-2.72 -0.034 0.0096-1.69 0.066 0.0089-0.71 0.161 0.00840.40 0.271 0.00841.37 0.387 0.00852.36 0.491 0.00873.43 0.595 0.00924.48 0.701 0.01015.49 0.803 0.01156.54 0.901 0.01317.55 0.991 0.01518.59 1.073 0.01769.55 1.146 0.0209

10.51 1.194 0.025811.63 1.229 0.0360

Run: bm05566 edit1Re = 400804.8

α Cl Cd

-7.76 -0.565 0.0196-6.82 -0.460 0.0151-5.77 -0.341 0.0123-4.80 -0.234 0.0108-3.70 -0.130 0.0097-2.69 -0.032 0.0092-1.69 0.069 0.0085-0.72 0.169 0.00790.35 0.274 0.0076

1.35 0.380 0.00842.36 0.492 0.00793.44 0.602 0.00874.42 0.701 0.00955.44 0.797 0.01086.53 0.898 0.01247.55 0.992 0.01448.57 1.075 0.01679.68 1.158 0.0206

10.55 1.205 0.025011.58 1.238 0.035312.57 1.238 0.0603

Run: bm05568 edit1Re = 500090.8

α Cl Cd

-7.81 -0.549 0.0179-6.88 -0.447 0.0139-5.78 -0.328 0.0115-4.83 -0.234 0.0101-3.70 -0.125 0.0091-2.76 -0.032 0.0086-1.67 0.080 0.0080-0.63 0.191 0.00740.34 0.284 0.00711.38 0.390 0.00722.40 0.502 0.00763.45 0.610 0.00844.46 0.710 0.00935.57 0.816 0.01056.54 0.911 0.01207.60 1.002 0.01408.57 1.076 0.01639.65 1.157 0.0201

10.65 1.212 0.025211.63 1.244 0.035312.51 1.252 0.0611

CAL2263mFig. 4.105

Run: bb05416 fullRe = 59801.9

α Cl Cd

-6.17 -0.452 0.0595-5.20 -0.391 0.0391-4.14 -0.289 0.0286-3.10 -0.168 0.0273-2.11 -0.053 0.0234-1.07 0.071 0.0225-0.05 0.220 0.02471.01 0.356 0.0267


2.00 0.459 0.02793.03 0.533 0.03114.06 0.665 0.03885.11 0.769 0.03846.13 0.767 0.04836.16 0.772 0.04757.10 0.897 0.05598.16 1.017 0.03869.17 1.114 0.0373

10.21 1.183 0.035611.20 1.227 0.038312.21 1.264 0.0414

Run: bm05439 editRe = 99887.9

α Cl Cd

-6.15 -0.440 0.0697-5.16 -0.316 0.0357-4.18 -0.197 0.0253-3.11 -0.061 0.0235-2.06 0.070 0.0208-1.01 0.209 0.0185-0.01 0.332 0.01901.01 0.433 0.01972.10 0.533 0.01963.08 0.626 0.01934.07 0.712 0.01835.12 0.802 0.01906.18 0.891 0.01987.18 0.967 0.02128.17 1.049 0.02309.22 1.127 0.0257

10.23 1.187 0.029011.19 1.228 0.031912.28 1.227 0.0401

Run: bb05422 editRe = 200162.8

α Cl Cd

-6.13 -0.266 0.0291-5.13 -0.155 0.0191-4.15 -0.052 0.0164-3.10 0.051 0.0147-2.02 0.145 0.0122-1.01 0.229 0.01040.02 0.368 0.01001.02 0.467 0.01052.07 0.567 0.01093.09 0.664 0.01134.10 0.768 0.01195.13 0.859 0.01286.14 0.948 0.01427.17 1.032 0.0160

8.24 1.107 0.01819.17 1.172 0.0205

10.23 1.224 0.023911.24 1.239 0.030212.18 1.237 0.0402

Run: bb05424 editRe = 299574.3

α Cl Cd

-6.11 -0.209 0.0168-5.14 -0.123 0.0141-4.02 -0.023 0.0129-3.02 0.076 0.0119-2.05 0.168 0.0108-0.96 0.273 0.00890.07 0.402 0.00821.03 0.501 0.00862.05 0.607 0.00893.15 0.710 0.00944.15 0.803 0.01025.11 0.885 0.01136.16 0.969 0.01297.21 1.053 0.01478.26 1.139 0.01689.23 1.192 0.0192

