low speed airfoil data v5 selig
TRANSCRIPT
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
Summary of Low-Speed Airfoil Data
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
Summary of Low-Speed Airfoil Data
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
4.5 Comparison between the true and actual AG16. . . . . . . . . . . . . . . . . . . . . . . . . . 36
4.6 Inviscid velocity distributions for the AG16. . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
4.7 Drag polar for the AG16. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37
4.8 Lift and moment characteristics for the AG16. . . . . . . . . . . . . . . . . . . . . . . . . . . 38
4.9 Comparison between the true and actual AG24. . . . . . . . . . . . . . . . . . . . . . . . . . 42
4.10 Inviscid velocity distributions for the AG24. . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
4.11 Drag polar for the AG24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
vi Summary of Low-Speed Airfoil Data
4.12 Lift and moment characteristics for the AG24. . . . . . . . . . . . . . . . . . . . . . . . . . . 44
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
4.27 Inviscid velocity distributions for the AG40d-02r with a 4 deg flap. . . . . . . . . . . . . . . . 72
4.28 Drag polar for the AG40d-02r with a 4 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . . 73
4.29 Lift and moment characteristics for the AG40d-02r with a 4 deg flap. . . . . . . . . . . . . . . 74
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
4.39 Inviscid velocity distributions for the AG40d-02r with a 5 deg flap. . . . . . . . . . . . . . . . 90
4.40 Drag polar for the AG40d-02r with a 5 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . . 91
4.41 Lift and moment characteristics for the AG40d-02r with a 5 deg flap. . . . . . . . . . . . . . . 92
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
4.45 Inviscid velocity distributions for the AG40d-02r with a 15 deg flap. . . . . . . . . . . . . . . 98
4.46 Drag polar for the AG40d-02r with a 15 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . 99
4.47 Lift and moment characteristics for the AG40d-02r with a 15 deg flap. . . . . . . . . . . . . . 100
4.48 Inviscid velocity distributions for the AG40d-02r with a 20 deg flap. . . . . . . . . . . . . . . 102
4.49 Drag polar for the AG40d-02r with a 20 deg flap. . . . . . . . . . . . . . . . . . . . . . . . . 103
4.50 Lift and moment characteristics for the AG40d-02r with a 20 deg flap. . . . . . . . . . . . . . 104
4.51 Inviscid velocity distributions for the AG40d-02r with a 30 deg flap. . . . . . . . . . . . . . . 106
4.52 Lift and moment characteristics for the AG40d-02r with a 30 deg flap. . . . . . . . . . . . . . 107
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
4.64 Inviscid velocity distributions for the AG455ct-02r with a −2.4 deg flap. . . . . . . . . . . . . 128
4.65 Drag polar for the AG455ct-02r with a −2.4 deg flap. . . . . . . . . . . . . . . . . . . . . . . 129
4.66 Lift and moment characteristics for the AG455ct-02r with a −2.4 deg flap. . . . . . . . . . . . 130
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
4.70 Inviscid velocity distributions for the AG455ct-02r with a 3.6 deg flap. . . . . . . . . . . . . . 140
4.71 Drag polar for the AG455ct-02r with a 3.6 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 141
4.72 Lift and moment characteristics for the AG455ct-02r with a 3.6 deg flap. . . . . . . . . . . . . 142
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
4.79 Inviscid velocity distributions for the AG455ct-02r with a −5.4 deg flap. . . . . . . . . . . . . 154
4.80 Drag polar for the AG455ct-02r with a −5.4 deg flap. . . . . . . . . . . . . . . . . . . . . . . 155
4.81 Lift and moment characteristics for the AG455ct-02r with a −5.4 deg flap. . . . . . . . . . . . 156
4.82 Inviscid velocity distributions for the AG455ct-02r with a 4.6 deg flap. . . . . . . . . . . . . . 158
4.83 Drag polar for the AG455ct-02r with a 4.6 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 159
4.84 Lift and moment characteristics for the AG455ct-02r with a 4.6 deg flap. . . . . . . . . . . . . 160
4.85 Inviscid velocity distributions for the AG455ct-02r with a 9.6 deg flap. . . . . . . . . . . . . . 162
4.86 Drag polar for the AG455ct-02r with a 9.6 deg flap. . . . . . . . . . . . . . . . . . . . . . . . 163
4.87 Lift and moment characteristics for the AG455ct-02r with a 9.6 deg flap. . . . . . . . . . . . . 164
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
4.147 Lift and moment characteristics for the S1223 with Gurney flap of h/c = 3.12%. . . . . . . . 259
4.148 Lift and moment characteristics for the S1223 with Gurney flap of h/c = 2.08%. . . . . . . . 263
4.149 Lift and moment characteristics for the S1223 with Gurney flap of h/c = 1.56%. . . . . . . . 265
4.150 Lift and moment characteristics for the S1223 with Gurney flap of h/c = 1.04%. . . . . . . . 267
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
4.153 Comparison between the true and actual S8064. . . . . . . . . . . . . . . . . . . . . . . . . . 272
4.154 Inviscid velocity distributions for the S8064. . . . . . . . . . . . . . . . . . . . . . . . . . . . 272
4.155 Drag polar for the S8064. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273
4.156 Lift and moment characteristics for the S8064. . . . . . . . . . . . . . . . . . . . . . . . . . . 274
4.157 Comparison between the true and actual S9000. . . . . . . . . . . . . . . . . . . . . . . . . . 278
4.158 Inviscid velocity distributions for the S9000. . . . . . . . . . . . . . . . . . . . . . . . . . . . 278
4.159 Drag polar for the S9000. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 279
4.160 Lift and moment characteristics for the S9000. . . . . . . . . . . . . . . . . . . . . . . . . . . 280
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.
4 Summary of Low-Speed Airfoil Data
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
6 Summary of Low-Speed Airfoil Data
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.
8 Summary of Low-Speed Airfoil Data
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
UIUC model E: tested Spring 2003
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
UIUC model E: tested Spring 2003
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
UIUC model E: tested Spring 2003
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.
Chapter 2: Experimental Methods 9
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 = 460,000
Figure 2.4: Continued.
10 Summary of Low-Speed Airfoil Data
−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
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 = 200,000
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 = 300,000
LTPT modelLift Moment
UIUC model E: tested Spring 2003Lift Moment
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.
Chapter 2: Experimental Methods 11
−10 0 10 20−0.5
0.0
0.5
1.0
1.5
α (deg)
Cl
E387 Re = 460,000
LTPT modelLift Moment
UIUC model E: tested Spring 2003Lift Moment
Cm
0.1
0.0
−0.1
−0.2
−0.3
Figure 2.5: Continued.
12 Summary of Low-Speed Airfoil Data
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
14 Summary of Low-Speed Airfoil Data
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
16 Summary of Low-Speed Airfoil Data
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.
Chapter 3: Summary of Airfoil Data 17
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.
18 Summary of Low-Speed Airfoil Data
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
Hole diameter = 0.026c
LE to hole center = 0.023c
Spacing = 0.052c
Small Cubes
Cube height = 0.005c
LE to cube front edge = 0.010c
Spacing = 0.010c
Large Cubes
Cube height = 0.021c
LE to cube front edge = 0.010c
Spacing = 0.042c
Fig. 3.4: Baseline flat plate and various leading-edge configurations drawn to scale (but not full span).
Chapter 3: Summary of Airfoil Data 19
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.
20 Summary of Low-Speed Airfoil Data
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
ancePlots
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
Table 4.1: Continued
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
ancePlots
25
Table 4.1: Continued
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
Table 4.1: Continued
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
ancePlots
27
Table 4.1: Continued
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
Table 4.1: Continued
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
ancePlots
29
Table 4.1: Continued
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
ancePlots
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
AG12 (M. Drela, et al.)
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
Avg. Difference = 0.0020 in.
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
Fig.4.6:Inviscidvelocity
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
AG16 (M. Drela, et al.)
−10 0 10 20 −0.5
0.0
0.5
1.0
1.5
2.0
α (deg)
Cl
Fig.4.7: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
AG16 (M. Drela, et al.)
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
AG16 (M. Drela, et al.)
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.8:Liftand
mom
entcharacteristicsforthe
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
AG16 (M. Drela, et al.)
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
AG16 (M. Drela, et al.)
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
AG16 (M. Drela, et al.)
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
AG16 (M. Drela, et al.)
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
AG16 (M. Drela, et al.)
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
Avg. Difference = 0.0047 in.
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
Fig.4.10:Inviscidvelocity
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
AG24 (M. Drela, et al.)
−10 0 10 20 −0.5
0.0
0.5
1.0
1.5
2.0
α (deg)
Cl
Fig.4.11: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
AG24 (M. Drela, et al.)
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
AG24 (M. Drela, et al.)
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.12:Liftand
mom
entcharacteristicsforthe
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
AG24 (M. Drela, et al.)
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
AG24 (M. Drela, et al.)
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
AG24 (M. Drela, et al.)
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
AG24 (M. Drela, et al.)
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
AG24 (M. Drela, et al.)
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)
Avg. Difference = 0.0042 in.
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
Fig.4.14:Inviscidvelocity
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
AG35−r (M. Drela, et al.)
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
AG35−r (M. Drela, et al.)
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.16:Liftand
mom
entcharacteristicsforthe
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
AG35−r (M. Drela, et al.)
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
AG35−r (M. Drela, et al.)
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
AG35−r (M. Drela, et al.)
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
AG35−r (M. Drela, et al.)
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
Avg. Difference = 0.0069 in.
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
Fig.4.18:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.20:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
AG40d−02r (M. Drela, et al.)Flap = 0 deg
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
Fig.4.21:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.23:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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
AG40d−02r (M. Drela, et al.)Flap = −2 deg
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∞
AG40d−02r Flap 2 deg
Fig.4.24:Inviscidvelocity
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
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.4.25: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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.26:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
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
AG40d−02r (M. Drela, et al.)Flap = 2 deg
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∞
AG40d−02r Flap 4 deg
Fig.4.27:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
−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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.29:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
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
AG40d−02r (M. Drela, et al.)Flap = 4 deg
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∞
AG40d−02r Flap −15 deg
Fig.4.30:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = −15 deg
−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
AG40d−02r (M. Drela, et al.)Flap = −15 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.32:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap −10 deg
Fig.4.33:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = −10 deg
−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
AG40d−02r (M. Drela, et al.)Flap = −10 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.35:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap −5 deg
Fig.4.36:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = −5 deg
−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
AG40d−02r (M. Drela, et al.)Flap = −5 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.38:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap 5 deg
Fig.4.39:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 5 deg
−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
AG40d−02r (M. Drela, et al.)Flap = 5 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.41:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap 10 deg
Fig.4.42:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 10 deg
−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
AG40d−02r (M. Drela, et al.)Flap = 10 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.44:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap 15 deg
Fig.4.45:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 15 deg
−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
AG40d−02r (M. Drela, et al.)Flap = 15 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.47:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap 20 deg
Fig.4.48:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 20 deg
−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
AG40d−02r (M. Drela, et al.)Flap = 20 deg
Re = 40,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.50:Liftand
mom
entcharacteristicsforthe
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∞
AG40d−02r Flap 30 deg
Fig.4.51:Inviscidvelocity
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
AG40d−02r (M. Drela, et al.)Flap = 30 deg
Re = 40,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.52:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 0 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = 0 deggap sealed
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.54:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 0 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = −2 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = −2 deggap sealed
Re = 200,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.56:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = −2 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = −2 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = 4 deggap sealed
−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
AG40d−02r (M. Drela, et al.)Flap = 4 deggap sealed
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
AG40d−02r (M. Drela, et al.)Flap = 4 deggap sealed
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.58:Liftand
mom
entcharacteristicsforthe
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
AG40d−02r (M. Drela, et al.)Flap = 4 deggap sealed
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
Avg. Difference = 0.0080 in.
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
Fig.4.61:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.63:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deg
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∞
AG455ct−02r Flap −2.4 deg
Fig.4.64:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
Re = 80,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.66:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deg
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
Fig.4.67:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.69:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 1.6 deg
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∞
AG455ct−02r Flap 3.6 deg
Fig.4.70:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.72:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deg
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∞
AG455ct−02r Flap −15.4 deg
Fig.4.73:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = −15.4 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = −15.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −15.4 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.75:Liftand
mom
entcharacteristicsforthe
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∞
AG455ct−02r Flap −10.4 deg
Fig.4.76:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = −10.4 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = −10.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −10.4 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.78:Liftand
mom
entcharacteristicsforthe
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∞
AG455ct−02r Flap −5.4 deg
Fig.4.79:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = −5.4 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = −5.4 deg
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
AG455ct−02r (M. Drela, et al.)Flap = −5.4 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.81:Liftand
mom
entcharacteristicsforthe
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∞
AG455ct−02r Flap 4.6 deg
Fig.4.82:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = 4.6 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = 4.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 4.6 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.84:Liftand
mom
entcharacteristicsforthe
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∞
AG455ct−02r Flap 9.6 deg
Fig.4.85:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = 9.6 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = 9.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 9.6 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.87:Liftand
mom
entcharacteristicsforthe
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∞
AG455ct−02r Flap 14.6 deg
Fig.4.88:Inviscidvelocity
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
AG455ct−02r (M. Drela, et al.)Flap = 14.6 deg
−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
AG455ct−02r (M. Drela, et al.)Flap = 14.6 deg
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
AG455ct−02r (M. Drela, et al.)Flap = 14.6 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.90:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deggap sealed
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
AG455ct−02r (M. Drela, et al.)Flap = −0.4 deggap sealed
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.92:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deggap sealed
−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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deggap sealed
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
AG455ct−02r (M. Drela, et al.)Flap = −2.4 deggap sealed
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.94:Liftand
mom
entcharacteristicsforthe
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deggap sealed
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
AG455ct−02r (M. Drela, et al.)Flap = 3.6 deggap sealed
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.96:Liftand
mom
entcharacteristicsforthe
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
Avg. Difference = 0.0080 in.
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
Fig.4.100:Inviscidvelocity
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
entcharacteristicsforthe
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
Fig.4.102:Continued.
CA
L1215j
Chapter4:
AirfoilProfiles
andPerform
ancePlots
183
184Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
CAL2263m
Avg. Difference = 0.0063 in.
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
Fig.4.104:Inviscidvelocity
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
CAL2263m (C. Greaves)
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
CAL2263m (C. Greaves)
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.106:Liftand
mom
entcharacteristicsforthe
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
CAL2263m (C. Greaves)
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
CAL2263m (C. Greaves)
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.106:Continued.