10.30 1.233 0.023711.16 1.250 0.028512.22 1.258 0.0408

Run: bb05428 editRe = 399877.8

α Cl Cd

-6.17 -0.241 0.0150-5.16 -0.147 0.0126-4.08 -0.049 0.0114-3.10 0.051 0.0109-2.03 0.158 0.0096-0.95 0.270 0.00840.05 0.362 0.00730.96 0.482 0.00752.03 0.588 0.00793.13 0.689 0.00854.14 0.773 0.00945.05 0.861 0.01056.17 0.953 0.01207.26 1.041 0.01388.20 1.117 0.01569.26 1.174 0.0185

10.31 1.206 0.023111.30 1.240 0.028412.23 1.253 0.0378

Run: bb05441 editRe = 499823.0

α Cl Cd

-5.02 -0.126 0.0116-4.09 -0.037 0.0109-2.99 0.077 0.0102-2.04 0.175 0.0093-0.95 0.278 0.00810.02 0.369 0.00701.09 0.500 0.00722.12 0.605 0.00763.16 0.702 0.00834.07 0.779 0.00915.13 0.868 0.01036.21 0.969 0.01187.20 1.046 0.01338.22 1.120 0.01539.25 1.164 0.0186

10.30 1.220 0.023411.27 1.241 0.029312.27 1.258 0.0425

CAL4014lFig. 4.109

Run: bb05545 editRe = 99950.5

α Cl Cd

-6.18 -0.531 0.0293-5.19 -0.416 0.0212-4.11 -0.296 0.0185-3.09 -0.192 0.0164-2.11 -0.107 0.0148-1.01 0.015 0.0142-0.01 0.111 0.01651.01 0.205 0.01872.06 0.291 0.01853.01 0.374 0.01834.11 0.470 0.01735.10 0.556 0.01746.15 0.653 0.01927.14 0.751 0.02048.19 0.840 0.02319.22 0.924 0.0279

10.25 1.000 0.032711.23 1.058 0.0413

Run: bm05547 fullRe = 200284.9

α Cl Cd

-6.14 -0.520 0.0201-5.20 -0.437 0.0172


-4.15 -0.335 0.0159-3.21 -0.243 0.0144-2.09 -0.138 0.0127-1.00 -0.041 0.0108-0.03 0.066 0.01040.98 0.178 0.01062.01 0.281 0.01103.07 0.396 0.01124.11 0.505 0.01155.11 0.609 0.01246.14 0.720 0.01367.11 0.811 0.01508.15 0.914 0.01759.27 1.008 0.0213

10.22 1.086 0.027111.15 1.137 0.032912.73 1.185 0.0577

Run: bm05550 editRe = 300082.1

α Cl Cd

-6.13 -0.559 0.0180-5.12 -0.468 0.0146-4.13 -0.364 0.0135-3.10 -0.255 0.0125-2.07 -0.149 0.0112-1.04 -0.045 0.0102-0.03 0.050 0.00821.01 0.180 0.00842.02 0.285 0.00863.04 0.389 0.00894.04 0.493 0.00965.08 0.610 0.01066.12 0.721 0.01157.20 0.828 0.01318.16 0.927 0.01569.21 1.023 0.0189

10.22 1.105 0.023311.28 1.179 0.029612.23 1.204 0.041713.24 1.235 0.0593

Run: bb05555 editRe = 399909.5

α Cl Cd

-6.20 -0.571 0.0164-5.09 -0.457 0.0135-4.15 -0.361 0.0123-3.13 -0.250 0.0113-2.11 -0.145 0.0104-1.03 -0.030 0.0096-0.03 0.064 0.00831.00 0.191 0.0073

1.97 0.295 0.00753.03 0.410 0.00804.14 0.528 0.00885.12 0.640 0.00976.10 0.740 0.01077.15 0.845 0.01238.17 0.948 0.01469.23 1.045 0.0176