CA
L2263m
188Sum
mary
ofLow
-SpeedA
irfoilData
−10 0 10 20 −0.5
0.0
0.5
1.0
1.5
α (deg)
Cl
CAL2263m (C. Greaves)
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
CAL2263m (C. Greaves)
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.106:Continued.
CA
L2263m
Chapter4:
AirfoilProfiles
andPerform
ancePlots
189
190Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
CAL4014l
Avg. Difference = 0.0094 in.
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
Fig.4.108:Inviscidvelocity
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
CAL4014l (J. Thomas)
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
CAL4014l (J. Thomas)
Re = 200,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.110:Liftand
mom
entcharacteristicsforthe
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
CAL4014l (J. Thomas)
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
CAL4014l (J. Thomas)
Re = 400,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.110:Continued.
CA
L4014l
194Sum
mary
ofLow
-SpeedA
irfoilData
−10 0 10 20 −0.5
0.0
0.5
1.0
1.5
α (deg)
Cl
CAL4014l (J. Thomas)
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.110:Continued.
CA
L4014l
Chapter4:
AirfoilProfiles
andPerform
ancePlots
195
196Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
E387 (E)
Avg. Difference = 0.0091 in.
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
Fig.4.112:Inviscidvelocity
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
entcharacteristicsforthe
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
Fig.4.114:Continued.
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
Fig.4.114:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickBaseline
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
Flat Plate (C. Goudeseune)3.1% thickBaseline
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
Flat Plate (C. Goudeseune)3.1% thickBaseline
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.116:Continued.
FlatPlateB
aseline
Chapter4:
AirfoilProfiles
andPerform
ancePlots
205
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.116:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations A
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations A
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations A
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.118:Continued.
FlatPlateSerrations
A
Chapter4:
AirfoilProfiles
andPerform
ancePlots
209
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.118:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations B
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.120:Liftand
mom
entcoefficientcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations B
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations B
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.120:Continued.
FlatPlateSerrations
B
Chapter4:
AirfoilProfiles
andPerform
ancePlots
213
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.120:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations C
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.122:Liftand
mom
entcoefficientcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations C
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations C
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.122:Continued.
FlatPlateSerrations
C
Chapter4:
AirfoilProfiles
andPerform
ancePlots
217
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.122:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations D
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.124:Liftand
mom
entcoefficientcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations D
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
Flat Plate (C. Goudeseune)3.1% thickLeading Edge Serrations D
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.124:Continued.
FlatPlateSerrations
D
Chapter4:
AirfoilProfiles
andPerform
ancePlots
221
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.124:Continued.
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
Flat Plate (C. Goudeseune)3.1% thickSquare Wave
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.126:Liftand
mom
entcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickSquare Wave
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
Flat Plate (C. Goudeseune)3.1% thickSquare Wave
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.126:Continued.
FlatPlateSquare
Wave
Chapter4:
AirfoilProfiles
andPerform
ancePlots
225
−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 = 120,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.126:Continued.
FlatPlateSquare
Wave
226Sum
mary
ofLow
-SpeedA
irfoilData
Small Holes
Hole diameter = 0.008c
LE to hole center = 0.014c
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
Flat Plate (C. Goudeseune)3.1% thickSmall Holes
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.128:Liftand
mom
entcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickSmall Holes
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.128:Continued.
FlatPlateSm
allHoles
Chapter4:
AirfoilProfiles
andPerform
ancePlots
229
230Sum
mary
ofLow
-SpeedA
irfoilData
Large Holes
Hole diameter = 0.026c
LE to hole center = 0.023c
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
entcharacteristicsfora
flatplatew
ithlarge
holeson
theleading
edge.
FlatPlateL
argeH
oles
232Sum
mary
ofLow
-SpeedA
irfoilData
Small Cubes
Cube height = 0.005c
LE to cube front edge = 0.010c
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
Flat Plate (C. Goudeseune)3.1% thickSmall Cubes
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
Cm
Fig.4.132:Liftand
mom
entcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickSmall Cubes
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
Cm
Fig.4.132:Continued.
FlatPlateSm
allCubes
Chapter4:
AirfoilProfiles
andPerform
ancePlots
235
236Sum
mary
ofLow
-SpeedA
irfoilData
Large Cubes
Cube height = 0.021c
LE to cube front edge = 0.010c
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
Flat Plate (C. Goudeseune)3.1% thickLarge Cubes
Re = 60,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.134:Liftand
mom
entcharacteristicsfora
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
Flat Plate (C. Goudeseune)3.1% thickLarge Cubes
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.134:Continued.
FlatPlateL
argeC
ubes
Chapter4:
AirfoilProfiles
andPerform
ancePlots
239
240Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
MA409
Avg. Difference = 0.0358 in.
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
Fig.4.136:Inviscidvelocity
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
entcharacteristicsforthe
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
Fig.4.138:Continued.
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
Fig.4.138:Continued.
MA
409
Chapter4:
AirfoilProfiles
andPerform
ancePlots
245
246Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
NACA 43012A
Avg. Difference = 0.0430 in.
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
Fig.4.140:Inviscidvelocity
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
NACA 43012A (M. Nankivil)
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
NACA 43012A (M. Nankivil)
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.142:Liftand
mom
entcharacteristicsforthe
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
NACA 43012A (M. Nankivil)
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
NACA 43012A (M. Nankivil)
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.142:Continued.
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
NACA 43012A (M. Nankivil)
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
NACA 43012A (M. Nankivil)
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.142:Continued.
NA
CA
43012A
Chapter4:
AirfoilProfiles
andPerform
ancePlots
251
252Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
S1223
Avg. Difference = 0.0099 in.
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
Fig.4.144:Inviscidvelocity
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
S1223 (Y. Tinel)clean
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.145:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)clean
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
S1223 (Y. Tinel)clean
Re = 140,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.145:Continued.
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
S1223 (Y. Tinel)clean
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)clean
Re = 180,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.145:Continued.
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
S1223 (Y. Tinel)clean
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
S1223 (Y. Tinel)clean
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.145:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 4.17%
Re = 180,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.146:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)Gurney Flap: h/c = 4.17%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 4.17%
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.146:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 3.12%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 3.12%
Re = 160,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.147:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)Gurney Flap: h/c = 3.12%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 3.12%
Re = 200,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.147:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 3.12%
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.147:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 2.08%
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)Gurney Flap: h/c = 2.08%
Re = 180,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.148:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)Gurney Flap: h/c = 2.08%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 2.08%
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.148:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.56%
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)Gurney Flap: h/c = 1.56%
Re = 180,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.149:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.56%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.56%
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.149:Continued.
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.04%
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)Gurney Flap: h/c = 1.04%
Re = 180,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.150:Liftand
mom
entcharacteristicsforthe
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.04%
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
S1223 (Y. Tinel)Gurney Flap: h/c = 1.04%
Re = 250,000
0.1
0.0
−0.1
−0.2
−0.3
−0.4
−0.5
Cm
Fig.4.150:Continued.
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
entcharacteristicsforthe
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
entcharacteristicsforthe
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
Avg. Difference = 0.0065 in.
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
Fig.4.154:Inviscidvelocity
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
entcharacteristicsforthe
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
Fig.4.156:Continued.
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
Fig.4.156:Continued.
S8064
Chapter4:
AirfoilProfiles
andPerform
ancePlots
277
278Sum
mary
ofLow
-SpeedA
irfoilData
Diff
eren
ce (
in)
−0.03
0.00
0.03
S9000
Avg. Difference = 0.0055 in.
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
Fig.4.158:Inviscidvelocity
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
S9000 (T. Akers)Flap = 0 deg
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
S9000 (T. Akers)Flap = 0 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.160:Liftand
mom
entcharacteristicsforthe
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
S9000 (T. Akers)Flap = 0 deg
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
S9000 (T. Akers)Flap = 0 deg
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.160:Continued.
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
S9000 (T. Akers)Flap = 0 deg
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
S9000 (T. Akers)Flap = 0 deg
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.160:Continued.
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
Fig.4.161:Inviscidvelocity
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
S9000 (T. Akers)Flap = 2.5 deg
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
S9000 (T. Akers)Flap = 2.5 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.163:Liftand
mom
entcharacteristicsforthe
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
S9000 (T. Akers)Flap = 2.5 deg
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
S9000 (T. Akers)Flap = 2.5 deg
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.163:Continued.
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
S9000 (T. Akers)Flap = 2.5 deg
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
S9000 (T. Akers)Flap = 2.5 deg
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.163:Continued.
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
Fig.4.164:Inviscidvelocity
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
S9000 (T. Akers)Flap = 5 deg
−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
S9000 (T. Akers)Flap = 5 deg
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
S9000 (T. Akers)Flap = 5 deg
Re = 100,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.166:Liftand
mom
entcharacteristicsforthe
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
S9000 (T. Akers)Flap = 5 deg
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
S9000 (T. Akers)Flap = 5 deg
Re = 300,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.166:Continued.
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
S9000 (T. Akers)Flap = 5 deg
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
S9000 (T. Akers)Flap = 5 deg
Re = 500,000
0.1
0.0
−0.1
−0.2
−0.3
Cm
Fig.4.166:Continued.
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.
296 Summary of Low-Speed Airfoil Data
[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.
298 Summary of Low-Speed Airfoil Data
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
300 Summary of Low-Speed Airfoil Data
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
Appendix A: Tabulated Airfoil Coordinates 301
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
302 Summary of Low-Speed Airfoil Data
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
Appendix A: Tabulated Airfoil Coordinates 303
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
304 Summary of Low-Speed Airfoil Data
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
Appendix A: Tabulated Airfoil Coordinates 305
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
306 Summary of Low-Speed Airfoil Data
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
Appendix A: Tabulated Airfoil Coordinates 307
0.484548 0.0415450.470723 0.0424690.456908 0.0433540.443086 0.0441990.429265 0.0450000.415452 0.0457550.401634 0.0464620.387821 0.0471180.374011 0.0477210.360206 0.0482690.346397 0.0487550.332597 0.0491780.318804 0.0495350.305013 0.0498220.291233 0.0500320.277461 0.0501630.263691 0.0502060.249936 0.0501560.236200 0.0500070.222471 0.0497520.208767 0.0493800.195088 0.0488820.181432 0.0482470.167822 0.0474620.154250 0.0465140.140736 0.0453840.127294 0.0440550.113930 0.0425000.100688 0.0406980.087581 0.0386110.074669 0.0361990.061994 0.0334060.049680 0.0301760.037910 0.0264600.027072 0.0222690.017878 0.0178380.011050 0.0136800.006524 0.0101790.003631 0.0073210.001787 0.0049340.000668 0.0028820.000113 0.0011200.000013 -0.0003590.000337 -0.0017260.001239 -0.0030950.002858 -0.0043900.005244 -0.0056650.008721 -0.0070250.013948 -0.0085490.021760 -0.0102080.032064 -0.0117830.043924 -0.0130760.056554 -0.014057
0.069555 -0.0147760.082773 -0.0152860.096139 -0.0156330.109606 -0.0158540.123148 -0.0159750.136750 -0.0160130.150401 -0.0159770.164083 -0.0158790.177803 -0.0157300.191544 -0.0155370.205308 -0.0153100.219101 -0.0150560.232927 -0.0147810.246776 -0.0144880.260659 -0.0141860.274574 -0.0138720.288521 -0.0135510.302495 -0.0132260.316489 -0.0128970.330500 -0.0125660.344528 -0.0122350.358569 -0.0119040.372617 -0.0115720.386681 -0.0112410.400763 -0.0109090.414858 -0.0105790.428946 -0.0102500.443025 -0.0099230.457091 -0.0095990.471152 -0.0092740.485206 -0.0089540.499244 -0.0086360.513274 -0.0083210.527300 -0.0080060.541320 -0.0076940.548893 -0.0075260.558509 -0.0073110.568829 -0.0070800.582291 -0.0067810.595470 -0.0064880.609368 -0.0061770.622968 -0.0058750.636569 -0.0055740.646155 -0.0053630.655175 -0.0051660.663777 -0.0049780.677452 -0.0046820.691121 -0.0043880.696993 -0.0042630.699968 -0.0042000.702041 -0.0041590.706272 -0.0040690.718805 -0.003804
0.732681 -0.0035150.746544 -0.0032300.760397 -0.0029500.774251 -0.0026720.788103 -0.0023990.801961 -0.0021280.815822 -0.0018640.829684 -0.0016090.843548 -0.0013740.857420 -0.0011550.871290 -0.0009560.885162 -0.0007780.899044 -0.0006220.912921 -0.0004940.926805 -0.0003910.940689 -0.0003210.954576 -0.0002850.968440 -0.0002850.982058 -0.0003320.994081 -0.0003891.000000 -0.000422
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
308 Summary of Low-Speed Airfoil Data
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