10.20 1.129 0.021211.27 1.197 0.027712.28 1.243 0.0389

Run: bb05553 edit1Re = 499702.6

α Cl Cd

-6.23 -0.587 0.0146-5.17 -0.483 0.0125-4.13 -0.376 0.0113-3.12 -0.266 0.0104-2.06 -0.149 0.0096-1.09 -0.055 0.0090-0.03 0.053 0.00790.96 0.163 0.00662.01 0.288 0.00703.06 0.396 0.00754.04 0.505 0.00845.04 0.618 0.00916.12 0.731 0.01037.19 0.835 0.01218.26 0.947 0.01459.16 1.038 0.0170

10.19 1.124 0.020811.27 1.202 0.028812.20 1.244 0.044113.26 1.266 0.0684

E387 (E)Fig. 4.113

Run: bb05385 editRe = 59888.7

α Cl Cd

-5.11 -0.238 0.0489-4.15 -0.155 0.0329-3.07 -0.036 0.0252-2.04 0.084 0.0273-1.00 0.207 0.0214-0.01 0.348 0.02691.06 0.442 0.03082.03 0.530 0.03463.06 0.571 0.04104.10 0.619 0.0485

5.11 0.669 0.05806.11 0.752 0.06107.09 0.829 0.05728.19 1.157 0.03919.24 1.212 0.0267

10.20 1.245 0.035011.26 1.261 0.0486

Run: bm05374 editRe = 99932.8

α Cl Cd

-6.27 -0.359 0.0756-5.24 -0.299 0.0577-4.24 -0.139 0.0316-3.16 0.014 0.0219-2.13 0.171 0.0186-1.06 0.284 0.0170-0.06 0.377 0.02050.96 0.480 0.02301.99 0.565 0.02572.95 0.645 0.02883.97 0.730 0.02745.03 0.832 0.02376.08 0.920 0.02077.04 1.016 0.02088.10 1.106 0.02079.17 1.166 0.0250

10.13 1.187 0.033411.08 1.194 0.044912.11 1.213 0.0564

Run: bm05376 editRe = 199747.1

α Cl Cd

-6.28 -0.385 0.0771-5.23 -0.184 0.0396-4.21 -0.048 0.0196-3.12 0.058 0.0150-2.20 0.151 0.0129-1.10 0.249 0.0106-0.09 0.353 0.01051.01 0.465 0.01121.95 0.560 0.01183.00 0.666 0.01284.03 0.771 0.01355.02 0.880 0.01406.04 0.974 0.01407.08 1.080 0.01488.12 1.156 0.01699.07 1.197 0.0244

10.08 1.221 0.034311.22 1.222 0.051712.14 1.229 0.0688


Run: bb05381 editRe = 299688.0

α Cl Cd

-6.30 -0.338 0.0697-5.27 -0.183 0.0387-4.17 -0.050 0.0175-3.12 0.052 0.0131-2.14 0.155 0.0113-1.12 0.255 0.0096-0.11 0.355 0.00831.01 0.478 0.00892.05 0.580 0.00922.96 0.676 0.00974.06 0.786 0.01055.04 0.893 0.01116.06 0.991 0.01187.05 1.080 0.01288.03 1.156 0.01669.13 1.209 0.0239

10.15 1.237 0.033411.11 1.243 0.051712.15 1.246 0.0749

Run: bm05384 editRe = 459413.5

α Cl Cd

-6.31 -0.326 0.0702-5.26 -0.182 0.0387-4.28 -0.066 0.0156-3.11 0.051 0.0115-2.14 0.153 0.0103-1.14 0.252 0.0089-0.06 0.360 0.00730.93 0.477 0.00721.99 0.584 0.00752.97 0.691 0.00793.99 0.792 0.00865.04 0.908 0.00926.08 1.009 0.00997.15 1.097 0.01228.12 1.165 0.01629.13 1.215 0.0219