0.692137 -0.0064520.729257 -0.0053930.765908 -0.0044450.797483 -0.0036900.824447 -0.0030990.849501 -0.0024890.871630 -0.0020330.892923 -0.0017640.913887 -0.0013550.931490 -0.0011370.946132 -0.0008840.966531 -0.0006020.985756 -0.0002720.991795 -0.0002200.995876 -0.0001280.998199 -0.0000721.000000 0.000000
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
Appendix A: Tabulated Airfoil Coordinates 309
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
310 Summary of Low-Speed Airfoil Data
0.078505 -0.0244800.101286 -0.0260020.126632 -0.0270420.154419 -0.0276580.184458 -0.0278670.216596 -0.0276450.250678 -0.0270290.286535 -0.0260620.323976 -0.0248400.362746 -0.0234420.402597 -0.0219180.443231 -0.0203940.484280 -0.0189050.525453 -0.0174020.566477 -0.0159030.607073 -0.0144130.646976 -0.0129440.685916 -0.0115130.723625 -0.0101320.759847 -0.0088110.794324 -0.0075580.826807 -0.0063490.857133 -0.0051840.885126 -0.0041330.910519 -0.0032470.933066 -0.0025200.952540 -0.0019020.968826 -0.0012790.981917 -0.0006600.991686 -0.0002080.997858 -0.0000011.000000 0.000000
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
0.421284 -0.0225880.457378 -0.0212700.501846 -0.0196350.549214 -0.0179520.584185 -0.0167910.623149 -0.0154130.659459 -0.0141120.691980 -0.0128510.726601 -0.0117120.757760 -0.0105660.788109 -0.0093580.817902 -0.0081740.843384 -0.0071660.869655 -0.0064220.892902 -0.0055240.913603 -0.0049420.931142 -0.0044050.948070 -0.0038310.963402 -0.0031420.977236 -0.0024330.989560 -0.0016720.995584 -0.0010010.998133 -0.0006671.000000 -0.000519
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
Appendix A: Tabulated Airfoil Coordinates 311
0.314605 0.0677210.281293 0.0682110.249304 0.0679330.218748 0.0669010.189818 0.0651230.162606 0.0625910.137210 0.0593830.113769 0.0554880.092328 0.0509470.072978 0.0458910.055797 0.0403520.040807 0.0344320.028111 0.0282620.017739 0.0219340.009713 0.0156350.004152 0.0094440.000990 0.0034980.000203 -0.0015710.002780 -0.0058270.008926 -0.0100180.017894 -0.0142290.029532 -0.0182620.043702 -0.0219600.060381 -0.0250750.079676 -0.0274940.101680 -0.0293150.126257 -0.0306830.153203 -0.0316230.182434 -0.0321420.213752 -0.0323050.246948 -0.0321010.281858 -0.0315970.318239 -0.0308380.355875 -0.0298250.394527 -0.0286220.433927 -0.0272600.473841 -0.0257650.513997 -0.0241800.554120 -0.0225210.593957 -0.0208230.633232 -0.0191120.671683 -0.0174040.709053 -0.0157280.745082 -0.0140930.779530 -0.0125120.812161 -0.0109990.842749 -0.0095580.871088 -0.0081940.896983 -0.0069090.920257 -0.0056990.940753 -0.0045590.958332 -0.0034820.972895 -0.002429
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
0.009944 0.0169990.008645 0.0155920.007365 0.0140000.005963 0.0120250.003811 0.0090500.002755 0.0074650.000759 0.0036790.000413 0.0027240.000164 0.0017390.000018 -0.0006590.000210 -0.0022990.001311 -0.0055990.003509 -0.0090910.007155 -0.0122030.010249 -0.0140120.016673 -0.0170920.025837 -0.0208800.035880 -0.0239840.051590 -0.0269920.071165 -0.0295210.092222 -0.0310250.115775 -0.0322540.148051 -0.0335140.177335 -0.0339840.206423 -0.0339200.248809 -0.0332170.287053 -0.0322890.322651 -0.0311980.354971 -0.0301050.392030 -0.0285640.427645 -0.0273320.465210 -0.0259290.506588 -0.0242610.556367 -0.0220300.593242 -0.0203340.626137 -0.0187480.660805 -0.0172250.705228 -0.0153350.737137 -0.0139390.767945 -0.0126320.793929 -0.0114620.822971 -0.0103140.855302 -0.0092530.878095 -0.0085380.899624 -0.0078200.920093 -0.0069840.937930 -0.0060480.954096 -0.0048560.979982 -0.0024880.985094 -0.0020430.990136 -0.0015470.993556 -0.0012400.996992 -0.000891
312 Summary of Low-Speed Airfoil Data
1.000000 -0.000683
E387True
x/c y/c1.000000 0.0000000.996770 0.0004300.987290 0.0018000.971980 0.0042300.951280 0.0076300.925540 0.0118400.895100 0.0167900.860350 0.0224200.821830 0.0286600.780070 0.0354000.735670 0.0424900.689220 0.0497500.641360 0.0569600.592720 0.0639000.543940 0.0702000.495490 0.0754600.447670 0.0793600.400770 0.0817300.355050 0.0824700.310780 0.0815600.268130 0.0790800.227420 0.0752900.189060 0.0703700.153450 0.0644800.120940 0.0577500.091850 0.0503300.066430 0.0423800.044930 0.0340800.027480 0.0256200.014230 0.0172600.005190 0.0093100.000440 0.0023400.000000 0.0000000.000910 -0.0028600.007170 -0.0068200.018900 -0.0101700.035960 -0.0126500.058270 -0.0142500.085690 -0.0150000.118000 -0.0150200.154900 -0.0144100.195990 -0.0132900.240830 -0.0117700.288920 -0.0099800.339680 -0.0080400.392520 -0.0060500.446790 -0.004100
0.501820 -0.0022800.556940 -0.0006500.611470 0.0007400.664720 0.0018600.716020 0.0026800.764750 0.0032000.810270 0.0034200.852020 0.0033700.889440 0.0030700.922050 0.0025800.949420 0.0019600.971180 0.0013200.987050 0.0007100.996740 0.0002101.000000 0.000000
E387 (E)Actual
x/c y/c1.000000 0.0007060.999678 0.0007910.995017 0.0018420.991921 0.0024370.982490 0.0040610.973277 0.0055910.958633 0.0080620.926344 0.0133430.909852 0.0160570.889160 0.0194650.866125 0.0232220.842670 0.0270470.817387 0.0311540.787247 0.0360400.759532 0.0404960.726804 0.0457020.694176 0.0508030.659004 0.0561980.622656 0.0615600.589064 0.0662420.551907 0.0709690.510670 0.0755320.476873 0.0786470.437853 0.0813960.397082 0.0831510.359760 0.0836900.284936 0.0813420.249194 0.0785510.209825 0.0741820.179829 0.0698430.144677 0.0635110.118041 0.0576820.092257 0.050985
0.066657 0.0428780.045215 0.0344720.027607 0.0258730.015156 0.0182580.008243 0.0128480.006684 0.0113660.004741 0.0092840.003353 0.0076090.002206 0.0060120.001354 0.0046320.000639 0.0031480.000136 0.0014890.000002 0.0001640.000017 -0.0005230.000245 -0.0017060.000972 -0.0031670.001656 -0.0040320.004841 -0.0063610.007808 -0.0077240.013461 -0.0095420.018565 -0.0106770.030536 -0.0123380.045514 -0.0135060.065671 -0.0144110.089604 -0.0149360.114136 -0.0150070.143028 -0.0146870.171384 -0.0140740.202734 -0.0131780.235535 -0.0120820.273934 -0.0106610.306347 -0.0094420.342832 -0.0080820.378570 -0.0067760.416413 -0.0054310.457857 -0.0039840.496093 -0.0027020.539972 -0.0013240.583917 -0.0000820.620305 0.0008110.656351 0.0015770.688568 0.0021290.723435 0.0026270.754643 0.0029820.785321 0.0031850.811327 0.0032920.839117 0.0033220.864230 0.0032550.888306 0.0030760.910334 0.0028410.926935 0.0025100.944962 0.0019380.961010 0.001195