10.14 1.248 0.031111.11 1.263 0.0518

MA409Fig. 4.137

Run: 06269rd interpRe = 40282.6

α Cl Cd

-5.84 -0.383 0.0625-4.81 -0.300 0.0440-3.62 -0.199 0.0296-2.62 -0.111 0.0228-1.73 -0.042 0.0165-0.54 0.061 0.01670.36 0.141 0.01811.54 0.251 0.01972.44 0.330 0.02223.52 0.484 0.02414.58 0.623 0.02415.62 0.701 0.02656.60 0.768 0.02927.68 0.827 0.04068.53 0.857 0.05019.49 0.898 0.0640


α Cl Cd

-5.86 -0.398 0.0623-4.76 -0.292 0.0311-3.65 -0.203 0.0256-2.63 -0.129 0.0183-1.60 -0.052 0.0146-0.70 0.028 0.01560.43 0.122 0.01671.47 0.219 0.01852.48 0.384 0.02133.41 0.491 0.01814.57 0.587 0.02065.54 0.667 0.02166.46 0.734 0.02627.60 0.795 0.03068.54 0.852 0.04369.61 0.898 0.0614


α Cl Cd

-5.63 -0.386 0.0541-4.45 -0.256 0.0252-3.72 -0.200 0.0208-2.54 -0.109 0.0168-1.64 -0.035 0.0144-0.53 0.062 0.0129

0.49 0.168 0.01251.45 0.316 0.01312.55 0.451 0.01293.56 0.547 0.01344.62 0.651 0.01385.64 0.732 0.01526.53 0.803 0.01897.55 0.878 0.02398.59 0.941 0.03249.57 0.990 0.0451


α Cl Cd

-5.66 -0.391 0.0527-4.65 -0.282 0.0173-3.63 -0.195 0.0147-2.57 -0.100 0.0127-1.73 -0.016 0.0110-0.51 0.150 0.00990.43 0.249 0.00881.52 0.355 0.00872.60 0.466 0.00863.49 0.559 0.00964.57 0.659 0.01125.41 0.728 0.01336.47 0.818 0.01677.57 0.893 0.02238.69 0.960 0.02989.74 1.009 0.0454

10.72 1.036 0.1234


α Cl Cd

-6.86 -0.527 0.0911-5.66 -0.390 0.0359-4.58 -0.286 0.0136-3.75 -0.207 0.0125-2.76 -0.111 0.0112-1.70 0.025 0.0097-0.70 0.149 0.00810.32 0.273 0.00681.40 0.380 0.00712.58 0.499 0.00793.57 0.594 0.00894.58 0.688 0.01095.52 0.772 0.01296.66 0.868 0.01597.63 0.940 0.02068.45 0.990 0.02539.73 1.059 0.0396

10.60 1.088 0.1208


NACA 43012AFig. 4.141


α Cl Cd

-8.59 -0.478 0.0966-7.59 -0.483 0.0624-6.56 -0.442 0.0467-5.55 -0.377 0.0345-4.54 -0.316 0.0282-3.47 -0.254 0.0245-2.49 -0.150 0.0214-1.38 0.026 0.0176-0.47 0.147 0.01980.66 0.237 0.02301.62 0.305 0.02242.68 0.394 0.02883.69 0.470 0.03354.72 0.537 0.03755.65 0.600 0.04396.67 0.645 0.05837.68 0.633 0.11028.69 0.597 0.13389.66 0.585 0.1401

10.71 0.594 0.1495


α Cl Cd

-8.62 -0.526 0.1146-7.55 -0.510 0.0823-6.50 -0.449 0.0495-5.49 -0.371 0.0356-4.51 -0.305 0.0275-3.47 -0.200 0.0211-2.48 -0.081 0.0178-1.46 0.015 0.0134-0.44 0.127 0.01430.65 0.235 0.01501.61 0.307 0.01592.68 0.394 0.01783.66 0.479 0.02024.70 0.572 0.02305.58 0.655 0.02486.79 0.764 0.02967.80 0.844 0.03468.73 0.790 0.11479.68 0.712 0.1376

10.68 0.694 0.1442


α Cl Cd

-8.62 -0.576 0.1190-7.65 -0.564 0.0855-6.55 -0.493 0.0519-5.54 -0.404 0.0338-4.48 -0.285 0.0225-3.48 -0.171 0.0172-2.43 -0.074 0.0146-1.37 0.016 0.0119-0.41 0.092 0.01040.58 0.202 0.01121.66 0.332 0.01202.67 0.427 0.01293.69 0.521 0.01414.69 0.618 0.01545.69 0.717 0.01706.73 0.819 0.01907.76 0.913 0.02138.78 1.007 0.02439.83 1.098 0.0269

10.85 1.182 0.029111.88 1.262 0.031712.86 1.330 0.033313.87 1.395 0.035414.87 1.393 0.1241


α Cl Cd

-8.18 -0.602 0.1205-7.19 -0.544 0.0857-6.12 -0.453 0.0535-5.04 -0.330 0.0317-3.99 -0.216 0.0199-3.01 -0.127 0.0155-2.00 -0.036 0.0134-0.98 0.056 0.01200.07 0.137 0.00931.07 0.252 0.00992.12 0.397 0.01063.14 0.492 0.01124.16 0.592 0.01195.20 0.695 0.01326.24 0.797 0.01457.25 0.901 0.01608.26 0.999 0.01779.36 1.101 0.0198

10.35 1.187 0.021611.37 1.278 0.023412.33 1.348 0.024913.33 1.429 0.0267


α Cl Cd

-6.13 -0.428 0.0473-5.09 -0.318 0.0296-4.04 -0.220 0.0187-3.00 -0.128 0.0142-2.04 -0.038 0.0126-1.00 0.060 0.01140.02 0.151 0.00861.10 0.261 0.00852.10 0.389 0.00973.17 0.516 0.01014.17 0.614 0.01105.18 0.716 0.01206.21 0.819 0.01337.28 0.920 0.01478.30 1.027 0.01619.24 1.112 0.0176

10.40 1.223 0.019411.34 1.311 0.020612.43 1.400 0.022313.45 1.460 0.0276


α Cl Cd

-6.11 -0.413 0.0467-5.05 -0.312 0.0288-4.02 -0.216 0.0178-2.94 -0.115 0.0135-1.99 -0.027 0.0121-0.97 0.066 0.01130.02 0.167 0.00861.08 0.267 0.00782.16 0.391 0.00913.15 0.512 0.00964.18 0.622 0.01045.25 0.727 0.01136.33 0.839 0.01247.20 0.921 0.01348.23 1.026 0.01479.25 1.129 0.0161

10.42 1.238 0.017711.39 1.333 0.019012.42 1.415 0.021013.39 1.473 0.0299


S8064Fig. 4.155


α Cl Cd

-6.15 -0.599 0.0264-5.11 -0.523 0.0205-4.07 -0.431 0.0170-3.10 -0.325 0.0187-2.06 -0.272 0.0182-1.03 -0.161 0.0194-0.05 -0.024 0.02061.04 0.137 0.02052.04 0.228 0.02333.09 0.304 0.02614.09 0.462 0.02295.12 0.567 0.01896.18 0.655 0.01617.20 0.733 0.01768.20 0.770 0.02489.18 0.818 0.0317

10.21 0.868 0.039211.24 0.901 0.0527


α Cl Cd

-7.14 -0.558 0.0238-6.10 -0.470 0.0187-5.12 -0.384 0.0151-4.05 -0.297 0.0123-3.10 -0.206 0.0130-2.04 -0.104 0.0133-0.98 -0.014 0.0142-0.04 0.067 0.01450.98 0.150 0.01512.06 0.247 0.01433.13 0.374 0.01284.12 0.514 0.01165.15 0.619 0.01006.19 0.704 0.01267.13 0.754 0.01618.17 0.819 0.02149.23 0.873 0.0264

10.25 0.915 0.033311.19 0.947 0.0434


α Cl Cd

-7.16 -0.536 0.0195-6.12 -0.455 0.0163-5.02 -0.362 0.0140-4.07 -0.287 0.0114-3.01 -0.188 0.0108-1.98 -0.093 0.0109-1.03 -0.004 0.01100.06 0.095 0.01091.05 0.193 0.01042.05 0.293 0.00993.08 0.397 0.00874.09 0.506 0.00875.15 0.636 0.00976.21 0.724 0.01237.15 0.786 0.01528.22 0.861 0.01999.25 0.908 0.0249