Appendix A: Tabulated Airfoil Coordinates 313
0.973032 0.0005430.988335 -0.0002560.993532 -0.0004920.997747 -0.0006190.999577 -0.0005880.999959 -0.000563
MA409True
x/c y/c1.000000 0.0003400.997540 0.0009400.990700 0.0025900.980370 0.0049800.966980 0.0079300.950440 0.0113600.930640 0.0152100.907750 0.0193700.882020 0.0237300.853700 0.0281700.823090 0.0326100.790480 0.0369400.756160 0.0411200.720430 0.0450500.683590 0.0486900.645940 0.0519800.607780 0.0548600.569370 0.0573200.530990 0.0593300.492650 0.0608900.454350 0.0620200.416380 0.0627000.378870 0.0629100.342040 0.0626000.306090 0.0617200.271200 0.0602500.237600 0.0581900.205490 0.0555500.175040 0.0523400.146480 0.0485900.119990 0.0443300.095760 0.0396100.073950 0.0345000.054680 0.0291300.038110 0.0236900.024330 0.0183100.013380 0.0130500.005480 0.0079700.000980 0.0031800.000000 -0.0000400.000980 -0.0030200.005480 -0.006640
0.013380 -0.0096800.024330 -0.0121300.038110 -0.0140000.054680 -0.0152700.073950 -0.0159000.095760 -0.0158900.119990 -0.0152700.146480 -0.0141600.175040 -0.0126400.205490 -0.0108100.237600 -0.0087500.271200 -0.0065300.306090 -0.0042100.342040 -0.0018600.378870 0.0003900.416380 0.0024500.454350 0.0042200.492650 0.0056500.530990 0.0067400.569370 0.0075400.607780 0.0080800.645940 0.0083700.683590 0.0084400.720430 0.0083000.756160 0.0079800.790480 0.0074900.823090 0.0068700.853700 0.0061300.882020 0.0052900.907750 0.0043900.930640 0.0034700.950440 0.0025600.966980 0.0017000.980370 0.0009300.990700 0.0002800.997540 -0.0001801.000000 -0.000360
MA409Actual
x/c y/c1.000000 0.0005090.997499 0.0009250.994830 0.0011390.989837 0.0015180.982926 0.0020520.966884 0.0034380.947351 0.0055540.934012 0.0071860.914492 0.0096240.895225 0.0120250.871906 0.014819
0.850598 0.0172290.825905 0.0199620.792359 0.0235640.763752 0.0264260.731226 0.0294140.694149 0.0325690.658309 0.0355270.623556 0.0384120.592373 0.0409830.559586 0.0432790.524180 0.0456350.496579 0.0472640.462524 0.0488880.437954 0.0500770.400386 0.0517180.372610 0.0525760.334397 0.0528700.307960 0.0529820.274972 0.0529850.248182 0.0519210.214771 0.0504250.183639 0.0480790.151810 0.0451080.128411 0.0425780.113265 0.0406390.095435 0.0379990.082268 0.0356350.069908 0.0330690.056090 0.0297670.043087 0.0259010.032164 0.0219090.023606 0.0183110.017513 0.0154050.009968 0.0107500.005977 0.0074860.003195 0.0051260.003079 -0.0049610.007431 -0.0077040.013830 -0.0119270.021809 -0.0140740.030907 -0.0159150.044523 -0.0175880.066194 -0.0189150.086095 -0.0193030.101862 -0.0191990.122932 -0.0191110.139429 -0.0188040.160927 -0.0182860.185238 -0.0171300.201283 -0.0165310.225861 -0.0157940.246533 -0.0149660.276677 -0.013559
314 Summary of Low-Speed Airfoil Data
0.308552 -0.0122290.340310 -0.0109270.374385 -0.0098530.396903 -0.0091260.421929 -0.0082590.453788 -0.0074500.484129 -0.0070720.510810 -0.0069230.538209 -0.0067940.568729 -0.0079430.613202 -0.0069700.638337 -0.0054870.664232 -0.0061220.690338 -0.0064280.715749 -0.0065290.733734 -0.0066610.757821 -0.0067140.778920 -0.0067180.800193 -0.0065340.817179 -0.0062850.850217 -0.0057080.868449 -0.0053590.885743 -0.0050460.900293 -0.0048550.924159 -0.0044160.936847 -0.0040520.950413 -0.0034730.959932 -0.0029920.969027 -0.0024730.975625 -0.0020590.982723 -0.0016190.989785 -0.0011370.995597 -0.0007370.998783 -0.0005761.000000 -0.000472
NACA 43012ATrue
x/c y/c1.000000 0.0000000.900000 0.0125000.800000 0.0244000.700000 0.0370000.600000 0.0520000.500000 0.0656000.400000 0.0770000.300000 0.0857000.250000 0.0895000.200000 0.0927000.150000 0.0933000.100000 0.0877000.075000 0.080300
0.050000 0.0692000.025000 0.0514000.012500 0.0389000.000000 0.0000000.012500 -0.0083000.025000 -0.0114000.050000 -0.0154000.075000 -0.0185000.100000 -0.0211000.150000 -0.0257000.200000 -0.0295000.250000 -0.0327000.300000 -0.0350000.400000 -0.0392000.500000 -0.0403000.600000 -0.0392000.700000 -0.0360000.800000 -0.0274000.900000 -0.0150001.000000 0.000000
NACA 43012AActual
x/c y/c1.000000 0.0015180.997750 0.0019000.995724 0.0022560.990020 0.0030190.979796 0.0042440.968135 0.0054710.955557 0.0070180.940824 0.0086190.925987 0.0102870.904721 0.0132320.884663 0.0162570.865924 0.0189200.843371 0.0224050.817029 0.0264410.782493 0.0312650.753529 0.0352830.733420 0.0382790.702819 0.0425870.667395 0.0479280.633085 0.0526330.605645 0.0562060.567177 0.0608810.530326 0.0648780.488272 0.0692430.451381 0.0725690.410885 0.0760450.365460 0.0799180.326461 0.082886
0.290173 0.0843530.248795 0.0850570.228965 0.0849570.202726 0.0842840.174660 0.0829400.153001 0.0811040.133118 0.0785640.109289 0.0740970.089771 0.0689040.072680 0.0629320.063317 0.0589930.050340 0.0526750.038853 0.0463190.030350 0.0410260.024916 0.0373900.019806 0.0336340.013172 0.0280070.010362 0.0252380.007582 0.0221440.004979 0.0186100.002851 0.0147110.000980 0.0086210.000242 0.0024090.000089 -0.0014350.003055 -0.0084770.008634 -0.0123350.015128 -0.0147280.021822 -0.0163340.028047 -0.0175530.032667 -0.0182940.056372 -0.0210950.076542 -0.0233720.105869 -0.0261880.133483 -0.0281910.161079 -0.0297910.201880 -0.0316710.234082 -0.0329250.268328 -0.0338970.302858 -0.0348230.343432 -0.0356710.384234 -0.0362490.420098 -0.0358700.449387 -0.0351750.493851 -0.0337870.531309 -0.0325440.572644 -0.0308900.608351 -0.0290240.644957 -0.0272100.679432 -0.0257570.713098 -0.0239030.741853 -0.0219900.771712 -0.0197990.797786 -0.017839
Appendix A: Tabulated Airfoil Coordinates 315
0.826321 -0.0154610.829250 -0.0151710.849452 -0.0136100.871984 -0.0118760.890292 -0.0104310.911533 -0.0089370.929652 -0.0075660.942286 -0.0065940.951796 -0.0058690.960961 -0.0050440.968828 -0.0043990.975406 -0.0039650.982705 -0.0033020.987785 -0.0028040.995158 -0.0021560.998513 -0.0018121.000000 -0.001422
S1223True
x/c y/c1.000000 0.0000000.998380 0.0012600.994170 0.0049400.988250 0.0103700.980750 0.0164600.971110 0.0225000.958840 0.0285300.943890 0.0347600.926390 0.0411600.906410 0.0476800.884060 0.0542700.859470 0.0608900.832770 0.0674900.804120 0.0740200.773690 0.0804400.741660 0.0867100.708230 0.0927700.673600 0.0985900.637980 0.1041200.601580 0.1093500.564650 0.1142500.527440 0.1188100.490250 0.1230300.453400 0.1268300.417210 0.1301100.381930 0.1327100.347770 0.1344700.314880 0.1352600.283470 0.1350500.253700 0.1334600.225410 0.130370