10.24 0.955 0.031511.24 0.982 0.0415


α Cl Cd

-7.61 -0.600 0.0200-6.63 -0.524 0.0170-5.53 -0.426 0.0144-4.56 -0.342 0.0124-3.47 -0.243 0.0098-2.54 -0.151 0.0093-1.51 -0.052 0.0090-0.49 0.046 0.00890.53 0.151 0.00851.57 0.261 0.00802.58 0.367 0.00763.63 0.472 0.00774.67 0.576 0.00875.66 0.673 0.01106.71 0.762 0.01367.72 0.839 0.01698.77 0.906 0.02129.71 0.952 0.0262

10.78 0.990 0.033711.75 1.024 0.0458


α Cl Cd

-8.09 -0.657 0.0206-7.07 -0.575 0.0168-6.08 -0.493 0.0146

-5.12 -0.401 0.0126-4.01 -0.296 0.0106-3.01 -0.203 0.0087-1.97 -0.101 0.0081-1.07 -0.013 0.00770.03 0.108 0.00741.09 0.222 0.00722.03 0.326 0.00693.09 0.433 0.00694.16 0.542 0.00775.18 0.632 0.00876.18 0.723 0.01197.19 0.805 0.01478.26 0.884 0.01839.27 0.952 0.0223

10.28 1.006 0.027111.28 1.032 0.034612.25 1.054 0.0503

S9000 fp0Fig. 4.159


α Cl Cd

-7.13 -0.389 0.0355-6.13 -0.326 0.0258-5.04 -0.222 0.0185-4.07 -0.149 0.0148-3.08 -0.073 0.0141-2.03 0.041 0.0159-0.98 0.148 0.01740.04 0.259 0.01961.13 0.434 0.02682.12 0.583 0.02903.14 0.676 0.02984.21 0.765 0.03045.21 0.838 0.03146.19 0.915 0.03437.29 0.993 0.03448.25 1.044 0.03799.30 1.097 0.0484

10.28 1.127 0.0687


α Cl Cd

-8.20 -0.496 0.0662-7.11 -0.383 0.0343-6.10 -0.307 0.0232-5.04 -0.219 0.0175-4.06 -0.133 0.0137


-3.04 -0.063 0.0120-2.02 0.033 0.0138-0.98 0.170 0.01470.04 0.353 0.01541.13 0.483 0.01462.15 0.571 0.01543.17 0.662 0.01524.13 0.740 0.01645.19 0.838 0.01856.20 0.913 0.02197.23 0.981 0.02618.22 1.045 0.03329.26 1.095 0.0404

10.27 1.121 0.052511.17 0.958 0.1872


α Cl Cd

-8.19 -0.489 0.0683-7.17 -0.389 0.0329-6.11 -0.308 0.0186-5.04 -0.222 0.0144-4.02 -0.124 0.0121-3.04 -0.000 0.0101-2.02 0.140 0.0086-0.94 0.267 0.00860.04 0.370 0.00891.05 0.469 0.00932.13 0.577 0.00993.12 0.671 0.01104.17 0.771 0.01245.21 0.864 0.01456.22 0.950 0.01737.21 1.027 0.02088.20 1.090 0.02439.28 1.152 0.0306

10.23 1.190 0.038811.18 1.191 0.0601


α Cl Cd

-7.17 -0.404 0.0332-6.06 -0.310 0.0166-5.07 -0.211 0.0128-4.06 -0.078 0.0109-2.99 0.057 0.0084-1.97 0.157 0.0070-0.94 0.264 0.0071-0.02 0.359 0.00741.10 0.475 0.00782.10 0.575 0.0085

3.16 0.682 0.00954.16 0.779 0.01115.17 0.880 0.01306.14 0.963 0.01507.22 1.060 0.01818.26 1.126 0.02139.32 1.192 0.0273

10.28 1.228 0.034711.27 1.238 0.0649


α Cl Cd

-7.11 -0.380 0.0301-6.09 -0.289 0.0146-5.02 -0.150 0.0107-4.08 -0.046 0.0095-3.00 0.062 0.0076-2.00 0.157 0.0065-0.99 0.249 0.00620.04 0.360 0.00661.10 0.467 0.00722.10 0.567 0.00803.14 0.668 0.00904.18 0.761 0.01045.22 0.861 0.01216.26 0.947 0.01437.24 1.031 0.01698.27 1.112 0.02019.28 1.167 0.0248

10.31 1.220 0.031911.02 1.223 0.0462


α Cl Cd

-7.10 -0.386 0.0236-6.08 -0.257 0.0126-5.02 -0.135 0.0100-3.99 -0.022 0.0079-2.97 0.075 0.0070-1.97 0.172 0.0061-0.95 0.271 0.00590.10 0.387 0.00621.15 0.494 0.00702.13 0.593 0.00773.19 0.695 0.00874.25 0.806 0.01015.26 0.900 0.01176.27 0.985 0.01387.27 1.068 0.01628.29 1.151 0.01939.32 1.228 0.0236

10.39 1.292 0.029511.32 1.298 0.0588

S9000 fp2.5Fig. 4.162


α Cl Cd

-8.10 -0.419 0.0510-7.09 -0.328 0.0325-6.07 -0.239 0.0250-5.10 -0.156 0.0200-4.04 -0.056 0.0155-2.99 0.026 0.0150-1.96 0.142 0.0163-0.94 0.242 0.01590.07 0.358 0.02201.10 0.521 0.02872.12 0.670 0.02883.20 0.778 0.02624.17 0.850 0.02885.24 0.950 0.03126.25 1.022 0.03357.26 1.091 0.03488.29 1.156 0.04349.29 1.209 0.0475

10.30 1.213 0.072311.22 1.038 0.1767


α Cl Cd

-8.12 -0.461 0.0661-7.09 -0.344 0.0335-6.06 -0.252 0.0233-5.05 -0.160 0.0182-4.00 -0.058 0.0142-3.02 0.013 0.0129-1.97 0.125 0.0132-0.95 0.263 0.01560.11 0.449 0.01571.09 0.566 0.01462.13 0.656 0.01503.16 0.737 0.01584.22 0.825 0.01805.23 0.919 0.01986.21 0.987 0.02437.24 1.046 0.02888.25 1.103 0.03569.25 1.147 0.0461

10.30 1.151 0.0559



α Cl Cd

-8.17 -0.446 0.0612-7.12 -0.344 0.0252-6.02 -0.244 0.0174-5.06 -0.144 0.0138-4.06 -0.023 0.0115-3.00 0.103 0.0110-1.97 0.240 0.0097-0.92 0.376 0.00960.09 0.480 0.00961.10 0.579 0.00962.16 0.678 0.01083.16 0.773 0.01234.19 0.865 0.01415.22 0.956 0.01636.25 1.040 0.01967.27 1.111 0.02338.25 1.166 0.02779.27 1.221 0.0337

10.34 1.234 0.045811.28 1.222 0.0751


α Cl Cd

-7.11 -0.293 0.0217-6.07 -0.173 0.0143-5.02 -0.058 0.0115-3.99 0.059 0.0099-3.03 0.169 0.0084-1.94 0.277 0.0079-0.98 0.379 0.00730.06 0.481 0.00751.13 0.590 0.00842.12 0.689 0.00953.20 0.796 0.01124.21 0.895 0.01265.23 0.983 0.01486.24 1.062 0.01747.26 1.141 0.02048.23 1.204 0.02449.25 1.254 0.0308

10.29 1.279 0.0426


α Cl Cd

-7.07 -0.247 0.0159-6.04 -0.136 0.0120-5.02 -0.027 0.0100-4.05 0.074 0.0086

-2.98 0.187 0.0071-1.95 0.285 0.0069-0.94 0.384 0.00680.08 0.496 0.00711.12 0.599 0.00802.18 0.706 0.00913.19 0.802 0.01054.21 0.896 0.01215.26 0.992 0.01426.26 1.074 0.01657.32 1.157 0.01938.28 1.221 0.02329.33 1.278 0.0286