0.198460 0.1259400.172860 0.1202600.148630 0.1135500.125910 0.1059800.104820 0.0977000.085450 0.0887900.067890 0.0794000.052230 0.0696500.038550 0.0596800.026940 0.0496600.017550 0.0396100.010280 0.0295400.004950 0.0196900.001550 0.0103300.000050 0.0017800.000440 -0.0056100.002640 -0.0112000.007890 -0.0142700.017180 -0.0155000.030060 -0.0158400.046270 -0.0153200.065610 -0.0140400.087870 -0.0120200.112820 -0.0092500.140200 -0.0056300.170060 -0.0007500.202780 0.0053500.238400 0.0121300.276730 0.0192800.317500 0.0265200.360440 0.0335800.405190 0.0402100.451390 0.0461800.498600 0.0512900.546390 0.0553400.594280 0.0582000.641760 0.0597600.688320 0.0599400.733440 0.0587200.776600 0.0561200.817290 0.0521900.855000 0.0470600.889280 0.0408800.919660 0.0338700.945730 0.0262400.966930 0.0182200.982550 0.0106000.992680 0.0046800.998250 0.0011501.000000 0.000000
S1223Actual
x/c y/c1.000000 0.0001300.997196 0.0033250.991871 0.0083900.986890 0.0126010.980752 0.0169150.973390 0.0212700.962928 0.0265950.953698 0.0307980.941761 0.0358070.925483 0.0420370.910530 0.0472840.889687 0.0539900.871785 0.0592390.844957 0.0663750.817442 0.0730530.783341 0.0806120.748017 0.0877880.718685 0.0932800.681526 0.0997580.650377 0.1048310.610719 0.1108670.590706 0.1136820.564072 0.1172000.535418 0.1207420.505785 0.1241890.478633 0.1271760.446406 0.1303950.416096 0.1330400.387522 0.1350860.359622 0.1366460.328535 0.1377110.296479 0.1379440.263559 0.1367860.232630 0.1339260.212833 0.1311380.195609 0.1280230.174734 0.1233570.157333 0.1186850.140322 0.1134790.124973 0.1081940.110091 0.1024700.095501 0.0961310.080175 0.0886260.067997 0.0819870.048966 0.0700040.035680 0.0599870.028461 0.0536860.019887 0.0448710.015032 0.038838
316 Summary of Low-Speed Airfoil Data
0.010818 0.0325670.007061 0.0258700.004819 0.0210030.002214 0.0136550.000797 0.0074510.000187 0.0028570.000046 -0.0013780.000833 -0.0059320.001395 -0.0068980.003122 -0.0087180.007902 -0.0106810.014494 -0.0114980.021376 -0.0118490.033297 -0.0119620.046269 -0.0115350.063894 -0.0104040.081174 -0.0088810.103535 -0.0066250.129482 -0.0036850.164036 0.0013430.201003 0.0078140.251147 0.0169010.296102 0.0247900.333923 0.0310020.381597 0.0381570.430060 0.0445490.479399 0.0500440.526540 0.0543170.576836 0.0574740.631716 0.0593890.676492 0.0598040.715586 0.0592080.736746 0.0584410.763328 0.0570290.789418 0.0550650.809470 0.0531000.830277 0.0506220.869940 0.0444570.899245 0.0383790.921488 0.0327440.944791 0.0256090.957067 0.0211660.971768 0.0149260.989001 0.0057040.992314 0.0036300.997998 -0.0001271.000000 -0.001777
S8064True
x/c y/c1.000000 0.0000000.997927 0.0001250.991844 0.0006610.981950 0.0017590.968438 0.0035650.951588 0.0061930.931716 0.0096860.909164 0.0140350.884293 0.0191680.857469 0.0249540.829053 0.0311980.799390 0.0376300.768826 0.0437600.737098 0.0491610.703928 0.0539090.669588 0.0581800.634301 0.0619610.598271 0.0652290.561701 0.0679470.524799 0.0700790.487769 0.0716010.450799 0.0725000.414069 0.0727550.377770 0.0723740.342087 0.0713870.307218 0.0698310.273369 0.0677360.240738 0.0651340.209521 0.0620540.179900 0.0585310.152050 0.0546000.126136 0.0503020.102312 0.0456780.080718 0.0407720.061481 0.0356360.044723 0.0303220.030576 0.0248730.019091 0.0193100.010271 0.0136700.004145 0.0080470.000637 0.0026920.000328 -0.0019110.003568 -0.0063250.009851 -0.0110820.018933 -0.0158120.030792 -0.0204020.045404 -0.0247680.062686 -0.0288600.082555 -0.032657
0.104908 -0.0361430.129620 -0.0392990.156536 -0.0421180.185485 -0.0445850.216282 -0.0466830.248732 -0.0483880.282633 -0.0496800.317775 -0.0505300.353959 -0.0509050.390991 -0.0507950.428663 -0.0501840.466779 -0.0490710.505162 -0.0474910.543597 -0.0454960.581815 -0.0431330.619551 -0.0404090.656518 -0.0372800.692670 -0.0335390.728117 -0.0293280.762681 -0.0250640.796024 -0.0209430.827817 -0.0170970.857738 -0.0136170.885475 -0.0105580.910730 -0.0079440.933228 -0.0057690.952715 -0.0039910.969006 -0.0025020.982067 -0.0012570.991810 -0.0004040.997911 -0.0000421.000000 0.000000
S8064Actual
x/c y/c1.000000 0.0017150.993871 0.0026530.987827 0.0032850.978927 0.0038740.973715 0.0043440.966389 0.0050240.956849 0.0060260.941828 0.0076590.922531 0.0107760.902777 0.0144690.881039 0.0187170.851314 0.0251000.826902 0.0303580.796545 0.0368710.769121 0.0422930.739279 0.047552
Appendix A: Tabulated Airfoil Coordinates 317
0.707893 0.0521190.673035 0.0565880.637012 0.0604040.600096 0.0637940.565283 0.0665690.524114 0.0690320.486636 0.0707520.442310 0.0721330.402680 0.0724930.365608 0.0722550.329832 0.0711800.293320 0.0691710.256990 0.0665070.221551 0.0632870.186629 0.0594190.157005 0.0555610.130559 0.0513680.106077 0.0467900.092831 0.0439580.082007 0.0414470.069265 0.0381700.058443 0.0350430.050040 0.0323000.039568 0.0284730.033012 0.0257140.027019 0.0230300.020723 0.0200140.015385 0.0170860.009778 0.0132850.007961 0.0118830.005660 0.0098960.004071 0.0083830.002707 0.0068970.002381 0.0064650.000844 0.0039630.000447 0.0030560.000078 0.0016730.000042 -0.0012350.000215 -0.0022510.000408 -0.0031650.000742 -0.0040180.002460 -0.0063630.005238 -0.0089230.009825 -0.0118490.016648 -0.0152600.022790 -0.0178630.032365 -0.0212390.042250 -0.0239790.054804 -0.0268660.077840 -0.0312850.107408 -0.0360960.131843 -0.0392900.162688 -0.042769
0.195861 -0.0456540.229856 -0.0479190.261993 -0.0496360.308995 -0.0514150.338291 -0.0522040.375956 -0.0525970.415489 -0.0523150.450402 -0.0513280.490038 -0.0499270.540336 -0.0476030.578930 -0.0453260.615025 -0.0426740.649557 -0.0394560.683166 -0.0358800.715991 -0.0319960.747453 -0.0283680.778342 -0.0249020.806257 -0.0217240.834720 -0.0185640.858408 -0.0158450.882547 -0.0130610.902375 -0.0107900.918378 -0.0088150.936260 -0.0065760.953385 -0.0043450.967481 -0.0027440.974625 -0.0022220.987561 -0.0015310.993813 -0.0012830.996015 -0.0012101.000000 -0.001088
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
318 Summary of Low-Speed Airfoil Data
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”.