10.33 1.306 0.0385


α Cl Cd

-7.07 -0.235 0.0141-5.99 -0.119 0.0113-5.01 -0.014 0.0090-4.01 0.087 0.0075-2.93 0.191 0.0067-1.97 0.287 0.0065-0.90 0.391 0.00660.13 0.503 0.00691.14 0.606 0.00772.18 0.705 0.00883.16 0.803 0.01004.20 0.896 0.01165.25 0.993 0.01346.31 1.083 0.01567.28 1.156 0.01838.31 1.231 0.02179.34 1.291 0.0260

10.32 1.334 0.0335

S9000 fp5Fig. 4.165


α Cl Cd

-8.15 -0.332 0.0468-7.11 -0.242 0.0334-6.08 -0.144 0.0242-5.07 -0.055 0.0193-4.06 0.027 0.0162-3.00 0.119 0.0153-1.97 0.216 0.0164-0.90 0.312 0.01780.07 0.404 0.0241

1.17 0.608 0.02932.20 0.741 0.02543.24 0.836 0.02624.25 0.913 0.02815.20 0.985 0.02996.28 1.048 0.03167.28 1.087 0.03568.23 1.141 0.04039.28 1.179 0.0570

10.27 0.988 0.0749


α Cl Cd

-8.10 -0.410 0.0611-7.06 -0.295 0.0303-6.06 -0.187 0.0213-5.04 -0.074 0.0173-4.02 0.026 0.0141-3.03 0.111 0.0125-1.98 0.218 0.0139-0.94 0.358 0.01700.10 0.546 0.01691.17 0.680 0.01552.13 0.770 0.01413.21 0.853 0.01654.26 0.958 0.01855.23 1.034 0.02166.22 1.091 0.02517.29 1.157 0.03058.27 1.202 0.03799.30 1.241 0.0477

10.28 1.258 0.0600


α Cl Cd

-8.14 -0.365 0.0536-7.09 -0.236 0.0218-6.06 -0.132 0.0156-5.02 -0.018 0.0127-4.03 0.091 0.0117-3.02 0.191 0.0112-1.91 0.321 0.0105-0.92 0.441 0.01000.14 0.553 0.00941.14 0.655 0.00972.18 0.751 0.01133.20 0.850 0.01334.25 0.945 0.01485.24 1.027 0.01766.28 1.097 0.02137.26 1.162 0.0244


8.30 1.217 0.02989.29 1.253 0.0376

10.27 1.269 0.0504


α Cl Cd

-8.03 -0.256 0.0289-7.06 -0.160 0.0165-6.02 -0.053 0.0127-5.06 0.039 0.0111-4.02 0.150 0.0098-2.97 0.262 0.0089-1.92 0.376 0.0079-0.91 0.476 0.00700.10 0.581 0.00771.11 0.680 0.00882.16 0.779 0.01023.15 0.870 0.01194.22 0.967 0.01375.26 1.055 0.01616.23 1.126 0.01867.26 1.203 0.02228.27 1.257 0.02709.32 1.299 0.0344

10.32 1.312 0.0492


α Cl Cd

-7.03 -0.128 0.0137-6.06 -0.030 0.0111-4.99 0.080 0.0094-3.96 0.188 0.0082-2.91 0.291 0.0073-1.92 0.393 0.0069-0.98 0.483 0.00670.14 0.593 0.00731.22 0.702 0.00852.08 0.786 0.00953.17 0.890 0.01124.30 0.984 0.01305.22 1.068 0.01516.24 1.147 0.01767.29 1.217 0.02078.33 1.285 0.02529.36 1.330 0.0308

10.32 1.337 0.044111.21 1.309 0.0913


α Cl Cd

-7.09 -0.127 0.0127-6.12 -0.036 0.0105-5.00 0.071 0.0089-3.99 0.180 0.0075-2.98 0.285 0.0066-1.99 0.388 0.0065-0.94 0.484 0.00670.10 0.591 0.00711.10 0.686 0.00802.19 0.788 0.00933.26 0.890 0.01074.20 0.977 0.01225.32 1.073 0.01446.32 1.151 0.01667.33 1.219 0.01968.31 1.285 0.02339.36 1.347 0.0278

10.38 1.366 0.037011.20 1.321 0.0828

Low Speed Airfoil Data V5 SELIG

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