320 Summary of Low-Speed Airfoil Data
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
Run: bb05637 interpRe = 59971.8
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
Run: bb05639 interpRe = 100021.3
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
Run: jb05645 interpRe = 200448.7
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
Run: jb05647 interpRe = 300060.2
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
Run: bb05695 interpRe = 40171.5
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
Run: ts05699 interpRe = 60060.2
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
Run: jb05687 interpRe = 79763.7
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
Run: ts05689 interpRe = 100281.3
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
Run: bm05693 interpRe = 200562.2
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
Run: jb05697 interpRe = 300032.3
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
322 Summary of Low-Speed Airfoil Data
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
Appendix B: Tabulated Drag Polar Data 323
-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
324 Summary of Low-Speed Airfoil Data
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
Run: ts05712 interpRe = 59910.7
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
Run: bm05714 interpRe = 79894.3
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
Run: bm05716 interpRe = 99849.0
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
Run: bb05718 interpRe = 150237.6
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
Run: bb05720 interpRe = 200377.9
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
Run: jb05722 interpRe = 299777.3
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
Run: jb05724 interpRe = 449782.5
α Cl Cd
-2.99 0.073 0.0075-1.97 0.177 0.0065-1.47 0.229 0.0062
Appendix B: Tabulated Drag Polar Data 325
-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
Run: jb05735 interpRe = 59761.2
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
Run: jb05749 interpRe = 79627.0
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
Run: bb05731 interpRe = 99882.9
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
Run: jb05751 interpRe = 150245.8
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
Run: bb05733 interpRe = 199757.4
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
Run: jb05730 interpRe = 299723.3
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
Run: jb05737 interpRe = 498769.6
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
326 Summary of Low-Speed Airfoil Data
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
AG40d-02r fp2Fig. 4.25
Run: ts05740 interpRe = 60191.0
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
Run: bb05823 interpRe = 80069.2
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
Run: ts05743 interpRe = 100097.3
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
Run: jb05821 interpRe = 150168.8
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
Run: bb05745 interpRe = 199713.0
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
Run: bb05747 interpRe = 299613.7
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
Appendix B: Tabulated Drag Polar Data 327
AG40d-02r fp4Fig. 4.28
Run: ts05753 interpRe = 59857.3
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
Run: bb05830 interpRe = 79922.1
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
Run: bm05755 interpRe = 100157.2
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
Run: bb05828 interpRe = 150055.9
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
Run: bb05757 interpRe = 199955.9
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
Run: jb05759 interpRe = 299653.4
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
Run: jb05762 interpRe = 99973.0
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
328 Summary of Low-Speed Airfoil Data
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
Run: bb05779 interpRe = 99996.1
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
AG40d-02r fn5Fig. 4.37
Run: bb05767 interpRe = 99879.9
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
AG40d-02r fp5Fig. 4.40
Run: bm05770 interpRe = 99856.8
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
Run: jb05773 interpRe = 99980.6
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
Run: ts05776 interpRe = 99859.8
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
Run: bb05825 interpRe = 39997.3
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
Appendix B: Tabulated Drag Polar Data 329
-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
AG40d-02r fp0Fig. 4.53
Run: bm05815 interpRe = 60061.0
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
Run: ts05799 interpRe = 100158.5
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
Run: bb05801 interpRe = 200026.8
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
AG40d-02r fn2Fig. 4.55
Run: jb05819 interpRe = 100093.1
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
Run: jb05795 interpRe = 199963.0
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
Run: bm05817 interpRe = 299877.7
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
Run: ts05797 interpRe = 500311.7
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
330 Summary of Low-Speed Airfoil Data
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
AG40d-02r fp4Fig. 4.57
Run: bb05813 interpRe = 59818.4
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
Run: jb05809 interpRe = 99881.2
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
Run: jb05811 interpRe = 199865.3
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
Appendix B: Tabulated Drag Polar Data 331
-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
332 Summary of Low-Speed Airfoil Data
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
Appendix B: Tabulated Drag Polar Data 333
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
334 Summary of Low-Speed Airfoil Data
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
Appendix B: Tabulated Drag Polar Data 335
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
336 Summary of Low-Speed Airfoil Data
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
Appendix B: Tabulated Drag Polar Data 337
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
338 Summary of Low-Speed Airfoil Data
-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
Appendix B: Tabulated Drag Polar Data 339
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
Run: 06265rd interpRe = 59874.3
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
Run: 06267rd interpRe = 99904.9
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
Run: 06271rd interpRe = 199825.8
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
Run: 06274rd interpRe = 300087.5
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
340 Summary of Low-Speed Airfoil Data
NACA 43012AFig. 4.141
Run: bb05887 interpRe = 59967.6
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
Run: jb05890 interpRe = 100181.6
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
Run: ts05892 interpRe = 200094.5
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
Run: bb05897 interpRe = 299666.9
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
Run: jb05899 interpRe = 399727.7
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
Run: jb05901 interpRe = 499444.3
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
Appendix B: Tabulated Drag Polar Data 341
S8064Fig. 4.155
Run: jb05614 interpRe = 99850.6
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
Run: bb05617 interpRe = 200214.1
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
Run: ts05619 interpRe = 299862.0
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
Run: bm05622 interpRe = 399890.6
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
Run: bm05624 interpRe = 499910.8
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
Run: bm05781 interpRe = 59948.2
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
Run: jb05784 interpRe = 100270.5
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
342 Summary of Low-Speed Airfoil Data
-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
Run: ts05786 interpRe = 200140.6
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
Run: bb05788 interpRe = 299511.0
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
Run: bm05790 interpRe = 399536.2
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
Run: bm05793 interpRe = 500351.2
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
Run: jb05840 interpRe = 60088.0
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
Run: ts05842 interpRe = 99945.5
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
Appendix B: Tabulated Drag Polar Data 343
Run: jb05838 interpRe = 200067.2
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
Run: bb05846 interpRe = 299971.4
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
Run: jb05848 interpRe = 399815.5
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
Run: jb05850 interpRe = 500103.6
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
Run: ts05853 interpRe = 60069.0
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
Run: bm05856 interpRe = 100144.3
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
Run: bb05858 interpRe = 200322.6
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
344 Summary of Low-Speed Airfoil Data
8.30 1.217 0.02989.29 1.253 0.0376
10.27 1.269 0.0504
Run: bb05860 interpRe = 300099.7
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
Run: ts05881 interpRe = 399786.0
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
Run: bm05884 interpRe = 500121.8
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